[American Institute of Aeronautics and Astronautics 47th AIAA/ASME/SAE/ASEE Joint Propulsion...

Development of Experimental Techniques for theCharacterization of Helicoidal Fin Arrays in Transonic Flow

Conditions

Laura Villafane�, Tolga Yasay, Guillermo Paniaguaz, Glauco Bonfantix,von Karman Institute for Fluid Dynamics, Rhode-Saint-Genese, 1640, Belgium

The current research focuses on the aerodynamic investigation of the three dimensional turbofan bypass-flow and its interaction with a finned air/oil surface heat exchanger integrated at the inner wall of the secondaryduct.

This paper addresses the experimental methodology employed in a new transonic facility characterizedby a complex 3D test section. The development of accurate and relatively short time response measurementtechniques is essential for the evaluation of the flow by means of map measurements at different locations alongthe test section.

Flow angles, total and static pressures are computed using a new methodology for the processing of thepressure readings acquired by means of hemispherical five-hole probes. The sensibility of the probes andthe errors on the angle determination are analyzed. Errors lower than 0.4 deg. for the determination ofyaw and flow angles are obtained. The response of the shielded thermocouples designed for the temperaturemeasurements are studied by means of conjugate heat transfer simulations. The numerical methodology isdescribed and the steady and transient temperature effects evaluated.

An innovative procedure for shear stress measurements based on oil-dot techniques is discussed and appliedon a representative test bench for validation. Results are comparable with theoretical and empirical wall shearstress correlations for the case of a flat plate in subsonic flow conditions.

I. Introduction

Developments in ultra high bypass ratio engines aim to achieve higher performances while further decreasingspecific fuel consumption over a wide range of missions. The evolution towards more efficient engine architecturesrequires advanced thermal management technologies to cover the increasing demand of refrigeration and lubrication.Oil systems hold a double function as lubricant and coolant circuits. They require efficient heat removal solutions for acontinued operation of the aero-engine. Fuel is the preferred heat sink for the lubricating oil systems, as illustrated byStreifinger,1 although its capacity is limited by several factors. In addition to the traditional limitations the developmentof components in composite material limits its the utilization. Air as a heat sink is extensively used in aeroengines.However within the oil systems the use of air coolers is generally limited to complement the capacity of fuel coolers2 incertain range of operations. The limitations presented by air cooled heat exchangers are mainly drag penalty, increasein mixing losses and decrease of the main mass flow if the air is bled from the compressor. The development ofair coolers optimized for certain locations within the aeroengine would reduce the adverse aerodynamic effects andincrease the required thermal capabilities.

Research on an efficient Air Cooled Oil Cooler heat exchanger integrated at the inner wall of the secondary duct ofa turbofan, as shown in Fig. 1, is currently performed at the von Karman Institute. The global objective is the analysisand understanding of the flow structure resulting from the interaction between the bypass flow and different geome-tries of finned heat exchangers. This knowledge will lead to the design of an efficient air cooled oil cooler locateddownstream of the air separator nose, to help the cooling of the oil circuit while introducing minimum aerodynamiclosses.�Ph.D. Candidate, Turbomachinery and Propulsion Department, Chaussee de Waterloo 72, [email protected] Engineer,Turbomachinery and Propulsion Department, Chaussee de Waterloo 72, [email protected] Professor, Turbomachinery and Propulsion Department, Chaussee de Waterloo 72, [email protected], AIAA Member.xPostgraduate Student, Turbomachinery and Propulsion Department, Chaussee de Waterloo 72, [email protected].

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American Institute of Aeronautics and Astronautics

47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit31 July - 03 August 2011, San Diego, California

AIAA 2011-6093

Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Figure 1. Heat exchanger location and research domain

Little public research exist regarding engine bypass flows. Authors addressing this topic focuss mainly on theimpact of the outlet guide vanes on the fan performances and on the acoustic effects , or on inlet sonic boom attenuationin supersonic flight. Aerodynamic characterization of this transonic flow and the effects of an obstacle interacting withit have not being addressed. The literature is also scarce on friction effects on heat exchangers at high Reynolds andhigh Mach number flows. Nevertheless , the utilization of heat exchangers in the same context as in this investigationhas been proposed in several patents, Refs.3–10

In order to evaluate the flow/finned heat exchanger interaction and its optimization a new wind tunnel test facilityhas been specifically designed as described in Ref.11 The new annular sector type test section is characterized by acomplex three-dimensional geometry to simulate the bypass-flow downstream of the engine fan.

The flow targeted in the present investigation involves certain experimental challenges, due to the characteristics ofthe flow, to the intermittent type wind tunnel (limited testing time), and to its complex test section with 3D walls andlimited transversal area. The confined 3D transonic flow and the flow structures developing within it downstream of thefins’ array will be evaluated using intrusive techniques that must provide sufficient spatial resolution accurately, andcausing minimum flow disturbances. The flow will be characterized by 2D traverses conforming map measurementsat different locations along the test section. All the techniques must provide a response time sufficiently low to reducethe waiting time in a measurement location. The response time should be also low enough to allow characterizing thestructures emerging from the flow/fins interaction. Conventional instrumentation is adapted and characterized to re-solve this particular flow. Special attention has been paid to the development of accurate directional and thermocoupleprobes. In addition, efforts have been also directed to the development of a new procedure to experimentally determinethe wall shear stresses. The analysis of the end-wall flow downstream of the fins’ array is of main importance to studythe flow evolution between different flow measurement planes.

