aero-thermal investigation of a multi-splitter axial turbine

International Journal of Heat and Fluid Flow 32 (2011) 1036–1046

Contents lists available at ScienceDirect

International Journal of Heat and Fluid Flow

journal homepage: www.elsevier .com/ locate / i jhf f

Aero-thermal investigation of a multi-splitter axial turbine

J.P. Solano ⇑, V. Pinilla 1, G. Paniagua, S. Lavagnoli, T. Yasavon Karman Institute for Fluid Dynamics, Turbomachinery and Propulsion Department, Chaussée de Waterloo, 72, B1640 Rhode-Saint-Genèse, Belgium

a r t i c l e i n f o a b s t r a c t

Article history:Received 15 September 2010Received in revised form 28 February 2011Accepted 11 May 2011Available online 20 July 2011

Keywords:TransitionThin film gaugesTurbineConvective heat transfer

0142-727X/$ - see front matter � 2011 Elsevier Inc. Adoi:10.1016/j.ijheatfluidflow.2011.05.011

⇑ Corresponding author. Present address: Dep. IngUniversidad Politécnica de Cartagena, Campus MuralSpain. Tel.: +34 968 325938; fax: +34 968 325999.

E-mail address: [email protected] (J.P. Solano)1 Present address: Industria de Turbopropulsores

Zamudio, Edificio, 300, 48170 Zamudio, Spain.

This paper reports the external convective heat transfer in an innovative low-pressure vane, designedwith a multi-splitter configuration. Three aerodynamic airfoils are positioned in between larger struc-tural vanes, replacing struts presently used in current aero-engines, which results in superior aerody-namic performance. Static pressure and heat flux measurements were carried out in the largecompression tube facility of the von Karman Institute, using pneumatic taps and single-layered thin filmgauges respectively. The steady and unsteady heat transfer distributions were obtained at representativeconditions of modern aero-engines, with M2,is close to unity and a Reynolds number of approximately106. This facility is specially suited to control the gas-to-wall temperature ratio that drives transitionmechanisms. The heat transfer across the multi-splittered passages is confronted with correlations onducts to further characterize the boundary layer status. The data will be used to guide code developersby verifying their boundary layer transition models, and designers by showing the areas of the vanewhere heat transfer is most sensitive to the off-design conditions.

� 2011 Elsevier Inc. All rights reserved.

1. Introduction

The quest for high performance aero-engines has led to ultra-high by-pass ratio architectures. In this perspective, larger fandiameters demand a reduction in rotational speed to maintain anacceptable tip velocity. The low-pressure-turbine radius must beincreased in order to preserve high levels of efficiency. Further-more, to reduce the axial length of the turbine, a swan-neck dif-fuser is adopted to link the high-pressure and low-pressureturbines (Dominy et al., 1998; Arroyo Osso et al., 2010). Hence, ahigh hade angle appears in the transition duct between the high-pressure and low-pressure turbines. This is particularly evidentin the first low-pressure vane or intermediate pressure vane in athree-shaft configuration (Miller et al., 2001). In this paper, aninnovative low-pressure vane with multi-splitter configuration isstudied. Small aero-vanes are contained within large structuralvanes, which are used to support the engine shaft and house theoil circuit as well as the instrumentation. The large size of thestructural vanes represents an obstacle for the compressible flowreleased by the high-pressure turbine. The assessment of the heatflux load on this configuration is crucial to both design the cooling

ll rights reserved.

eniería Térmica y de Fluidos,la del Mar, 30202 Cartagena,

.S.A., Parque Tecnológico de

requirements to refrigerate this component and to assess theboundary layer status.

The adequate prediction of the boundary layer status is neces-sary to evaluate the profile loss across turbine passages. Thereforethe location of boundary layer transition is vital in order to obtainan accurate prediction. According to Denton (2010) novel aerody-namic configurations, such as the one presented, must be firsttested and understood in experimental facilities before beingexposed to CFD research. Experimental data such as the presentinformation represents an excellent database to validate transitionmodels.

The published research has provided insight into the heat trans-fer mechanisms in a highly unsteady environment, flow migrationacross turbine rows and wake and shock interactions. Ladisch et al.(2009) investigated the heat transfer related to pressure side recir-culation bubbles in a linear cascade. Additionally, Joslyn and Dring(1992) and Johnston and Fleeter (1999) performed experiments atlow speed in multistage configurations. Since the late 1990s, high-speed aero-thermal research on multistage interactions has beenconducted by groups at Oxford University (Povey, 2003), Ohio StateUniversity (Haldeman et al., 2004), von Karman Institute (Billiardet al., 2008), Wright Patterson Air Force Research Laboratory (Pol-anka et al., 2002) and QinetiQ in Farnborough (Haller and Hildith,2007).

