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Propulsion and Power Research
Propulsion and Power Research 2015;4(4):179–189
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ORIGINAL ARTICLE
Numerical investigation on heat transferin an advanced new leading edgeimpingement cooling configuration
G. Lina,n, K. Kusterera, A.H. Ayeda, D. Bohnb, T. Sugimotoc, R. Tanakac,M. Kazarid
aB&B-AGEMA GmbH, Aachen D-52070, GermanybRWTH Aachen University, Aachen D-52062, GermanycKawasaki Heavy Industries, LTD., Gas Turbine & Machinery Company, Akashi 673-8666, JapandKawasaki Heavy Industries, LTD., Technical Institute, Akashi 673-8666, Japan
Received 4 March 2015; accepted 4 August 2015Available online 11 December 2015
KEYWORDS
Gas turbine;Impingement cooling;Leading edge;Double swirl chambers
tional Laboratory foe (http://creativecom16/j.jppr.2015.10.00
or. Tel.: (þ49) 241
@bub-agema.de (G
esponsibility of Nat
Abstract It is known that the leading edge has the most critical heat transfer area of a gasturbine blade. The highest heat transfer rates on the airfoil can always be found on thestagnation region of the leading edge. In order to further improve the gas turbine thermalefficiency the development of more advanced internal cooling configurations at leading edge isvery necessary. As the state of the art leading edge cooling configuration a concave channelwith multi inline jets has been widely used in most of the blades. However, this kind ofconfiguration also generates strong spent flow, which shifts the impingement off the stagnationpoint and weakens the impingement heat transfer. In order to solve this problem a new internalcooling configuration using double swirl chambers in gas turbine leading edge has beendeveloped and introduced in this paper. The double swirl chambers cooling (DSC) technologyis introduced by the authors and contributes a significant enhancement of heat transfer due tothe generation of two anti-rotated swirls. In DSC-cooling, the reattachment of the swirl flowsalways occurs in the middle of the chamber, which results in a linear impingement effect.Compared with the reference standard impingement cooling configuration this new coolingsystem provides a much more uniform heat transfer distribution in the chamber axial directionand also provides a much higher heat transfer rate. In this study, the influences of differentgeometrical parameters e.g. merging ratio of two cylinder channels, the jet inlet hole
r Aeronautics and Astronautics. Production and hosting by Elsevier B.V. This is an open access article under themons.org/licenses/by-nc-nd/4.0/).3
5687860.
. Lin).
ional Laboratory for Aeronautics and Astronautics, China.
G. Lin et al.180
configurations and radius of blunt protuberances in DSC have been investigated numerically.The results show that in the DSC cooling system the jet inlet hole configurations have largeinfluences on the thermal performance. The rectangular inlet holes, especially those with higheraspect ratios, show much better heat transfer enhancement than the round inlet holes. However,as the price for it the total pressure drop is increased. Using blunt protuberances instead ofsharp edges in the DSC cooling can improve the heat transfer enhancement and reduce the totalpressure drop.& 2015 National Laboratory for Aeronautics and Astronautics. Production and hosting by Elsevier B.V.
This is an open access article under the CC BY-NC-ND license
(http://creativecommons.org/licenses/by-nc-nd/4.0/).