Multi-hole probes are extensively used to provide flow direction and Mach numbers due to their robustness, sim-plicity and reduced cost when compared with other techniques for flow determination as laser Doppler anemometers(LDA) or hot wire anemometers. Numerous authors have reported works related with the multi-hole technology duringthe last decades, addressing the effects of probe geometry, calibration techniques and methodologies to process of thepressure measurements. Multi-hole probes allow determining not only the flow angles but also the static pressure andMach number of the flow. Hemispherical five hole probes have been selected in this case for showing better perfor-mances at transonic conditions as stated by Houtman and Bannink12 and Mukhopadhya et al.13 For the determinationof the flow quantities from the pressure readings, a new methodology developed by Yasa and Paniagua14 is employed.Calibration of the hemispherical five-hole probes is performed at different Mach numbers from 0.4 to 0.9, in the range-20 < pitch < 20, -30 < yaw < 30. Flow sensitivity analysis of the different pressure coefficients are performedat different angles and Mach numbers, and compared with the calibration results of a conical five-hole probe of 60�

cone angle at M=0.4. The uncertainty of the flow measurements, and the sensitivity of the flow angles computation isevaluated.

The utilization of thermocouples is the most extended technique for the determination of the gas temperaturethanks to its low cost, reliability, high precision, small size and suitability for almost all kind of flows. However theachievement of the gas temperature in the sensing probe is affected by several sources of error that have to be consid-ered in each particular application for a proper design of the probe. In the present application at high flow velocityand demanding moderate high response times for the map measurements within the testing time, a compromise solu-tion has taken to the design of shielded thermocouple rakes. Dynamic experimental calibrations are complex and not

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always accurate. A numerical procedure to analyze the response of the new shielded thermocouple and to evaluatethe influence of the different error sources is presented. Steady and unsteady temperature errors and the heat fluxesbetween flow, junction, wires and shield are studied.

To this date, measurement based on skin-friction lines constitute the main source of information to determine theboundary layer status and drag, and in general, for the understanding of the 3D separation. Surface flow visualizationtechniques are widely used to visualize wall streamlines, or skin-friction lines, due to the simplicity and low cost oftheir application. However the quantitative determination of the shear stress is a complex task in most of the exper-imental applications, preventing their routinely application in complex geometries. Within the present investigationeffort is directed to develop a technique to determine wall shear stresses applicable to surfaces with 3D shapes, withouteasy optical accessibility, and without involving sophisticated or expensive equipments. The procedure investigated isbased on thee thin film theory, and in the work of Tanner and Blows,15 who derived the wall shear stress as a functionof the oil viscosity and the pattern of the oil displacement. The new method is based on the conventional oil dot visu-alization technique and on the image analysis of the displacement of oil dots over a smooth surface under the presenceof a flow.

The flow measurements obtained with those techniques in the transonic bypass flow, in combination with the nu-merical results, will allow characterizing the interaction of the flow with the different fins’ array geometries tested, andevaluating the effectiveness of a new refrigerative solution for turbofans. Additionally this investigation proportionatesvaluable experimental data for the validation of engine models.

II. Experimental Methodology

The region of the turbofan under investigation is the portion of the engine comprised between the fan trailing edgeand the leading edge of the Outlet Guide Vanes. The air flow conditions to be reproduced correspond to the aerodesignpoint characterized by cruise flow velocities and standard pressure and temperature conditions at the turbofan inlet. Inthis domain flow magnitudes present variations of the same order of magnitude in radial and axial directions and alsothe same order of magnitude show the azimuthal and axial velocity components.

The methodology followed for the flow investigation includes the performance of aerodynamic flow measurementsand surface flow characterization. The selection and characterization of the measurement techniques is a major concernin this study due to the features of the flow and the complex test rig involved.

A. Wind Tunnel

The new transonic blowdown facility is driven by a pressurized storage of 72 m3 at 40 bar. The pressurized air flowsalong a duct circuit to a settling chamber, passing through a pressure control valve. The valve is operated to keep thetotal to static pressure ratio defining the operational Mach number in the test section constant. The turbulence intensityof the flow in the settling chamber is controlled by the use of screens and honeycombs. From the settling chamberthe air is first accelerated in a contraction. A set of Inlet Guide Vanes turn and accelerate the flow to reproduce theradial distributions of Mach number and flow angle downstream of the engine fan. The flow then evolves along thethree-dimensional test section. The air under the splitter is evacuated through a variable area restriction, allowing thecontrol of the mass flow and hence the pressure distribution around the stagnation nose. The core flow is expanded ina diffuser and later exhausted to the atmosphere.

B. Description of the test article

The new test facility aims to replicate the 3D turbofan bypass-flow environment, which requires simulating the flowgradients. To this goal a complex annular sector type test section, sketched in Fig 2, has been designed. Detailedinformation regarding the wind tunnel design is described in Ref.11

At the design conditions and in the absence of the fins, the test section Mach number is expected to vary between0.6 and 0.9, the maximum variation of total pressure is in the order of 600 mbar, and the flow angle is expected toshow a maximum variation around 15 degrees. Variations of temperature are expected downstream of the finned arraydue to the flow disturbances introduced. The characteristic area perpendicular to the flow at the location of the heatexchanger is about 0.013 mm2.

For the experimental flow characterization aerodynamic map measurements are performed at the five transversalplanes shown in Fig 3, to measure the three-dimensional flow with and without the presence of the investigated finnedarray. Traverse measurements of total temperature, total pressure, flow angle and turbulence intensity will be carriedout in those five planes and also upstream of the array of blades at the entrance of the test section. Static pressure

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Figure 2. Test Section Sketch with instrumentation carriage

will be measured at the walls along the wind tunnel and circumferentially at the location of the measurement planes.Visualization of skin-friction lines and measurements of wall shear stress will be carried out over the splitter surfacedownstream of the array of fins.

Figure 3. Measurement planes, a) 3D view, b) meridional view.