Fig. 1 displays three possible architectures for a turbine stagesubmitted to large structural loads, which are particularlyprevalent in the first stage of a high-pressure steam turbine. Zhouet al. (1993a,b) compared numerically an architecture consisting ofstrut plus vane (Fig. 1a) versus a multi-splitter geometry (Fig. 1c),

http://dx.doi.org/10.1016/j.ijheatfluidflow.2011.05.011

mailto:[email protected]

http://dx.doi.org/10.1016/j.ijheatfluidflow.2011.05.011

http://www.sciencedirect.com/science/journal/0142727X

http://www.elsevier.com/locate/ijhff

Nomenclature

C airfoil chord (m)cp specific heat at constant pressure (J/(kg K))Dh hydraulic diameter (m)F frequency (kHz)g pitch (m)H airfoil height (m)h convective heat transfer coefficient (W/(m2 K))k thermal conductivity (W/(m K))N rotational speed (RPM)P pressure (bar)q heat flux (W/m2)R radius (mm)s airfoil curvilinear coordinate (m)T temperature (K)t time (s)U peripheral speed (m/s)v absolute velocity (m/s)x axial coordinate (m)

Dimensionless groupsNuDh Nusselt number, hDh/k (–)ReDh Reynolds number, qvDh/l (–)

Greek symbolsa thermal diffusivity (m2/s)l dynamic viscosity (Pa s)q density (kg/m3)X numerical domain

Subscripts0 free stream or total conditions1 stator inlet2 stator–rotor interface3 rotor outlet4 low pressure vane outletax axialLE leading edgeTE trailing edgew wall

J.P. Solano et al. / International Journal of Heat and Fluid Flow 32 (2011) 1036–1046 1037

quoting a profile loss reduction from 2.97% (in the strut plus vane,Fig. 1a) to 2.75% in the structural vane (Fig. 1b) and 2.33% in theaero-vane (Fig. 1c). However, the published research on multi-splitter low-pressure vanes is scarce from an aerodynamic pointof view and nonexistent for thermal analysis.

The current paper is focused on the boundary layer status andtransition of a multisplitter vane integrated in a high-pressureand low-pressure turbine layout. This analysis is of the utmostinterest during the design of compact s-shaped transition ducts,and remains a major concern for CFD validation at the presenttime.

In particular, experimental results on heat transfer and staticpressure distributions across the passages of the low pressure vaneare presented. The aim of the experimental methodology is to de-scribe the mechanisms driving boundary layer transition, for twodifferent pressure ratios across the stage. The potential effects ofthe upstream rotor on the heat transfer are further analyzed onthe basis of time-resolved data processing.

Fig. 1. Evolution of the m

2. Experimental apparatus

2.1. Turbine model

Measurements were carried out on a one and a half turbine stage.The single stage high-pressure turbine consists of 43 stator vanesand 64 shroudless rotor blades. The low-pressure vane is displayedin Fig. 2, where three small aerodynamic airfoils are located betweenstructural vanes. The characteristic dimensions of both airfoils aresummarized in Table 1. In total there are 16 big structural vanesand 48 small aero-vanes. Three measurement planes are identified,plane 1 at the high-pressure turbine inlet. Plane 3 and 4 representthe inlet and outlet of the low-pressure turbine vane.

2.2. Turbine wind tunnel

The von Karman Institute rotating rig can house turbine rows ofabout 900 mm in diameter. Fig. 3 displays the test section, located

ultisplitter design.

Fig. 2. High-pressure turbine and multi-splittered low-pressure turbine configuration.

Table 1Dimensions of the structural vane and the aerovane.

H/C g/C RLE (mm) RTE (mm) Dh,throat/g

Structural vane 0.52 1.43 8.24 0.53 0.67Aerovane 1.10 0.73 3.53 0.53 0.67

1038 J.P. Solano et al. / International Journal of Heat and Fluid Flow 32 (2011) 1036–1046

between a piston pressurized cylinder and the downstream dumptank at low pressure. All experiments were carried out using a tur-bine inlet temperature and pressure of 480 K and 2.3 bar respec-tively. A variable sonic throat is used to adjust the downstreampressure. Consequently, the Reynolds number and Mach numberare selected independently to match the engine conditions. Thisis necessary to reproduce the same boundary layer developmentand the pressure gradient which are present in the aero-engine.The Reynolds number based on the hydraulic diameter of the pas-sages and outlet velocity is approximately 106. The Mach numbers

Fig. 3. Turbin

at the low-pressure turbine inlet and outlet are 0.4 and 0.7 respec-tively. When the initially closed shutter valve opens, hot gas flowsacross a cold turbine, allowing heat flux measurements. The tem-perature ratio Tgas/Twall is similar to the actual aero-engine affect-ing the location of the boundary layer transition from laminar toturbulent (Arnal et al., 2008). The nominal turbine power of1.5 MW is released to a rotor with large momentum of inertia.The monitoring of the angular phase during the test enables todetermine the mechanical power (Paniagua and Yasa, 2007). Theturbulence intensity level is 5% at the turbine inlet and the lengthscale is 40 mm.

