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mpact of Surface Roughness onompressor Cascade Performance

eung Chul Backechanical and Aerospace Engineering,

eoul National University,eoul 151-744, Korea

une Hyuk Sohnenault Samsung Motors,ongin,46-796 Gyunggi, Korea

eung Jin Songechanical and Aerospace Engineering,

eoul National University,eoul 151-744, Korea

his paper presents an experimental investigation of roughnessffects on aerodynamic performance in a low-speed linear com-ressor cascade. Equivalent sandgrain roughnesses of 12 �m,80 �m, 300 �m, 425 �m, and 850 �m have been tested. Inondimensional terms, these roughnesses represent compressorlade roughnesses found in actual gas turbines. Downstreamressure, velocity, and angle have been measured with a five-holerobe at 0.3 chord downstream of the blade trailing edge. For theested roughnesses of 180 �m, 300 �m, 425 �m, and 850 �m,he axial velocity ratio across the blade row decreases by 0.1%,.1%, 2.5%, and 5.4%, respectively. For the same cases, the exitow angle deviation increases by 24%, 38%, 51%, and 70%, re-pectively. Finally, the mass-averaged total pressure loss in-reases by 12%, 44%, 132%, and 217%, respectively. Also, theoss increases more rapidly in the transitionally rough region.hus, among the three parameters, the loss responds most sensi-

ively to changes in compressor blade roughness.DOI: 10.1115/1.4001788�

IntroductionThe need to reduce the overall life cycle cost of operating a gas

urbine and the increasingly stringent regulation to reduce green-ouse gas �GHG� emissions are two of the most pressing con-traints for a gas turbine operator. Maintenance and fuel expendi-ures constitute approximately 90% of the overall life cycleacquisition, installation, fuel, maintenance, and operation� costsf a gas turbine. Therefore, gas turbine users have a strong incen-ive to minimize such costs by maintaining high efficiency opera-ion. Retaining high efficiency operation can also simultaneouslyeduce GHG emissions.

One of the main reasons of performance deterioration in gasurbines is the roughening of blades. For example, fouling of com-ressor blades due to particles ingested into the compressor de-rades gas turbine performance. Compressor fouling is defined ashe deposition of airborne particles onto compressor blades. Airntering a compressor contains various particles such as dust,and, and salts. In the compressor, these particles can combineith oil and/or moisture and get deposited onto the blades. Accu-ulated particles alter blade geometry, including the shape of the

eading edge and surface roughness. In turn, the altered blade

Contributed by the Fluids Engineering Division of ASME for publication in theOURNAL OF FLUIDS ENGINEERING. Manuscript received October 18, 2009; final manu-cript received May 11, 2010; published online June 23, 2010. Assoc. Editor: Chunill
ah.
ournal of Fluids Engineering Copyright © 20

ded 16 Feb 2011 to 147.46.117.245. Redistribution subject to ASM

geometry degrades compressor performance such as pressure ra-tio, efficiency, and mass flow rate. Such degradation leads to re-duced thrust for airborne gas turbines, reduced power output, andincreased heat rate in power generation gas turbines �1�.

For these reasons, the impact of blade surface roughness onturbomachinery performance has recently received increasing at-tention. However, much of such attention has been focused on theimpact of roughness in turbines. Taylor �2� measured turbineblade roughness in two military engines and found great spatialvariation in roughness. Bogard et al. �3� also made profilometermeasurements of sample turbine vanes from military engines.Bons et al. �4� made measurements, including the centerline aver-aged roughness, maximum peak-to-valley roughness, and rough-ness shape/density parameter in many different turbines. Bons etal. �4� also categorized turbine blade roughness according to theircauses—foreign deposits, pitting/erosion, and thermal barriercoating spallation. Yun et al. �5� examined turbine roughness ef-fects in a rotating rig and found significant performance degrada-tion. Bammert and Sandstede �6� conducted a boundary layer sur-vey to figure out the effect of surface roughness on a turbine bladeand found that the momentum thickness on the suction side de-velops more than that on the pressure side. Zhang and Ligrani �7�tested the effect of roughness in a turbine cascade. They testeduniform and variable roughness sizes and found that the total pres-sure loss increases with roughness. Also, some researchers con-ducted experiments in flat plates to examine the effect of surfaceroughness on the boundary layer. Tay et al. �8� conducted a par-ticle image velocimetry �PIV� experiment to examine the bound-ary layer characteristics in a roughened flat plate with a pressuregradient. They found that surface roughness decreases the massand momentum flux. Pailhas et al. �9� used a hot wire to find outthe effect of surface roughness and Reynolds number in a flatplate.

However, less research has been conducted on the impact ofblade surface roughness in compressors. Millsaps et al. �10� pro-vided a good review of the available literature. Milsch �11� re-ported a significant increase in profile loss with increasing rough-ness in a cascade. Mal’tsev and Shakhov �12� presented adeviation for a roughened compressor cascade and suggested acorrelation. Leipold et al. �13� also presented performance degra-dation due to roughness on compressor blades. Thus, investigatorspresent either a total pressure or flow angle data but not both for agiven cascade. Therefore, a cascade investigation in which bothtotal pressure and flow angle are measured would fully character-ize cascade performance and enrich the scarce database.