Several fin arrays with different geometrical configurations will be tested under the same flow conditions to obtaina better understanding of the flow fins interaction, and to identify the optimal array configuration.

From the comparison of flow results at the different measurement planes with and without the influence of thefin arrays, the flow effects caused by the latter ones will be studied. Downstream of the finned arrays analysis willbe done of the vortical structures created, their intensity and development between the measurement planes MP4 andMP5. Blockage effects and upstream flow modifications will be analyzed by comparison of flow results at the twomeasurement planes upstream the fin array leading edge. The aerodynamic effectiveness of the heat exchanger will beevaluated in terms of the ratio of the total pressure loss between two measurement planes upstream and downstream, inthe presence of the array with respect to the clean configuration. Balances of entropy generation will be also analyzedto quantify the flow disturbances.

C. Measurement Techniques

The design criteria common for all the aerodynamic instrumentation is based on the requisites of robustness, minimumblockage, high spatial resolution, flow angle sensitivity and limited time responses. Fixed the design as a compromisebetween the previous factors, the calibration and characterization of the different probes is required. It allows correct-ing the measurements and thus minimizing the errors. The knowledge of the response time of each probe is essentialin order to correlate the measurements with upstream conditions. It is required in this experimental case due to thequasi-steady character of the flow, in which the flow conditions slightly change as the deposit discharges.

Rakes have been chosen to perform map measurements in shorter test duration. In order to minimize the blockageeffects interest was focused on the miniaturization of the probes. All probes have been designed to introduce similarfrontal blockage and thus similarly affecting the flow. Probe miniaturization is also beneficial for increasing the spatial

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resolution. Shielded pitot probe heads are used for total pressure determination since they are more insensitive toangle effects. For the same reason hemispherical five hole probe heads have been selected for flow angle and pressuremeasurements. For the temperature measurements plastic shielded head probes with the wires perpendicular to theflow and suspended within the plastic shields were manufactured. The use of shields increase the recovery factor andreduces the conduction errors due to the position of the wires, without compromising the structural integrity of thethermocouples. However it also causes a loss of frequency response.16 The different probes employed are sketchedin Fig. 4. They consist on rakes of shrouded pitots, five hole probes, shielded T-thermocouples and hot wires. Alsocombined temperature and pressure probes are used for reference measurements, and a single head five hole probe formeasurements in the proximity of the surfaces. The maximum head probe diameter is 2.2 mm.

Figure 4. a) hemispherical five hole probe rake, b) reference probe, c) shielded thermocouple rake, d) hot-wire rake, e) shielded pitot proberake.

III. Flow Direction and Pressure

The expected flow angular variation in the test section between measurement sections is about 15 degrees in yawangle. The characterization and improvement of the technique employed for the flow angle determination by means ofthe hemispherical head 5 hole-probes, its accuracy and sensibility, is of utmost importance to increase the reliabilityof the measurements.

A. Probe head geometry

Non-dimensional coefficients are defined in any kind of multi-hole probes which relate the pressures measured bythe different sensing holes in order to relate the probe calibrations to the pressures measured during a test. Differentdefinitions of the coefficients are used by different authors. Whatever the definition, the geometry of the probe headshould procure, in the angular range of operation, the highest pressure-coefficient sensitivity to angle variation, andthe lowest dependence between pressure coefficients in order to obtain the best accuracy of the measurements.

Bryer and Pankhurst17 provide a wide overview analyzing different probes for the investigation of three-dimensionalflows. The effects of several five hole probe tip shapes at different velocities are addressed by Dominy and Hodson.18

Conical heads with forward facing tubes are really common due to their easy manufacturing, however they presentseveral drawbacks among which can be highlighted the dependence between the pressure coefficients computed fromthe pressure measurements. Gaillard19 analyzes the influence of the cone angle on five-hole conical probes with per-pendicular holes, for different Mach numbers. In the particular case of the hemispherical head, Mukhopadhya et al.13

compare the e effect of conical and hemispherical heads, concluding that the hemispherical one is more sensitive.Houtman and Bannink12 recommend the use of the hemispherical heads instead of the conical ones in transonic flows,due to the smooth variation of the surface pressures with the variation of Mach number.

The geometric characteristics of the hemispherical heads are shown in Fig. 5. The holes are separated 90 degreesaround the head axis and at 50 degrees from the tip. The pressure tubes diameters change from 0.3 to 0.4 mm in theprobe head and later in the probe stem to 0.6mm. The response time of the pressure measurements based on probetubing estimations is less than 5 ms.20

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1

Side view

Front views

Side cross-section

2

3

4 5

Figure 5. Design of the hemispherical 5 hole probe head and hole number location

B. Calibration properties

The probe is operated in the non-null method, in which the probe is held in a fix position during the test and the angleis derived from the pressure readings. With this method, calibration of the probe at different pitch and way angles andMach numbers is required.

Calibrations of the single head 5-hole probe and the two 5-hole heads of the probe rake are performed in a vertical-nozzle free-jet facility in which Mach numbers up to 0.9 can be achieved. This blowdown wind tunnel is fed by a40 bar pressurized reservoir (72 m3). Several grids are mounted in the settling chamber to guarantee the uniformityof the flow and the turbulence. The calibration nozzle has an exit diameter equal to 50 mm. Two computerizedcarriages allow running a detailed angular (yaw and pitch) calibration. The error in angular position of each step of thecalibration is less than +/- 0.1 deg. Calibrations of each 5-hole head are performed for pitch and yaw angles between�20� < pitch < 20�, and �30� < yaw< 30�, at steps of 2� . Calibrations are performed at Mach numbers from 0.4to 0.9.