2.3. Measurement technique

2.3.1. Pressure measurementsStatic pressure is measured around the low-pressure vane sur-

faces by means of pneumatic tappings (0.8 mm in diameter). The

e test rig.


large and the small vane profiles are respectively instrumentedwith 27 and 19 pressure taps (see Fig. 4b). All instrumentation islocated at the mid-span of the airfoils. The reference total pressure(P03) and total temperature (T03) are measured at the low pressureturbine inlet plane by means of a Kiel probe and a K-type 25 lmwire thermocouple.

The overall measurement uncertainty is estimated to be±0.002 bar (20:1) for total pressure and 0.001 bar (20:1) for thesurface static pressure. The uncertainty associated with the totaltemperature measurement is ±2 K.

2.3.2. Temperature measurementsThe single layer thin-film gauges used at the von Karman Insti-

tute consist of thin platinum temperature resistors fired onto aceramic substrate. The entire blade could be manufactured in cera-mic for cascade testing (Arts et al., 1998), or inserts could be fittedinto metallic blades (Dénos, 1996).

In the current measurement campaign, Macor is selected as theceramic substrate, used to produce the inserts fitted into themetallic airfoils. Due to the resulting thickness of the thin-filmgauge and its high thermal conductivity, the frequency responseof the sensor can be as high as 50 kHz. In addition, many gaugescan be deposited on a small area of the same substrate, allowingfor the wall temperature history in sufficient locations of thecross-sectional area of the airfoil to be obtained. Fig. 4c presentsthe test specimens instrumented for heat transfer measurements.A sketch of the low pressure vane with the position of the sensorsis also included (d), which allows the definition of the four

Fig. 4. (a) Casing of the low-pressure vane. (b) Low pressure vanes instrumented with prthe low pressure vane: definition of passages A–D. Location of the gauges.

passages where heat transfer is analyzed. In order to be able tomeasure surface temperature changes at the time scale of the testduration (0.5 s) and the rotor blade passing events, an analoguefilter was implemented with a gain increasing in frequency. Ademodulation routine (Schultz and Jones, 1973) was implementedin Matlab to retrieve the temperature fluctuations.

3. CFD methodology

The numerical investigation has been performed using the Finesolver (Numeca, 2009). The model is identical to the geometryinvestigated experimentally. The computational domain consistsof 36 blocks, covering the NGV and rotor blade passages, plusone strut and three aero vanes. The mesh is generated usingaround five million grid points for the splitter-vane scheme (seeFig. 5). An O-type grid topology was selected around the airfoil,however a H-type grid is used in the rest of the domain.

The steady computation has been performed utilizing a mixingplane approach, with the Spalart–Allmaras turbulence model beingselected. A y+ value lower than 5 was calculated at the airfoil walls.The experimental total pressure and total temperature were im-posed as the inlet boundary conditions. The measured static pres-sure at the hub is used as outlet boundary condition and radialequilibrium is then applied. A periodic boundary condition is ap-plied at the sides of the domain. Two runs were performed for bothturbine operative conditions varying the pressure ratio of the HPturbine stage. The rotational speed of the rotor is the same forthe two runs, 6790 RPM.

essure taps. (c) Vanes instrumented with single-layer thin film gauges. (d) Sketch of

Fig. 5. Numerical model of the 1.5 stage.


4. Heat transfer data reduction

4.1. Evaluation of the time-mean heat flux

Several approaches are commonly employed for thin film gaugedata processing, namely analog-circuits (Doorly and Oldfield, 1986,1987), FFT techniques (Oldfield, 2007), and the numerical solutionof the unsteady heat conduction equation in the gauge substrate(Iliopoulou et al., 2004). Nevertheless, these methods rely on twoassumptions: the substrate can be considered as semi-infinite dur-ing the short duration test, and 1D heat conduction phenomenatake place. To account for radial conduction effects actually presentat the leading edge of the airfoil (Buttsworth and Jones, 1997) andfor the failure of the semi-infinite assumption in the trailing edge,the global approach developed by Solano and Paniagua (2009) isconsidered in this investigation. This data reduction techniqueconsists of solving the 2D unsteady heat conduction equation inthe cross sectional area of the vane, using as a boundary conditionthe reconstructed temperature history provided by the thin filmgauges in the contour of the airfoil:

1a@T@t¼ @

2T@x2 þ

@2T@y2 ð1Þ

Tðx; y;0Þ ¼ T0ðx; yÞ x; y 2 X

Tðx; y; tÞ ¼ Twallðx; yÞ x; y 2 @X 0 < t 6 tf

A weighted residual (Galerkin) approach is used to derive the fi-nite element equations from the governing differential equationEq. (1), as stated by Rao (1989). The solution of the resulting alge-braic system provides the time-dependent temperature distribu-tion inside the body and subsequently, the normal heat flux toits external boundary:

qwðx; y; tÞ ¼ �kmacor �~nrT x; y 2 @X ð2Þ

Due to the complexity of measuring the adiabatic wall temper-ature, the Nusselt number is expressed in function of the vane up-stream total temperature. The driving temperature T03 is obtainedusing a thermocouple probe placed upstream of the stator row at