1. Introduction
In order to achieve high process efficiencies during theoperation of stationary gas turbines and aero engines,extremely high turbine inlet temperatures at adjustedpressure ratios are applied. The maximum hot gas tempera-ture is limited by the allowable material temperature of thehot path components, in particular the vanes and blades ofthe turbine. Thus, intensive cooling is required to guaranteean acceptable life span of these components.A large number of techniques have been developed in
recent years to enhance the convective heat transfer ratio forinternal cooling of turbine airfoils, e.g. rib turbulators, pinfins, dimpled surfaces, impingement cooing and swirlchambers. According to Ligrani et al. [1,2], the commonpoints of all these techniques are that they all can increasesecondary flows and turbulence levels to enhance themixing of the flows. In all these cooling techniques, jetimpingement has the most significant potential to increasethe local heat transfer coefficient. It has been widely used ingas turbine blade leading edge area, where extremely highthermal load exists.Over the past 50 years, numerous experimental and
numerical investigations on flow and heat transfer char-acteristics of impinging jets have been carried out. Severalgood reviews have been published by Martin [3], Han andGoldstein [4], Zuckerman and Lior [5], which summarizedthe most important results of investigations on impingingjets before 2006. Baughn and Shimizu [6] and Cooper et al.[7] experimentally studied a single circular turbulent air jetimpinging on a flat stationary surface. Lee and Lee [8]experimentally studied the effect of nozzle configuration onheat transfer of a single jet on flat surface. Since theimpinging jets have been widely used in a variety ofengineering applications with a curved surface like gasturbine blade leading edge, Lee et al. [9] investigated theeffect of concave surface curvature on heat transfer of afully developed round impinging jet in their experiment.Jordan et al. [10] investigated the influences of different jetgeometries on impinging jet on a cylindrical concavesurface. The measurement results show that the squareedge racetrack holes provide the highest stagnation regionNusselt numbers for a given jet mass flow rate. Many of the
previous works found in the literature have also dealt withnumerical studies on single jet impingement. Abdon andSunden [11] and Jia et al. [12] numerically investigatedsingle jet impinging on flat and concave surfaces. In bothstudies the secondary peak of heat flux for impinging on aflat surface with a small nozzle-to-plate distance, which isrelated to the wall jet boundary layer transition, cannot bepredicted. In the numerical study by Ibrahim et al. [13],single jet impinging on flat, concave and convex surfacesand multiple jets impinging on flat surfaces have all beeninvestigated. The results of the calculation with turbulencemodel V2F showed overall the best agreement withexperimental data compared with other turbulence modelsused for single jet impinging on different surfaces. Thesecond peak for Nu cannot be found using k-ε, k-ω or V2Fturbulence models in Reynolds-averaged Navier-Stokes(RANS) calculations. For multiple jets in three rows k-ωturbulence model presents the best agreement with experi-mental data. For impingement cooling on gas turbine bladeleading edge, intensively experimental and numericalinvestigations have been carried out by Taslim et al.[14,15], Taslim and Khanicheh [16], Elebiay and Taslim[17] and Yang et al. [18].
Facing the challenge of continuously growing turbineinlet temperature, the development of some new coolingconfigurations that can provide higher heat transfer rates hasbecome necessary. Recently, an alternative internal coolingconfiguration in the family of swirl chambers named doubleswirl chambers has been developed by Kusterer et al. [19].Swirl chamber is a kind of internal flow passage, in whichlarge-scale swirling of the flow circling under most circum-stances around the main axis is generated by internal insertsor outlets configurations [1]. The swirl can significantlyenhance the heat transfer rate. Double swirl cooling con-figuration can be generated by merging two swirl chambers.The numerical result by Kusterer et al. [19] showed that thiscooling concept presented much higher local and globallyaveraged heat transfer rate than the values in a standardswirl chamber. The major physical phenomena in the DSC-cooling and the main reason for the improvement of heattransfer are: (1) Heat exchange can be enhanced betweenthe two swirl flows in the shared section of two chambers;(2) Cross effect between two swirl flows can generate a
Nomenclature