The characteristics of the flow at a point in a steady three-dimensional flow are defined by four independentparameters: the two flow angles, the local Mach number and the local total pressure. Thus, in order to calculate theflow characteristics, at least four pressures have to be measured, while their relationship to the flow parameters isestablished by calibration by means of non dimensional coefficients. The methodology applied in this case to computethe flow quantities is based on six non dimensional coefficients. Four of them corresponding to each of the lateralpressure readings and independent of the other three. They are defined as the ratio between the pressure reading ofthe hole and the reading of the central one , both referred to the static pressure. The sensibility of the coefficientslinked with the lateral holes to the yaw angle, for a fixed pitch position is analyzed in Fig. 6. The angular sensitivityis expressed by the slope of the plots.

Figure 6. Sensitivity of the non-dimensional pressure coefficients of the four lateral holes to yaw variation

Two new spherical heads (A, B) are compared at minimum (0.4 Mach) and maximum (0.9 Mach) flow speed to

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verify manufacturing irregularities and Mach number dependencies. In head A the lateral holes, still each 90 degreesare slightly rotated from the vertical and horizontal references, while in head B they are inline with the reference lines.This affects mainly the distributions of the upper and inner coefficients. The head B shows a symmetric distributionaround 0 yaw, it is not the case for head A. The sensitivity of the left and right pressure coefficients is only slightlyaffected since the effect of the yaw variation is predominant. In general, the angular sensitivity is slightly reducedwhen the flow speed is increased. The sensitivity remains unaffected if a different pitch angle is considered for theanalysis. The results of the five hole probes are compared with a conical head probe with a cone angle of 60 deg.The conical head shows symmetric distributions, as in the case of head B, but the sensibility to the yaw angle isconsiderably decreased.

C. Flow angle determination

Different techniques have been proposed for the postprocessing of the pressure measurements. In any cases pressurereadings from the different holes are reduced to a certain number of non-dimensional parameters that are then com-pared to the calibration maps by means of different approaches. Those approaches can be direct interpolation in thecalibration maps , graphical methods,21 Taylor series decompositions22 , polynomial curve fittings23or zonal decom-position of the calibration range24among others. The common technique based on four parameters presents certainlimitations, as its failure when the calibration maps are distorted, or when one of the non-dimensional parameters iserroneous due to a wrong lecture on one of the pressure readings. The same limitations appear in the majority of theproposed postprocessing techniques based on non dimensional parameters computed as pressure differences betweenthe different readings. It can be avoided by handling directly the raw pressure levels without computing left-right pres-sure differences. The technique employed in this case uses a calibration pressures database instead of calibration maps.The database contains the information of the probe pressure readings during the calibration process that are translatedto six non dimensional pressures using the definitions given in Eqs. 1,2,3. The new data processing algorithm devel-oped by Yasa and Paniagua,14 involves also an iterative process computed for every set of pressures provided by theprobe. In a simplificative way, the method is based on the comparison of the non dimensional parameters from theprobe measurements with the values registered in the data base, from which regression coefficients are defined for aposterior polynomial surface fitting. A distinct peak of the polynomial surface of the regression coefficients revealsthe flow angles. The sensibility of the method can be referred as the sharpness of the peak, or valley if ( 1-r2) isrepresented instead.

cp1 =p1 � psp5 � ps


(1)



(2)


cpave =mean(p1�4)� ps

p0 � ps(3)

In Fig. 7a) the distribution of (1-r2) is plotted over the calibration range for the determination of a certain flowangle. To analyze the sensitivity of the flow angle determination, measurements on the calibration jet at differentflow velocities, probe positions, and with the different probes have been performed. The probes were set at a certainknown pitch and yaw angle at the outlet of the free jet and two Mach numbers were tested, M=0.4 and M=0.9. Figures7a) and b) represents the regression coefficients (1-r2) over the calibration range, for one of the hemispherical heads(B) with a incident flow at pitch=6� and yaw=0�, for the two Mach numbers tested. The same is represented inFig. 7c) for the conical probe head. The 1D plot in Fig. 7d) shows the 1-r2 distributions along a line at constantpitch angle as the indicated in Fig. 7a), for the different cases. The comparison of the solid and lines proves thatthe hemispherical head A in which the lateral holes are not aligned with the stem reference and its perpendicular,is slightly more sensitive than the probe B. The accuracy of the flow determination is improved for higher Machnumbers. It can be directly appreciated by the comparison of the contour plots. For the same flow velocity and flowangles, the conical probe shows a much wider valley, which in agreement with the evaluation of the sensitivity of eachnon dimensional coefficient, shows its inferior capability for an accurate determination of the flow angle. The accuracyof the determination of the flow angle is not dependent on the particular flow angle for a certain range of pitch andyaw. The sensibility is the same if the flow angle correspondent to pitch and yaw of 4�, than if the investigated angleis pitch=6� and yaw=0�.

The accuracy of the flow angle determination for the case in which the Mach number is known can be analyzedusing the data from the calibration at this correspondent Mach number. For this purpose all the calibration points are

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Figure 7. Contour distributions of (1-r2) distributions for pitch=6�, yaw=0� determined with: a)Hemispherical head B M=0.4,b)Hemispherical head B M=0.9, c)Conical head M=0.4. d)1D (1-r2) distribution for a different flow conditions and angles

Figure 8. Error evaluation in yaw and pitch angles

treated one by one. For the evaluation of a certain point (pitch and yaw angle values), the correspondent values ofpressure are considered as a test measurement. The information related to this point is removed from the calibrationdatabase. The methodology for the determination of the flow angles is then performed with the test pressures as aninput. The result is shown in Fig. 8 in which the contour plots represent the difference between the computed yawand pitch angles and the exact ones. This plot corresponds to the calibration of one of the hemispherical heads (A) atM=0.9. The dots show the real angular positions. The exact flow angle is computed with a maximum error limitedto 0.2 for incidence values within the interval �20 < yaw < 20 approximately. For higher yaw angles the error inthe yaw determination increases up to 0.3-0.4 degrees. It can be justified by two reasons. The sensibility in the lateralholes coefficients is linear within this same range. For higher absolute values of the yaw angle at which separationmay occur in one of the lateral holes, the sensitivity is altered. The second reason is based of the new methodologyin which the surface fitting is done taking into consideration the points with higher regression coefficients from thedatabase. In the limits of the calibration map, the information from the surrounding points is not balanced, since itdoesn’t exist information for higher angles.