50% span-wise position. The local hydraulic diameter of the pas-sages is chosen as the dimension to define the Nusselt numberfor each gauge position:

NuDh ¼qw

T03 � �Tw

Dh

kairð3Þ

To highlight the shortcomings of a purely 1D approach, the heatflux distributions along the aero-vane and structural vane profileshave been computed, prescribing the analytical wall temperatureevolution in a flat plate which applies to a heat flux step. This solu-tion is reported by Doorly and Oldfield (1987) for a 1D, semi-infi-nite substrate. Therefore, it discards any lateral heat conductionphenomenon. Fig. 6a depicts this analytical wall temperature evo-lution evaluated for a Macor substrate (q = 2520 kg/m3, cp = 752 J/kg K, k = 1.672 W/m K) and a wall heat flux step qw = 30 kW/m2 ex-tended over 0.5 s. These magnitudes are representative of the shortduration tests in the compression tube facility. Fig. 6b presents thewall heat flux evolution retrieved by the 1D solution of the heatconduction across the Macor substrate (Iliopoulou et al., 2004)and the equivalent result obtained in the stagnation point of theaero-vane with the 2D methodology, developed by Solano andPaniagua (2009). Whereas the solution of the 1D semi-infinite sub-strate fully reconstructs the wall heat flux step of 30 kW/m2, the2D solution decays short after the commencement of the test, ow-ing to the radial heat conduction effects present in the stagnationpoint of the aero-vane. At tf = 0.5 s, the 2D solution is 16.6% lowerthan the prescribed heat flux for this position. The heat flux distri-bution at the end of the test is shown in Fig. 6c for the aero-vaneand for the structural vane. The results show that the expectedheat flux step is reconstructed in wide regions of the pressureand suction sides of both airfoils during the whole test duration,proving that these regions fulfill the 1D hypothesis.

As previously shown, the wall heat flux evolution around theleading edge does not comply with the 1D model. However, thecurvature effects are less pronounced in the structural vane as a re-sult of the higher leading edge radius. The results in the trailingedge of the airfoils also underpredict the 1D solution by approxi-mately 25%, owing to the failure of the semi-infinite assumption.

The present methodology solves the 2D effects, relevant to crit-ical sensors of the airfoil. Fig. 6d depicts 1D and 2D Nusselt numberdistributions around the first aero-vane (passages A–B), in nominalconditions. The temperature distribution in the cross sectional areaof the airfoil reveals the lateral heat conduction phenomena pres-ent in the stagnation point. The temperature diffusion phenomenathat actually occur in the trailing edge of the airfoil also justifiesthe adequacy of a 2D approach.

4.2. Ensemble averaging of the periodic unsteadiness

Unsteadiness in any flow property can be decomposed into adeterministic and random component.

q ¼ �qþ q0uns ¼ �qþ q0deterministic þ q0random ð4Þ

with �q ¼ 1N

PNi¼1q, considering N samples.

The random heat flux unsteadiness is originated by turbulence,while the deterministic component is related to the model design.In the present research, the rotor blade passing events upstream ofthe low-pressure vane dominates the deterministic component.This is revealed by the analysis of the frequency spectrum. Fig. 7shows two dominant frequencies at 1 and 64 rotor blade eventper revolution. The first dominant frequency is due to differencesin blade geometries, which implies perfect periodic conditionsare only achieved once the rotor completes a full rotation. The 64events per revolution reveals interactions caused by the passageof each rotor airfoil, namely the potential flow field, wake and

Fig. 6. (a) Analytical wall temperature evolution for qw = 30 kW/m2; (b) wall heat flux computed in 1D and 2D Macor substrates; (c) 2D computation of the wall heat fluxaround the aerovane and structural vane; (d) 2D computation of the wall heat flux and substrate temperature distribution.

Fig. 7. Spectrum of the unsteady temperature.


compression waves.The corresponding deterministic heat flux isobtained by means of averaging the measured unsteady evolutionat each rotor blade-passing event. Let us consider n number ofperiods, or rotor blade-passing events. The position of the rotor isdefined throughout the phase (/), where the phase is equal to 0when the rotor leading edge is aligned with the structural vaneleading edge. Alternatively, a phase of 1 symbolizes when the rotorturns across a rotor blade pitch.

q0detð/Þ ¼Xn

i¼1

q0unsð/Þn

ð5Þ

The associate RMS is derived:

RMSð/Þ ¼ffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi

1n� 1

Xn

i¼1ðq0unsð/Þ � �qÞ2

rð6Þ

Fig. 8 displays a typical unsteady data analysis. The final stepdemonstrates the results of the time-resolved heat flux comparedwith the raw unsteady signal. The method commences with theanalysis of the raw unsteady wall temperature (Fig. 8 a), followedby selecting the first dominant frequency (one fluctuation per rev-olution). The average expressed by Eq. (5) is performed over onerotor revolution on the unsteady wall temperature, where / variesfrom 0 to 64, as shown in Fig. 8b. This trace contains 64 oscilla-

tions, one related to each rotor blade. The low frequency tempera-ture variation implies that the low-pressure vane is exposed todifferent deterministic unsteady total temperatures for differentrotor blade passing events. Such phenomenon is dictated by thedifferences in the rotor blade tip clearance, which creates differenttip leakage jets and vortices. Consequently, variations in the bladetip clearance result in different heat fluxes on the low-pressurevane, across a rotor revolution.