D diameter of impinging inlet hole (unit: m)d diameter of double swirl chambers (unit: m)L length of impinging inlet hole (unit: m)L1 length of upstream channel (unit: m)L2 height of upstream channel (unit: m)L3 distance between two centrals in DSC (unit: m)Nu Nusselt number based on inlet jets diameterPout outlet static pressure (unit: bar)R radius of the circular arc (unit: m)r radius of blunt protuberance (unit: m)Rej Reynolds number based on jet hole diameterS space between two impinging jet inlets (unit: m)X coordinate in the target length direction (unit: m)Y coordinate in the target surface spanwise direction
(unit: m)Z space from the jet to the target surface (unit: m)yþ non-dimensional wall distance
qw wall heat flux (unit: W/m2)Tw wall surface temperature (unit: K)Tt,in upstream channel inlet air temperature (unit: K)
Greek letters
ρ density of fluid (unit: kg/m3)κ turbulent kinetic energy (unit: m2/s2)ω specific turbulence dissipation rate (unit: s-1)ε turbulent dissipation (unit: m2/s3)θ angle on the leading edge from the stagnation point
(unit: 1)
Abbreviation
DSC double swirl chambersMr merging ratioAr aspect ratio
Numerical investigation on heat transfer in an advanced new leading edge impingement cooling configuration 181
three dimensional “1” shape swirl flow, which improvesthe mixing of the fluids; (3) The reattachment of the flowwith the maximum velocity in the middle of two chambershas an impingement effect and downstream of the reattach-ment a restart of the boundary layer occurs, which results ina very high local heat transfer coefficient. The influence ofgeometry parameters e.g. merging ratio of chambers andaspect ratio of inlet duct on the cooling performance havebeen numerically investigated for the DSC cooling config-urations by Lin et al. [20,21]. Lin et al. [22] has also studiedthe influence of different inlet and outlet configurations onheat transfer and pressure loss characteristics of DSCcooling. In this study Lin et al. [22] investigated an inletslot in the middle of two chambers directly facing thereattachment region, which is very suitable for gas turbineblade leading edge impingement cooling. Kusterer et al.[23] and Lin et al. [24] have introduced how to use DSCcooling in the gas turbine blade. Figure 1 presents theschematic view of the application of double swirl chambersin blade leading edge impingement cooling.
This study presents the performance of double swirl cham-bers with focuses on the important thermal characteristics of a
Figure 1 Schematic view of double swirl chambers in blade leadingedge impingement cooling.
blade leading edge as the first stage development of a newinternal cooling method. At first, a systematic numericalvalidation in an impingement chamber is presented. Theexperimental geometry investigated by Yang et al. [18] hasbeen chosen as the physical models for this comparativecalculation. Several different turbulence models have been usedfor the calculations with STAR CCMþ, showing somevariations in the heat transfer prediction. After that, theinfluences of geometry parameters e.g. merging ratio ofchambers, impinging inlet hole configurations and radius ofblunt protuberance on the cooling performance and pressuredrops have been numerically investigated for the leading edgeDSC cooling configuration.
2. Computational setup
2.1. Geometry
Test case investigated by Yang et al. [18] has beenchosen for the validation, which is illustrated in Figure 2. Itconsists of an upstream channel and a concave targetchannel. The upstream and the target channel are connectedby 10 cylindrical impinging holes. Hot air flows from theupstream entrance and injects into the target channel to heatthe target surface. The target surface is a 2401 circular arc.The outlet of the target channel is directly open to theatmosphere. A summary of the geometrical details can befound in Table 1.
Figure 3 presents a schematic description of the leadingedge DSC cooling configuration. The main configurationse.g. upstream channel, impinging inlet hole positions, targetchamber length, circumferential length of the inlet hole andcircumferential length of the target chamber are all the samefor the investigated DSC cooling configurations and the testcase. The only difference between DSC cooling and the test
Figure 3 Geometry for leading edge DSC cooling configuration.
Table 2 Investigated cases for DSC cooling configuration.
Case No. Mr/% Ar r/D Rej
1 25 Cyclic 0 150002 23 Cyclic 0 150003 20 Cyclic 0 150004 15 Cyclic 0 150005 10 Cyclic 0 150006 20 1 0 150007 20 2 0 150008 20 3 0 150009 20 Cyclic 0.3 1500010 20 Cyclic 0.5 1500011 20 Cyclic 1 1500012 20 Cyclic 2 1500013 20 Cyclic 10 15000
Figure 2 Geometry for test case investigated by Yang et al. [18].