The influence of flow speed on the flow angle is analyzed in Fig. 9. The data collected during calibration atMach 0.9 is treated as measurement and the flow angle is derived using calibration data of Mach 0.4. The error iscomputed as difference between the computed and real incidence angles. When the yaw angle is lower than 16 deg.the maximum error in pitch and yaw angle is limited to 0.6 and 0.9 deg., respectively.

IV. Flow temperature measurement technique

A wide variety of techniques to determine the total temperature of a flow are available nowadays. Among them,thermocouples continues to be the most used devices thanks to the low cost, reliability, high precision, simplicity, small

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Figure 9. Velocity influence in the error determination

size and suitability for almost any kind of flow. Thermocouples have been widely investigated since many years, butit is still a challenging technique due to the multiple factors influencing the measured quantity. The performance of athermocouple depends on its geometrical design and configuration, material properties, but also on the flow conditionsin which it is immersed, and the kind of measurements that aim to be achieved.

The objective of a thermocouple inserted in a flow is to retrieve the total temperature sensed in its junction, wherethe flow is brought to rest in equilibrium. The difficulty is that in practise this equilibrium does not exist since thejunction temperature results from a balance of heat fluxes between the gas convection, the probe stem conduction, thewalls radiation and the conversion of kinetic energy to thermal energy in the boundary layer around the thermocouple.Thus the transducer responds to a temperature different from the stagnation temperature even if it is able to reach asteady value in a long duration test.

A. Thermocouple design and analysis considerations

In the new wind tunnel the total temperature of the flow being discharged to the atmosphere from a pressurizeddeposit will slightly decrease during the run due to the expansion in the deposit. Regarding the measurements, maptemperatures will be recorded using traversing mechanisms at different plane locations. The knowledge of the responsetime, and its minimization, is required to correlate the measurements with reference conditions upstream. Due to thehigh flow velocity involved, structural characteristics also play an important role when designing the probe. TypeT shielded thermocouples, to increase the sensitivity, have been designed with the geometrical characteristics shownin Fig. 10. Shielded thermocouples increase the recovery factor due to a more efficient flow deceleration. Theinlet/outlet area ratio is 4, and the junction position agrees with the values recommended by Rom and Kronzon25 andSaravanamutto.26 However shielded thermocouples have a slower response time than bare thermocouples. The shieldis made of polycarbonate, chosen for its low conductivity. The wire diameter being 0.0254 mm, gives a ratio l=dwof 78. The shield diameter is a compromise between area blockage and the interest of high l=dw values to minimizeconduction errors. The present thermocouple configuration aimed to maximize this ratio without compromising thestructural resistance of the wires due to the high flow velocities. At transonic conditions bare thermocouples mayresult damaged due to wire bending. The thermocouple junction is 2.7 bigger than the wire diameter. The diffusivetime for both wire and sphere is lower than the response time.

Side

cross-section

Front view Bottom cross-section

Figure 10. Thermocouple design

Errors can be roughly estimated by applying existing correlations. They are expressed generally in function ofprobe geometry and material characteristics. However, they do not predict accurately each test case. In order toanalyze the different sources of error and to be able to correct them when possible, experimental calibrations of thethermocouples are performed. However, the experimental procedures to determine the probe recovery factor and theresponse time, are complex and require high accuracy . Moreover, some procedures only allow to study the time

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response under a certain frequency bandwidth due to the difficulty to experimentally simulate an instantaneous flowstep. Many different procedures have been used in order to analyze the response time. They can be divided intointernal and external heating techniques. Internal heating is based in the Joule effect, either applying a continuousor pulsating DC current,27 or with AC current.28 However they present several disadvantages as Peltier effects andnon-uniform temperature distributions. Regarding external heating different techniques have been applied. Based onflow jets: cold air jet impulses,29 rotating wheel chopping hot and cold tubes,30 the utilization of fast opening gatesto retain the flow impinging in the probe,20 or the displacement of probes into a jet.31 External heating techniquesare also the utilization of a laser beam.32 A new procedure to characterize the thermocouple probe is proposed in thepreset case.

B. Numerical methodology for thermocouple characterization

The proposed technique for thermocouple characterization is based on numerical analysis by means of transient con-jugate heat transfer simulations determining the temporal evolution of the temperature both in the fluid and within thethermocouple. It provides all the information regarding heat transfer fluxes within the assembly junction-wire-support,for a given set of flow conditions, thermocouple geometry and specifications. Therefore, it allows also to optimize thethermocouple design prior to its manufacturing.