Fig. 8c depicts the unsteady deterministic heat flux across acomplete rotor revolution, with 64 oscillations per revolution. Notethat there are no low frequency oscillations, as the heat flux is thetime derivative of the wall temperature and therefore, the high fre-quency temperature fluctuations contribute more to the unsteadyheat flux. Fig. 8d finally presents the average of the 64 heat fluxoscillations per revolution, i.e. the phase locked averaged signalper rotor blade, or deterministic heat flux. In this case, the goodagreement between the raw and the periodic signal indicates alow contribution from random unsteadiness.

5. Results and discussion

5.1. Operating conditions

The experimental investigation has been carried out at twooperating conditions. The nominal conditions correspond to a stagepressure ratio (P01/P3) of 2.92, with a rotational speed of 6790 RPM.At this condition, the low-pressure vane is close to the chokingcondition, for which the turbine pressure ratio cannot be furtherincreased. Table 2 lists the investigated operating conditions. Theeffect of reducing the pressure ratio (P01/P3) = 1.92 is studied atnominal rotational speed (henceforth labeled off-designcondition).

5.2. Time-mean steady results

The presence of the large structural vane alters substantially theflow field entering the aero vanes, notably at the passages next tothe structural vanes (A and D). Therefore, the experimental resultsare analyzed passage per passage. The measured time-mean heattransfer are confronted with well-known correlations of local

Fig. 8. (a) Raw unsteady wall temperature; (b) unsteady deterministic wall temperature over one rotor revolution; (c) unsteady deterministic heat flux; (d) phase-lockedaverage signal.

Table 2Test conditions.

HPT rotational speed M3 HP exit angle P01/P4 M2,is,hub T03 (K)

Nominal 6785 0.42 12.5 3.76 1.09 354.7Off design 6788 0.31 �25.3 2.05 0.90 386.6


Nusselt number for simultaneously developing flow in ducts sub-ject to uniform heat flux. The solution for laminar developing flowis accounted for with the correlation of Churchill and Ozoe (1973):

Nux

4:364 1þðGz=29:6Þ2h i1

6¼ 1þ Gz=19:04

1þðPr=0:0207Þ23

h i12

1þðGz=29:6Þ13

h i13

2664

3775

32

8>>><>>>:

9>>>=>>>;

13

ð7Þ

In turbulent flow, the solution proposed by Bhatti and Shah(1987) is adapted for the computation of the local Nusselt number:

Nux

Nu1¼ 1þ 0:234

Dh

x

� �0:76

ð8Þ

where Nu1 denotes the fully developed Nusselt number.Static pressure (Ps/P03) and Nusselt number distributions along

the mid height channel are depicted in Fig. 9, at nominal condi-tions. The static pressure distribution describes the pressure gradi-ent driving the boundary layer development. Local values alongpressure side and suction side of passages A, B and D are presentedin function of the axial coordinate x/Cax,strut. The highest heat fluxlevel is measured in the stagnation region, where the boundarylayer begins to develop. Maximum values of Nusselt number ofthe order of 2000 are found at the leading edges. In the front sec-tion of the passage a steep acceleration occurs for both the suctionand pressure sides. Hence, the Nusselt number decays abruptly asthe laminar boundary layer develops along both sides of the pas-

sage. The comparison with the correlation data helps to identifythe transition location in each side.

Along passage A, a sudden increase of the Nusselt number in thepressure side for x/Cax,strut � 0.15 indicates transition onset withfurther turbulent boundary layer development. This is caused bythe flow deceleration from x/Cax,strut � 0.1 to x/Cax,strut � 0.5 in-ferred from the pressure distribution. Such adverse pressure gradi-ents promote the onset of transition, resulting in heat transferenhancement. This onset of transition is also noticeable via by apeak in the RMS of the unsteady data. The further acceleration ofthe flow from x/Cax,strut � 0.5 towards the rear part of the pressureside stabilizes and thickens the turbulent boundary layer, leadingto the decrease of the Nusselt number. Along the suction side ofthe passage, the Nusselt number decreases continuously from theleading edge of the aero-vane up to x/Cax,strut � 0.8. This evolutionis well predicted by the laminar correlation (Eq. (7)), with theexception of the gauge located at the leading edge of the aero-vane,submitted to the stagnation of the flow that approaches the aero-vane. From this point towards the rear part of the suction side,transition occurs and further thickening of the turbulent boundarylayer is observed.