Table 1 Geometrical details.
R/D 2L/D 1Z/D 3L1/D 5L2/D 3.2S/D 3
Figure 4 Calculated surface averaged heat flux at leading edge targetwall for mesh independency study.
G. Lin et al.182
case is the cross section of the target channel. In the DSCcooing the R/D ratio is 1.19.The geometry parameter merging ratio is defined as
Mr ¼ 2R�L32R
ð1Þ
to describe the overlap of two single chambers. Thisparameter has influences on the cross section of the targetchamber. Another investigated geometry parameter is theimpinging inlet hole configurations. The original circularhole and rectangular holes with different aspect ratios havebeen investigated and compared with each other. In order tomake these different inlet hole configurations comparablewith each other the cooling air mass flow rate and the
circumferential length of the cooling hole have been keptthe same, which results in the same Reynolds number basedon jet inlet hole hydraulic diameter. The third investigatedgeometry parameter is the radius of blunt protuberance. Allthe investigated cases are listed in Table 2.
2.2. Numerical method and boundary conditions
A commercial CFD code STAR CCMþ has been appliedfor all the numerical investigations. Unstructured polyhedralgrids for calculation regions and prism layers in the near-wall region have been generated in STAR CCMþ. A meshindependence study has been carried out in three levels(1 million, 3 million and 5 million). The dimensionless walldistance value yþ of the first cell in the boundary layers isoverall less than yþ¼1 in three meshes. Stretch ratio in thedevelopment of the prism layers in normal direction of thewall surface has been controlled to be around 1.2. Thespanwise-averaged Nusselt number along the chamber axishas been chosen to investigate the mesh independency. Theshear stress transport (SST) k-ω turbulence model has beenused, which shows that the result from the calculation using3 million cells agrees well with the case using 5 millioncells. As shown in Figure 4, the deviation between case
Figure 6 Spanwise-averaged Nusselt number along thechamber axis.
Figure 5 Local mesh distribution for the calculation model.
Numerical investigation on heat transfer in an advanced new leading edge impingement cooling configuration 183
1 million and case 3 million is however much larger. Themesh with 3 million cells has been therefore chosen as thesuitable mesh for further investigations. The grid conver-gence index for this study is around 0.02%.
Figure 5 shows the mesh distribution for thecalculation model.
In the experiment, the heat transfer coefficients on thesurface are measured using transient liquid crystal. Thesurface temperature uncertainty caused by TLC is70.1 K.The K type thermal couple has an uncertainty of 70.5 K atreference temperature. The time resolution associated withthe CCD camera is around 70.1 s. All these factors resultin an approximately 12% uncertainty of the Nusseltnumber. Different impingement Reynolds number basedon the jet diameter from 10000 to 20000 have beenexperimentally investigated. In the present study, the testcondition with jet Reynolds number 15000 has been chosenfor the numerical calculation. Constant temperature is givenat the chambers wall with Tw¼419.15 K. Incompressiblefluid (air) with constant thermal physical properties hasbeen used to carry out the numerical calculations. The inletair has a temperature of 348.15 K before entering theupstream channel. All the investigated cases in this studyhave the same boundary conditions.
Figure 7 Nusselt number distribution on the target surface.
3. Results and analysis
3.1. Validation of the numerical approach
The numerical validations have been conducted withseveral different turbulence models: SST k-ω, Realizablek-ε, Wilcox k-ω, V2F and Spalart Allmaras. The results ofpredictions have been compared with the experimental databy Yang et al. [18].
Figure 6 presents the comparison of the calculatedspanwise (�751�751 of stagnation line) averaged Nusseltnumber along the chamber axis with the experimental data.The Nusselt number can be calculated in the numerical andexperimental as
Nu¼ hD
λ¼ qw UD
ðTw�Tt;inÞλð2Þ
As shown in this figure, the results obtained by theapplication of the code CCMþ depend on the chosenturbulence model. All the calculation results show much
higher peak value of the Nusselt number at impingementstagnation points than the measurements data. The SpalartAllmaras turbulence model shows overall the best accuracyamong all the investigated turbulence models, especiallyalong the second half of the chamber.