The objective of the numerical analysis is the characterization of the response of the thermocouple under thesame conditions as in the experimental case. The analysis of the response time is based on the time reaction ofthe temperature of the thermocouple junction to a flow step. This is another advantage of the numerical method,since it is not possible to reproduce a real excitation step in experimental conditions. Dimensional analysis of theflow characteristic times, shows that that the characteristic time of the air flow to develop around the thermocouplefor the given flow velocities, is more than two orders of magnitude smaller than the characteristic time to get thermalequilibrium in the solids. It allows to partially decouple the fluid and solid problems facilitating the numerical problem.It is not necessary to impose a no-flow/flow step which is numerically complex to solve requiring a fine mesh andintegration times about 0.001 ms. Instead, it is equivalent to simulate a temperature step in the thermocouple froman already developed flowfield around it. Simulations are divided in two steps. First a steady simulation for the flowdevelopment around the solid. Secondly a transient conjugate heat transfer simulation including the thermocouplejunction, the wire and the shield on the thermal problem.

Figure 11. 3D mesh

According to those considerations two different cases have been simulated in order to evaluate the effect of thethermocouple wires. In one the thermocouple shield, wires and junction are considered within the fluid domain. A3D mesh is employed. In the second one only the shield and the sphere are considered within the fluid, in the absenceof wires. In this later case an antisymmetric 2D mesh is used. The domains and mesh refinements are identical inboth cases with the exception of the presence of the wires. The hybrid 3D mesh shown in Fig. 11 which refines alongthe walls of the solid parts is composed by about 1.75 million cells. The domain extends 6 shield diameters upstreamof the thermocouple and in radial direction, and 10 diameters downstream of it. The shield as well as the wire andthe junction are meshed independently and are concatenated in the NS solver used. The software employed, CFD++(v.8.1), uses a Reynolds Averaged Navier Stokes solver. The turbulence model considered is the k-epsilon model,with initial values of k and epsilon estimated as a function of the free stream nominal velocity, with a free-streamturbulence level of 1% and a turbulence length scale based on the tube inner diameter. The boundary conditions for thesteady simulation are velocity inlet conditions of M=0.7, Tt =300K, and atmospheric pressure, and atmospheric basepressure at the outlet. The solid are considered isothermal at 300 K. Convergence is achieved after 1000 iterations for

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the steady simulation, 5.5 hours CPU time running in 8 parallel Intel Core 2 Quad Q9400 (2.66 GHz) machines. Theunsteady simulations use the previous steady result as initial conditions. The temperature of the solids is released bychanging the boundary conditions to conjugate heat transfer boundary conditions. Different materials are consideredfor shield, wire and junction. The shield material is polycarbonate, the junction properties are considered as an averagebetween the two wire materials of type T thermocouples (copper-constantan), and two different possibilities have beenconsidered for the wires: copper and constantan. The thermal properties are summarized in table 1 correspond to areference temperature of 23�C. The time step in the unsteady calculations is adjusted as a function of the temperaturegradients starting from a time step of 0.1ms. The CPU time required to reach convergence in all the solids is defined bythe shield. Approximately 1000 iterations with different time steps are required, involving 51 hours CPU time runningin 8 parallel machines. However, the junction temperature remains steady after 9 CPU hours.

Copper Constantan PolycarbonateK, [W m-1 K-1] 401 19.5 0.2�, [kg m-3] 8930 8860 1210Cp, [J kg-1 K-1] 385 390 1250

Table 1. Material properties

C. Thermocouple characterization

1. Wire effects on recovery factor and time response

Comparison of the temporal evolution of the junction for the two mesh geometries, allows analyzing the effect ofthe wires in the present shielded thermocouple in which the wires are perpendicular to the flow. The effect of thewires orientation has been addressed, among other authors, by Moffat33 to asses the effect of increasing the heattransfer coefficient between the gas and the wire to minimize the steady temperature errors. Neglecting errors due toconduction, the comparison of the two configurations, can be considered an approximation to evaluate the effect of thewire orientation. The second case represents and ideal case in which the wires being parallel to the flow do not presentany interaction with the fluid.

Figure 12 shows the time evolution of the junction temperature for both cases. The material of the wires isconstantan and the Petit criterion34 to consider negligible conduction heat loss is satisfied. Attention must be paid forthe results interpretation to the fact that in this case it is not imposed a step of high temperature as is common practisewhen experimentally characterizing thermocouples. The total temperature or gas temperature to which the junctionshould arrive in the absence of any temperature error and an isentropic deceleration, computed at the entrance of thedomain is T0=299.8K. Regarding the steady effects, the recovery factor is 0.98 and 0.99 for the isolated junction andthe case including wires respectively. Thanks to the flow deceleration caused by the shield design and the high velocityflow for which the recovery factor is in general increased, the recovery factor achieved is satisfactory high. Howeverthe temperature decay in the case of the isolated junction is 1.5 times bigger than with the wires, resulting in a highertemperature error. The presence of the wires reduces the temperature error due to two main reasons. The first onealready introduced is the increase of conductive heat transfer caused by the wires orientated perpendicular to the flow.A theoretical value for hcv computed using correlations proposed by Moffat33 is hcv = 9884 W=m2K, relatively highthanks to the small wire diameter. The second one is due to the aerodynamic flow effect as illustrated in Fig 12. Thepresence of the wire forces the deceleration of the flow all around the junction, and not only in the stagnation pointand in the back of it, increasing the thermal boundary layer and therefore the gas recovery temperature.

Figure 12 allows evaluating the response time of the thermocouple. The dot lines represent the solution of a firstorder differential equation fitting the junctions temperature. In the absence of conduction along the wires, all theconvective heat flux is absorbed by the probe. Hence the energy balance of the junction temperature is reduced to afirst order differential equation. It is true for the case of the isolated sphere where no conduction effects are possible.However, the over-all response of a thermocouple will not be a first order system if the temperature of the junction isaffected by the temperature of the shields. In the general case, a combination of first order systems defines the transferfunction of the probe. As seen in the plot the presence of the wires modifies slightly the shape of the temperatureresponse. The response times indicated in the figure are defined as the time required for the thermocouple to complete63.2% of its response to the temperature step. The lower value of about 7 ms in the presence of the wires is expecteddue to the higher convective effects.