Concerning passage B, the Nusselt number distribution alongthe suction side in nominal conditions presents a similar patternto that reported for passage A, with an earlier onset of transitionat x/Cax,strut � 0.7. It is worth noting that both correlations predictaccurately the local Nusselt number values in regions where theboundary layer is purely laminar or turbulent, respectively. Inthe front part of the suction side, the Nusselt number decreases

Fig. 9. Steady results at nominal conditions.


continuously from the leading edge up to x/Cax,strut � 0.7, indicatingthe growth of a laminar boundary layer. Further enhancement ofthe Nusselt number towards the rear part of the suction side re-veals the onset of transition and the development of a fully turbu-lent boundary layer. Pressure measurements along pressure side ofpassage B indicate that large velocity peaks occur in the front sideof the passage. The adverse pressure gradient promotes early tran-sition to turbulence in the pressure side, at x/Cax,strut � 0.5, Fromthis point onwards, turbulent boundary layer stabilized and thick-ens progressively, leading to a decrease of the local Nusseltnumber.

In regards to passage D, the transition onset in the suction sideis found at about x/Cax,strut � 0.15, corresponding to the flow decel-eration observed in the pressure distribution. The increase of areain the transition channel is not counter-balanced by the low curva-ture of the structural vane profile in that zone, resulting in diffu-sion. The turning imposed by the airfoils is enough to oppose thediffusion generated by the inter-turbine channel and acceleratesthe flow up to the geometric throat. The heat transfer characteris-tics along the pressure side of the channel is similar to that ob-served in passage B.

When the turbine operates at off-design conditions, the rotoroutlet flow angle changes significantly. The flow impinges on thestructural vane with a large negative incidence (�25�). The detri-mental effect on the aerodynamics is clearly visible on the pressure

Fig. 10. Steady results at o

profile: the stagnation point moves towards the suction side in allpassages as displayed by Fig. 10. The pressure levels are very sim-ilar at the rear part of the passages.

Regarding passage A, the flow abruptly decelerates on the pres-sure side at x/Cax,strut � 0.1. This results in boundary layer detach-ment because of the strong adverse pressure gradient thatclosely follows the velocity peak. Early transition to turbulence inthe suction side is promoted at approximately x/Cax,strut � 0.65.The flow approaching towards the suction side is also over acceler-ated, and impinges in the leading edge at higher velocities than forthe other aero-vanes. This effect enhances the heat flux in thisregion.

The separated region extends along most part of the structuralvane pressure side, which is supported by both the CFD simula-tions and oil visualizations presented in Fig. 11. The detached flowregion evidently affects the suction surface. The flow distortion im-posed by the low-speed recirculation structure obstructs a wideportion of the vane passage. Hence, the resulting suction sidevelocity profile is dominated by a large velocity peak in the leadingedge. The Nusselt number distribution reveals the poor perfor-mance for this operating regime. The front part of the pressure sidepresents local values of Nusselt number higher than those found atthe leading edge, due to the separation of the boundary layer. Reat-tachment and development of the turbulent boundary layer isfound for values of x/Cax,strut � 0.5.

ff-design conditions.

Fig. 11. Mach number distribution and streamlines at 45% of the span for at offdesign conditions (left). Flow visualization result of structural vane pressure side(right).


Owing to the blockage effect originated by this low-velocity re-gion, the flow is steeply accelerated along the suction side of thepassage A. The flow still approaches the aerovane with a negativeincidence, and early transition is promoted from x/Cax,strut � 0.6to x/Cax,strut � 0.75.

On passage B, the trend is similar to the nominal conditions,although the flow still approaches the aero-vanes with negativeincidence. Thus, the velocity peak on the pressure side is largerthan what has been measured for the nominal conditions.

The correlations reported in Eqs. (7) and (8) agree with theexperimental observation, which is more evident for passage B,where the flow is guided by both aero-vane surfaces, thus closelyresembling the flow in a pipe. For all passages, the abrupt decreaseof the Nusselt number due to the thickening of the laminar bound-ary layer is also well predicted. Overall, the experimental data arefound to lie within the heat transfer levels estimated by the lami-nar and turbulent correlations. Therefore, results give confidencethat Nusselt levels predicted by correlations for laminar and turbu-lent flow in a pipe may still provide an effective indication of theheat transfer variations found in turbine rows. Additionally, the

Fig. 12. Frequency spectra at nomi

static pressure distributions provide a guideline to assess the onsetof boundary layer transition. On the other hand, the largest dis-crepancy between the correlations and the experimental data isfound in the region where transition occurs, suggesting that thecorrelations are only valid for fully laminar and turbulent bound-ary layers.

5.3. Time-resolved unsteady analysis

The multi-splittered vane is being exposed to two main typesof periodic unsteadiness: events at once per revolution and onceper blade passage (blade passing frequency 7.2 kHz). The largeheat flux variation once per rotor revolution is associated to therotor tip clearance and thus different hot leakage jets. However,the following analysis will focus on the time-periodic signal atthe rotor blade passing frequency at design and off-designconditions.