Figure 7 shows the Nusselt number distribution on thetarget surface. The concave surface near the stagnation line(�751�751) has been transformed to rectangles to showthe real area of all locations. Compared with the measure-ment data the calculated Nusselt number distributions showmuch stronger asymmetric effect along the second half ofthe chamber near the outlet. The positions of the stagnationpoints can be very well predicted by all the turbulencemodels, but the values are significantly over estimated. TheSpalart Allmaras turbulence model shows the best agree-ment with experimental data.
G. Lin et al.184
Figure 8 presents the calculated local Nusselt numberdistribution at axial positions X/D¼2, 8 and 24. Comparedwith all the other investigated turbulence models the SpalartAllmaras turbulence model shows again the best agreementwith the measurement data in all these three axial locations.Table 3 shows the static pressure loss from the air supplychannel inlet to the target channel outlet calculated bydifferent turbulence models. All the numerical calculationsshow an overestimation of the pressure loss.Based on the validation above, best results have been
obtained with the Spalart Allmaras turbulence model. In the
Figure 8 Nusselt number distribution on the target surface.(a) X/D¼2, (b) X/D¼8, and (c) X/D¼24.
following the alternative cooling configuration double swirlchambers for turbine blade leading edge cooling will beinvestigated numerically. Due to the similar main config-urations all the calculations and comparisons are based onSpalart Allmaras turbulence model.
3.2. Leading edge DSC cooling
In the study by Lin et al. [24], the DSC cooling presentsoverall a much higher heat transfer enhancement and also amuch more uniform distribution of Nusselt number alongthe target wall than the cooling configuration in the test caseby Yang et al. [18]. In the test case, due to the spent flowthe stagnation regions have been shifted to the downstreamdirection, which leads to lower heat flux values. In contrast,in DSC cooling the spent flow has almost no influence onthe positions of stagnation points and only a very littleinfluence on the peak values of the heat flux. Figures 9 and10 show the comparison between these two coolingconfigurations based on numerical calculations.
For both applications the same cooling air mass flow ratehas been applied. The target leading edge evaluation surfacein DSC cooling has the same area as in the test case(�751�751 of stagnation line). Thus, a comparison basedon heat transfer can be carried out directly. Table 4 presentsthe result of the comparison. The DSC cooling can generate24% more heat transfer than the referenced impingement
Table 3 Static pressure loss of the wholecooling configuration using different turbu-lence models.
Turbulence model Static pressure loss/Pa
Experiment 1500SST k-ω 1685V2F 1735Wilcox k-ω 1753Realizable k-ε 1702Spalart Allmaras 1698
Figure 9 Spanwise-averaged Nusselt number along the chamber axisfor two cooling configurations.
Numerical investigation on heat transfer in an advanced new leading edge impingement cooling configuration 185
cooling. The total pressure drop by DSC is also increasedby approximately 24%.
In the study by Lin et al. [24] the investigated DSCcooling configuration has a merging ratio of 20%. In thispaper, the influences of geometry parameters e.g. mergingratio, the impinging inlet hole configurations and radius ofblunt protuberance have been investigated and will bediscussed in the following chapters.
Figure 11 investigated configurations with different merging ratios.
Figure 12 Spanwise-averaged Nusselt number along the chamberaxis DSC cooling with different merging ratios.
3.2.1. Merging ratioLin et al. [20] has investigated the influence of geometry
parameter merging ratio on the heat transfer enhancementfor DSC cooling by using a tangential rectangular inlet slotas the swirl generator. The results show that increasingmerging ratios results in a smaller Nusselt number ratio andalso a smaller pressure drop.