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Figure 12. Recovery factor and response time. Wire effects

2. Conduction effect

Heat conduction from the probe stem to the junction, or viceversa for a positive step or temperature, occurs due tothe temperature gradients along the wire between them. In order to analyze the effect of conduction for a givenprobe geometry, two different wire materials are considered: constantan and copper. For the comparison presentedthe influence of two of the factors affecting the temperature error can be neglected. The first one is the temperatureerror due to radiation since the considered temperatures are low. The second one is the effect of the velocity errorsince for a given flow and identical probe geometrical characteristics with the only difference of the wire material,the effectiveness of the flow deceleration is the same. In this conditions, the junction temperature is defined by thethermal balance between the conductive heat flux (shield-junction) and the convective one (gas-junction). Since theheat transfer coefficient hcv is independent of the wire material, the junction temperature is only affected by theconduction errors. The conduction error is function of the flow properties, the conductivity of the wire, and the ratiol=dw. Following Petit et al.34 conduction errors can be neglected when Eq.4 is satisfied.

l

dw2

rkgaskw

Nu > 10 (4)

For the present values of l=dw = 78, Nusselt number computed Moffat correlations of 8.822 for a Reynoldsnumber about 477 based on the wire diameter, and the air thermal conductivity, Petit criterium is satisfied for theconstantan, but not for the copper. Given the thermal conductivity of the copper, a l=dw ratio higher than 200 isrequired to avoid conduction errors.

To check the conduction effect two simulations where performed imposing the same junction material averageof the properties of copper and constantan, but different wire materials for each of them. Figure 13 represents thejunction temperature evolution for the two wire materials considered. Due to the higher conductivity of the copper thefinal temperature achieved is higher since the thermal inertia of the shield (polycarbonate) is very big and its initialtemperature is 300K. The constantan wire, facing the same flow and being in contact with the same support showsa junction temperature lower than the copper. Conductivity errors are negligible in the constantan and the differencebetween the junction temperature achieved and isentropic total flow temperature is due to the velocity error. In theabsence of conductivity errors the junction temperature would be the same in both cases.

The present numerical analysis allow evaluating the temporal evolution of the complete map of temperature gra-dients within the shielded thermocouple. It gives an important understanding of the complex heat transfer interactionbetween the different solid and fluid parts. The previous conclusion ensuring the reliability of the junction temperaturein the case of the constantan is demonstrated by analyzing the temporal evolution of the temperature along the wire.Figure 14 shows the temperature along the wire central line for different times for the Conjugate Heat Transfer simu-lation including the shield, constantan wire, and junction. A simulation solving the Conjugate Heat Transfer problemonly on the wire and the junction, and imposing isothermal boundary conditions at T=300 K in the shield walls, hasbeen also performed for the comparison of the real case with a situation in which there is not heat conduction betweenthe wire and the shield. The steady temperature distributions in both simulations differ on the wire temperatures closerto the shield. However the junction temperature remains unaffected. All the distributions for the different times col-

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Figure 13. Temperature error cause by conduction through the wires

lapse to the final junction temperature after 0.04 s, time for the thermal equilibrium of the junction. At the shield, thetemperature decreases with time from the initial temperature, but the time the shield employs to get to equilibrium ismuch higher.

T, K

Y, m

Y, m

X, m

, K

Figure 14. Temporal evolution of the wire temperature

With this analysis it can be concluded that the conductive temperature error is negligible for the constantan wire.This study confirms the validity of the theoretical work of Petit et al., confirmed by them experimentally for low flowvelocities.

V. Shear stress measurement technique

The quantitative determination of the shear stress is a complex task in most of the experimental applications.Some precise techniques capable to measure the wall shear stress are based on the diverging fringe shear stress sensor(OFI)35 and on the micro-pillar shear stress sensor (MEMS).35 The complexity of those techniques prevents theirroutinely application in complex geometries. Tian et al.36 succeed in measuring the wall shear stress in a low speedlinear cascade by means of LDA measurements. Naughton37 demonstrated that, at least in two-dimensional flows OilFlow Interferometry may be effectively used for skin friction measurements. In the present investigation an innovativetechnique to determine the shear stress based, on conventional oil dot visualization is applied.

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A. Theoretical analysis

An oil film generally deforms under the action of gravity, pressure gradient, surface tension and skin friction resultingfrom the gas flow over the oil surface. If the film is thin enough the dominant effect is the skin friction and the otherterms can be neglected. The procedure investigated is based on the thin film theory applied to oil droplets, and in thework of Tanner and Blows,15 who derived the wall shear stress as a function of the oil viscosity and the pattern of theoil displacement. The hypothesis considered are: 2D flow, Newtonian oil behavior, dominant effect of the shear stressover gravity, pressure gradient and surface tension, constant shear stress in small area elements, and linear variation ofthe thinning rate on the small elements considered. Under those hypothesis, it is possible to compute the shear stressfrom the rate of deformation of the oil droplet (length to thickness ratio) and its viscosity, as indicated by Eq.5, being1 and 2 two different time instants.

� = �1s2

� 1s2

t2 � t1(5)

B. Procedure

In order to compute the thinning rate, some authors use interferometric techniques. But it is complex and expensive.Instead the new procedure is based on image analysis. A CCD camera is used to record top surface dot evolution.A post-processing algorithm applies to each video frame the required corrections due to distortion in necessary, andfilters the images to enhance the image quality. Once the temporal evolution of the length and the area are known, thethinning rate is reconstructed. It is function of the length, the area, and the initial volume, given a certain shape for thecross section al deformation. A triangular shape as show in the scheme on Fig15 is considered.