Fig. 12 displays the Fast Fourier Transform (FFT) of the heatfluxes at design and off-design in several gauges across the pas-sages, in the range of frequencies between 0 and 25 kHz. The fun-damental frequency of the rotor blade passing and four harmonicsare well captured in the measurements. The amplitude of the fun-damental frequency tends to increase along the suction side. Thisphenomenon is related to the boundary layer status. As is shownin the time averaged results, once transition occurs, the boundarylayer becomes more energetic and the convective terms are stron-ger than in a laminar boundary layer. In a laminar boundary layerthe amplitude of the unsteadiness is low in comparison to a turbu-lent boundary layer, due to a lack of convective movement towardsthe wall. The high convective movement in a turbulent boundarylayer enhances the transport of the external fluctuations towardsthe wall.

At off-design conditions, the blade passing frequency is also no-ticed in the FFT, where a frequency band at around 1 kHz appearson the suction side of passage A. This frequency band is related tothe appearance of a separation bubble in the pressure side of thestructural vane.

nal and off-design conditions.

Fig. 13. Phase-locked averaged heat flux on two different passages at design conditions.


This information can be translated to the phase domain in orderto have information of the different phenomena and their relativeposition in phase. Fig. 13 displays the time resolved fluctuationsalong the suction side in passages A and C.

The heat flux shows a similar trend along the rear suction sideof the two aero vanes. All heat flux signals are characterized by aclear peak as one rotor blade sweeps one aero-vane pitch. Theamplitude of the heat flux peak grows larger from the aero-vanemid passage, 2–3 kW/m2, towards the trailing edge, where the heatflux rises to approximately 8 kW/m2. In Fig. 13 the dotted lineshows the location of the local maximum in the heat transfertraces along the rear suction side of passages A and C. It shouldbe noted that the heat flux fluctuations do not manifest any con-vective pattern. As the rotor sweeps across on vane pitch, the per-turbation that originates the heat flux variations appears to moveupstream. Therefore, the unsteady heat flux field in the low-pres-sure vane is governed by a potential interaction between the rotorand stator rows.

The viscous effects due to the sweeping of the unsteady rotorwake in the rear suction side are not observed in the time-resolveddata at the blade passing frequency.

Additionally, the time-resolved heat transfer data differdepending on the airfoil position relative to the structural vane.The large airfoil generates a strong potential field that locally mod-ifies the interactions between the unsteady flow structures period-ically released at the rotor outlet and the aero-vanes.

6. Conclusions

The steady and unsteady heat transfer of an innovative splitterlow-pressure vane have been investigated in the presence of afully rotating high-pressure turbine. Experiments reproduced

the relevant engine non-dimensional numbers, namely Reynolds,Mach and temperature ratios. Two operating conditions were car-ried out. The combination of static pressure and heat transfermeasurements helps to characterize the boundary layer statusand the main driving mechanisms in the boundary layertransition.

The steady heat flux data was compared to existing correlationsfor ducted flows, which allowed the detection of the onset of tran-sition. The results show that by-pass transition is triggered by anadverse pressure gradient. For design and off-design conditions,the flow is able to follow the aero-vane profiles, but in off-designconditions the aerodynamic behavior of the structural vane affectsthe aerodynamics of neighboring aero-vanes.

The unsteady data reduction performed, using the average overone rotor revolution, yields information concerning the low fre-quency phenomena related to non-identical rotor blade geome-tries. The time resolved data at the blade passing frequencyyields information about potential effects on the vanes due tothe interaction with the rotor. The unsteady heat flux providesvaluable insight into the physical causes dominating the heat flux.

The present research should provide directions for the design ofmulti-splittered airfoils for turbine applications. Furthermore, thevalidity of correlations was assessed, providing powerful tools forengineering evaluations of the thermal loads and coolingrequirements.

Acknowledgments

We gratefully acknowledge the financial support of the Euro-pean Commission and industrial manufacturers that participatein the Project Turbine Aero-Thermal External Flows 2, coordinatedby Snecma. We would like to thank Prof. Buchlin from the von Kar-man Institute for the interesting discussions on this work.


References

Arnal, D., Casalis, G., Houdeville, R., 2008. In: Saric, W.S. (Ed.), Practical TransitionPrediction Methods: Subsonic and Transonic Flows. RTO Lecture Series onTransition. von Karman Institute.

Arroyo Osso, C., Johansson, T.G., Wallin, F., 2010. Heat Transfer Investigation of anAggressive Intermediate Turbine Duct – Part 1: Experimental Investigation.ASME Paper No. GT2010-23653.

Arts, T., Duboue, J.M., Rollin, G., 1998. Aero-thermal performance measurementsand analysis of a two-dimensional high turning rotor blade. ASME Journal ofTurbomachinery 120 (3), 404–499.