This study shows the influence of merging ratio on thethermal performance of the leading edge DSC cooling withinline multiple inlet holes. Figure 11 shows the investigatedconfigurations with different merging ratios.
Figure 12 presents the spanwise-averaged Nusselt num-ber along the chamber axis for the investigated cases withdifferent merging ratios. The results show that the influenceof merging ratio on the heat transfer enhancement is not thatsignificant as in DSC cooling with only one inlet. As shownin Figure 13, all the cases with different merging ratiospresent very similar Nusselt number distribution. They allhave the linear impingement effect and almost the sameimpingement positions with very small influences fromspent flows. Table 5 presents the heat transfer in the leadingedge evaluation surface (same surface area for all cases) andthe total pressure drop for the investigated cases with
Figure 10 Comparison of globally Nusselt number distribution(view from leading edge).
Table 4 Comparison between impingement cooling and DSCcooling.
Configuraiton Impingement DSC
Cooling air mass flow rate/(kg/s) 2.45� 10�2 2.45� 10�2
Heat transfer/W 168 208Total pressure drop/Pa 1384 1716
different merging ratios. In DSC cooling with inline multi-ple impinging inlet holes, the geometry parameter mergingratio has negligible influence on the heat transfer and thepressure loss.
Figure 13 Comparison of globally Nusselt number distribution fordifferent merging ratios (view from leading edge).
G. Lin et al.186
3.2.2. Impingement hole configurationsIn the test case by Yang et al. [18] only a circular inlet
hole has been used. In the present study, three alternativerectangular inlet holes with different aspect ratios have beennumerically investigated and compared with the circularinlet hole configuration. Figure 14 presents the investigatedinlet configurations in this study. The rectangular inlet holeshave been expanded along the stream wise direction.Figure 15 shows the spanwise-averaged Nusselt number
along the chamber axis for different impingement inlet holeconfigurations. It can be found in this figure that the test caseswith rectangular inlet holes accomplish overall much higherheat transfer enhancement, especially along the second half ofthe chamber near the outlet. By using the rectangular inletholes the fluid is more concentrated in the midsection of theDSC, it thus results in a higher mass density near the wall andalso a higher swirl number, which can improve the heattransfer. This effect is even much stronger by using largeraspect ratios for the inlet holes. With larger aspect ratios, thepeak values of the Nusselt number at the impingement positionare increased along the second half chamber. At the same time,due to the increasing of swirl numbers the shifting effect of theimpingement positions is weakened.Figure 16 presents the comparison of Nusselt number
distribution for cases with different impingement inlet hole
Figure 14 investigated cases with different inlet configurations.
Table 5 Comparison among different merging ratios.
Configuraiton Total pressure drop/Pa Heat transfer/W
Impingement 1384 168DSC Mr¼25% 1701 210DSC Mr¼23% 1715 210DSC Mr¼20% 1716 208DSC Mr¼15% 1687 207DSC Mr¼10% 1662 208
configurations. With larger aspect ratios the heat transferenhancement is overall higher, and also the linear impinge-ment region is expanded.
As shown in Table 6, rectangular impinging inlet holeconfigurations can significantly increase the heat transfer inleading edge at the price of a significantly increased totalpressure drop.
3.2.3. Blunt protuberanceDSC cooling configurations are generated by the merging
of two cyclic chambers. After the merging two sharp edgesare generated, the manufacturing of this configuration can
Figure 15 Spanwise-averaged Nusselt number along the chamberaxis for DSC cooling with different inlet hole configurations.
Figure 16 Comparison of globally Nusselt number distribution fordifferent impingement inlet hole configuraitons (view fromleading edge).
Numerical investigation on heat transfer in an advanced new leading edge impingement cooling configuration 187
be very challenging in actual casted blades. Thus, the effectof blunt protuberances on the heat transfer performance hasalso been investigated in this study. Figure 17 shows theinvestigated configurations with different radius for bluntprotuberances.