For the se5We need then to characterize the oil dot viscosity. The oil to be use depends on the flow conditions. Forthe validation tests we use silicon oil mixed with micrometric titanium bioxide

The viscosity also plays a fundamental role in the calculation of the wall shear stress. Therefore it is importantto have a precise characterization of the viscosity of the oil mixture used. The oil mixture to be used depends on theflow conditions. For the validation test at subsonic velocitues different mixtures were examined. The selected paint iscomposed by a base of silicon oil mixed with titan-bioxide pigment. Its dynamic viscosity is characterized in functionof the temperature and the velocity gradient.

Figure 15. Oil thickness reconstruction

The new methodology is first investigated over a flat plate in subsonic conditions. A flat plate test bench is set upat the VKI free jet facility . The plate is vertically located in the jet. The velocity distribution on the plate is measuredby means of pneumatic taps in the center line of the plate. A shutter plate is placed at the exit of the jet to prevent theexposure of the oil dots to the air flow while the velocity is being adjusted. The plate is monitored by means of a CCDcamera. The images are recorded with a rate of 200 fps with a resolution of 800x600 pix. A light source heading tothe flat plate is used to to increase the contrast between dots and background. The oil dots are manually placed next tothe pneumatic taps locations. Tests are performed at different velocities ranging from Mach 0.22 to 0.42. Several testsare done at each velocity to ensure measurement repeatability, and to analyze the effects of the different parameters.

C. Results

The Computed values of the shear stress are shown in figure 16 at two locations along the plate. They correspond toa test performed at Mach number 0.42. They are compared with two analytical and two empirical correlations for thelaminar and the turbulent cases. The mean value has an uncertainty value and an error related to the closest referencevalue of the theoretical correlation. The boundary layer is expected to be turbulent at the conditions tested. A goodagreement is found with the correlations.

An uncertainty analysis has been performed first considering the error on the thickness reconstruction of the thinsheet using the linear fit. The error made is one order of magnitude smaller than the thickness itself and the errorin terms of wall shear stress is less than 1 per cent. A second uncertainty analysis has been carried out consideringa small variation on the longitudinal deformation of the dot. The error on the area determination could be due tothe optical setup (camera, light source, etc) and the following post processing image analysis procedure. For a two

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Figure 16. Wall shear stress at M=0.24

pixels variation, which corresponds to 1.5% of the maximum overall longitudinal deformation of the dots, the errorcomputed on the mean value is close to 4%. The imposed longitudinal variation of two pixels is considerably bigsince the resolution of the high speed camera is big and the robust algorithm for the image analysis allows to preciselyextract the border of the dots. The influence of the volume of oil has been also analyzed placing oil dots of 2 and 7�m . It has been proved that there is not a significant variation on the wall shear stress computation changing the oildot quantity.

The results obtained show that the oil dot technique gives values close to the theoretical and empirical wall shearstress for the simple case of a flat plate in subsonic flow stream.

VI. Conclusions

An innovative high speed subsonic wind tunnel has been recently designed to reproduce the real bypass flowconditions downstream of the fan of a gas turbine engine. The characterization of the complex 3D transonic flow isand its development along the test section is targeted. The testing time in the new facility is limited by the change ofthe flow temperature extracted from a compressed air deposit. Under those conditions a careful definition of the mostappropriated experimental techniques and a full characterization of their response if of upmost importance.

The development of an accurate technique for the determination of the flow angles, total and static pressure isdescribed. Hemispherical five-hole probes are designed for their higher sensitivity at transonic conditions accordingto the literature. Head dimensions are reduced to avoid blockage effects. Calibrations are performed at Mach numbersfrom 0.4 to 0.9, and flow angles in the range -20< pitch < 20, -30< yaw < 30. Sensitivity of the pressures recordedby the four lateral holes, in a certain non dimensional form, is satisfactory for all the calibration range. Evaluation ofthe new post-processing methodology with the calibration data yields errors lower than 0.4 degrees for yaw angles inthe range of � 0.4 deg. Maximum errors due to velocity if only the calibration data base for a given Mach number isconsidered are lower than 0.7 degrees.

The response of a new design of shielded thermocouples is evaluated by means of conjugate heat transfer sim-ulations. Numerical simulations are presented as a procedure to characterize the steady and transient behavior of athermocouple, without the necessity of the complex experimental calibration required for the determination of thetransient characteristics, i.e.,the response time. The different temperature errors affecting the new thermocouple de-sign are studied and evaluated. Petit et al.’ criterium for neglecting conduction errors is numerically probed by meansof the analysis of the temporal evolution of the temperature gradients within the wire and the shield. The increaseof the temperature recovery factor in the case in which the wires are perpendicular to the flow due to the increaseof conductive heat transfer and higher flow deceleration is also addressed. Response times in the order of 6ms areobtained for flow steps characteristic of the expected ones in the test section of the new wind tunnel.

A new procedure based on the surface oil dot technique is developed and validated in the well known case of aflat plate. It allows to obtain quantitative information of the wall shear stress by recording the dot deformation. Thetheoretical background is based on the thin film theory. Image processing techniques have been developed for thereconstruction of the temporal evolution of the oil dot thickness. Application of the new methodology to a subsonicflow over a flat plate show a good agreement with theoretical models and experimental correlations.

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Acknowledgments

This work is sponsored by TechSpace Aero in the frame of INTELLIGENT COOLING SYSTEM project. Thefinancial support of the Region Wallone and the pole of competitiveness Skywin is acknowledged. The authors aregrateful to T. Boeyen for the manufacturing of the instrumentation, V. Van der Haegen for his valuable contribution onthe conjugate heat transfer simulation and to F. Panerai for the final redaction of the paper.

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