Bhatti, M.S., Shah, R.K., 1987. Turbulent and transition flow convective heat transferin ducts. In: Kaka, S., Shah, R.K., Aung, W. (Eds.), Handbook of Single-PhaseConvective Heat Transfer. Wiley-Interscience, New York.

Billiard, N., Paniagua, G., Dénos, R., 2008. Impact of clocking on the aero-thermodynamics of a second stator tested in a one and a half stage HPturbine. Journal of Thermal Science 17 (2), 97–110.

Buttsworth, D.R., Jones, T.V., 1997. Radial conduction effects in transient heattransfer experiments. The Aeronautical Journal (May), 209–212.

Churchill, S.W., Ozoe, H., 1973. Correlations for forced convection with uniformheating in flow over a plate and in developing and fully developed flow in atube. Journal of Heat Transfer 95, 78–84.

Dénos, R., 1996. Aerothermal Investigation of the Unsteady Flow in the Rotor of aTransonic Turbine Stage. PhD Thesis, University of Poitiers.

Denton, J.D., 2010. Some Limitations of Turbomachinery CFD. ASME Paper No.GT2010-22540.

Dominy, R.G., Kirkham, D.A., Smith, A.D., 1998. Flow development through inter-turbine diffusers. ASME Journal of Turbomachinery 120 (2), 298–304.

Doorly, J.E., Oldfield, M.L.G., 1986. New heat transfer gages for use on multilayeredsubstrates. ASME Journal of Turbomachinery 108, 153–160.

Doorly, J.E., Oldfield, M.L.G., 1987. The theory of advanced multi-layer thin film heattransfer gauges. International Journal of Heat and Mass Transfer 30 (6), 1159–1168.

Haldeman, C.W., Dunn, G.M., Barter, J.W., Green, B.R., Bergholz, R.F., 2004.Experimental Investigation of Vane Clocking in a One and 1/2 Stage HighPressure Turbine. ASME Paper No. GT2004-53477.

Haller, B.R., Hildith, M.A., 2007. External Heat Transfer on a Shrouded HP GasTurbine Stage. ASME Paper No. GT2007-27168.

Iliopoulou, V., Dénos, R., Billiard, N., Arts, T., 2004. Time-averaged and time-resolvedheat flux measurements on a turbine stator blade using two-layered thin-filmgauges. ASME Journal of Turbomachinery 126, 570–577.

Johnston, D.A., Fleeter, S., 1999. Turbine Blade Unsteady Heat Transfer Change Dueto Stator Indexing. ASME Paper No. 99-GT-376.

Joslyn, D., Dring, R., 1992. Three-dimensional flow in an axial turbine: Part 1:aerodynamic mechanisms. ASME Journal of Turbomachinery 144 (1), 61–70.

Ladisch, H., Schulz, A., Bauer, H., 2009. Heat Transfer Measurements on a TurbineAirfoil with Pressure Side Separation. ASME Paper No. GT2009-59904.

Miller, R.J., Moss, R.W., Ainsworth, R.W., Harvey, N.W., 2001. Time-resolved vane-rotor-vane interaction in a transonic one and a half stage turbine. Proceedingsof the Institution of Mechanical Engineers Part A: Journal of Power and Energy215 (6), 675–685.

Numeca International, 2009. User Manual Fine/Turbo v8.7, Brussels, Belgium.Oldfield, M.L.G., 2007. Impulse Response Processing of Heat Transfer Gauge Signals.

ASME Paper No. GT-2006-90949.Paniagua, G., Yasa, T., 2007. Accurate turbine inertia measurement. Journal of

Experimental Mechanics 47 (5), 693–700.Polanka, M.D., Hoying, D.A., Meininger, M., MacArthur, C.D., 2002. Turbine Tip and

Shroud Heat Transfer and Loading Part A: Parameter Effects Including ReynoldsNumber, Pressure Ratio, and Gas to Metal Temperature Ratio. ASME Paper No.GT-2002-30186.

Povey, T., 2003. On Advances in Annular Cascade Techniques. PhD Thesis, Universityof Oxford, UK.

Rao, S.S., 1989. The Finite Element Method in Engineering, second ed. PergamonPress.

Schultz, D.L., Jones, T.V., 1973. Heat transfer measurements in short durationfacilities. AGARD Report No. 165.

Solano, J.P., Paniagua, G., 2009. Novel 2D transient heat conduction calculation in acooled rotor: ventilation preheating – blowdown flux. ASME Journal of HeatTransfer 131, 81601-1–81601-9.

Zhou, D.M., Zhang, Z.G., Li, Y.S., 1993. Numerical Prediction of Profile and End WallLosses for Multi-Splitter Turbine Cascades. ASME Paper 93-GT-255.

Zhou, D.M., Zhang, Z.G., Li, Y.S., 1993. Design and Calculation for Splittered AxialTurbine Cascade. ASME 93-GT-311.

aero-thermal investigation of a multi-splitter axial turbine

Documents