Figure 18 presents the calculated spanwise-averagedNusselt number along the chamber axis for DSC coolingwith different radius of the blunt protuberances. The resultsshow that the radius of blunt protuberances has largeinfluences on the heat transfer enhancement in DSC. Thesharp edge model does not accomplish the best heat transferenhancement. With the increasing of the radius the heattransfer enhancement is also increased, because the flatwiseimpingement area directly facing the impinging jet inlets
Figure 18 Spanwise-averaged Nusselt number along the chamberaxis for DSC cooling with different blunt protuberances.
Table 6 Comparison among different impinging inlet holes.
Configuraiton Total pressure drop/Pa
Heat transfer/W
Impingement 1384 168DSC circular 1716 208DSC rectangular Ar¼1 2399 226DSC rectangular Ar¼2 3009 242DSC rectangular Ar¼3 4002 249
Figure 17 investigated configurations with different radius for bluntprotuberances.
grows. In the sharp edge model there is no flatwise surfacedirectly facing the impinging jet. The high heat transferenhancement mostly results from the reattachment of theflow to the wall and the new generation of boundary layers.The optimal heat transfer enhancement can be achieved inthe range of r/D¼0.5 to r/D¼1. With further increasing ofthe radius the heat transfer enhancement decreases due toweaker double swirl effect.
Figure 19 shows the comparison of globally Nusseltnumber distribution for cases with different blunt protuber-ance radiuses. With increasing radius the width of the linearimpingement region is increasing, but the effect of linearimpingement is weaker due to the unstable double swirls. Inthe case of r/D¼10, the double swirl fluid structure almostdisappears so that the Nusselt number distribution is verysimilar to that in the test case by Yang et al. [18].
Table 7 shows the heat transfer (in the same leading edgesurface) and the total pressure drop for all the cases with
Figure 19 Comparison of globally Nusselt number distribution fordifferent blunt protuberance radiuses (view from leading edge).
Table 7 Comparison among different blunt protuberanceradiuses.
Configuraiton Total pressure drop/Pa Heat transfer/W
Impingement 1384 168DSC sharp edge 1716 208DSC r/D¼0.3 1681 214DSC r/D¼0.5 1666 218DSC r/D¼1 1361 219DSC r/D¼2 1438 216DSC r/D¼10 1327 205
G. Lin et al.188
different blunt protuberance radius. The optimal radius forheat transfer enhancement is in the range of r/D¼0.5 to 1.The total pressure drop in DSC cooling is very close to thetest case with r/DZ1.
4. Conclusions
An alternative gas turbine blade leading edge internalcooling configuration has been introduced and numericallyinvestigated in the present work.At first, a systematic numerical validation on multiple
inline impinging jets on concave target channel is pre-sented. Several different turbulence models have been used,showing variations in the heat transfer prediction. And theones with Spalart Allmaras model show the best agreementwith the experimental data.After that, the influence of geometry parameters, e.g.
merging ratio, impinging inlet hole configurations and bluntprotuberance radius, have been investigated. The followingconclusions can be drawn:
1. By the application of the same boundary conditions andthe same main configurations the DSC cooling accom-plishes 24% higher heat transfer enhancement thanimpingement cooling with inline multiple jets.
2. The geometry parameter merging ratio has negligibleinfluence on heat transfer performance and total pressuredrop in the investigated leading edge DSC coolingconfigurations.
3. Rectangular impinging inlet hole configuration can sig-nificantly improve the heat transfer in DSC cooling at thecost of the increased pressure drop.
4. The blunt protuberance with radius ratio r/D from 0.5 to1 accomplishes the optimal heat transfer enhancement inDSC cooling. With r/DZ1 the pressure drop is veryclose to the impingement cooling with inline multiple jetinlets.
Acknowledgements
The numerical simulations presented in this paper havebeen carried out with the STAR CCMþ Software of CD-adapco. Their support is gratefully acknowledged.
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