To satisfy such needs, an experimental study of roughness ef-fects on the performance of a modern compressor blade has beenconducted. Aerodynamic measurements have been carried out in alinear cascade. The specific research questions are as follows:

1. How does surface roughness affect profile loss in a compres-sor cascade?

2. How does surface roughness influence flow turning �devia-tion� in a compressor cascade?

2 Experiment

2.1 Test Facility. A low-speed, linear cascade at Seoul Na-tional University �SNU� has been used to examine the influence ofblade surface roughness on the aerodynamic performance of com-pressor rotor blades �Fig. 1�. A total of six blades have beenmounted in the test section, and two middle blades have beenroughened for roughness tests �Fig. 2�. Details of the facility canbe found in Sohn �14�.

Compressor blading used in this study is representative of themidspan of a compressor rotor blade in a modern industrial gasturbine. The geometric parameters of the test blade are listed inTable 1. The chord length and span of the actual blade are 205mm and 314 mm, respectively. The chord length of the test blade
is approximately the same as that of the actual blade, but the test
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lades have higher aspect ratios because the focus of this experi-ent is on acquiring flow data �loss and turning� at the bladeidspan. At nominal test conditions, the upstream velocity is 30/s, and the Reynolds number based on the blade chord length is

00,000.

2.2 Instrumentation. Upstream stagnation temperature andressure have been measured at one chord length upstream of theeading edge and the midspan height for reference purposes. Staticressure has been measured on the blade pressure and suctionurfaces to obtain blade loading information. There are 24 pres-ure taps on the suction side of the third blade and 18 pressureaps on the pressure side of the fourth blade at the midspan height.tatic pressure has also been measured on the test section casing

o ensure periodicity among the passages at 0.1 chord lengthownstream from the trailing edge of blades. Downstream at 0.3hord lengths from the trailing edge of blades, a five-hole probeas been traversed across one pitch to obtain flow angles, total andtatic pressures, and velocity at the midspan height. The measure-ent locations are shown in Fig. 3.

2.3 Data Reduction. Based on the pressure data obtainedrom pressure taps, the pressure coefficients have been calculated,

Fig. 1 Schemati

Fig. 2 Picture of the test section

Table 1 Geometry of cascade blade

omenclature Test blade

umber of blades 6hord, c 200 mmpan, H 196 mmitch, S 168 mmolidity, � �c /S� 1.19tagger angle, � 50.0°lade inlet angle, k1 63.3°lade exit angle, k2 40.8°

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Cp =P − P1

P01 − P1�1�

P01 and P1 indicate the total and static pressures of the main-stream measured by the upstream probe, and P is the static pres-sure measured at each pressure tap.

From the upstream and downstream total pressure data, the losscoefficient Yp, as defined in Eq. �2�, can be determined, whereP0= P0�y /S� indicates the total pressure measured by the down-stream five-hole probe,

Yp =P01 − P0

P01 − P1�2�

From the yaw angle data, the deviation has been obtained, asdefined in Eq. �3� below,

� = �2 − k2 �3�The velocity data have been normalized by the upstream veloc-

ity. From calibration results and repeatability tests, the uncertaintyintervals with a 95% confidence level for the pressure coefficient,loss coefficient, deviation, and velocity are estimated to be 0.003,0.002, 0.2°, and 0.02 m/s, respectively. Calibration process hasbeen conducted in a free jet flow by adjusting both yaw and pitchangles of the five-hole probe between �40° and 40° in 5° incre-ments.

3 Smooth Blade ResultsThis section presents experimental results for the smooth blade

case. The smooth blade results are also compared with other stud-ies’ results.

3.1 Periodicity and Blade Pressure Distribution. Figure 4shows the static pressure coefficient plotted across the pitch fortwo passages. The casing pressure data of adjacent passages atcorresponding pitchwise locations are uniform to within 2% of

f the test facility

Fig. 3 Measurement locations

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pstream dynamic pressure. Thus, good periodicity has beenchieved. Figure 5 shows the blade surface pressure coefficientistribution at the nominal test condition.

3.2 Deviation and Loss. Figure 6 presents the deviation plot-ed versus nondimensionalized pitch. Pitchwise locations 0 and 1re as shown in Fig. 3. In the freestream region, the deviation isearly 6°, and the mass-averaged deviation is 6.1°. From Carter’sule �15�, the deviation is predicted to be 6.05°, which matches thexperimental data well.

Figure 7 shows the loss coefficient plotted versus the pitchwiseocation. The width of the wake region is about 20% of pitch, andhe maximum value of the loss coefficient is 0.34. For a Reynoldsumber of 450,000 �similar to this study’s 400,000� and at a 0.55hord behind a blade, Leipold et al. �13� measured a wake widthf nearly 20% of pitch, a maximum loss coefficient of 0.3, and anntegral total pressure loss coefficient of 0.036 �Table 2�. Bothata sets are quantitatively similar and thus build confidence in theew data.

Fig. 4 Pitchwise end wall static pressure distribution

Fig. 5 Nominal blade pressure distribution

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4 Roughened Blade Results

4.1 Roughness Application. To represent the roughness of areal compressor to a cascade blade, 2D surface roughness profiledata of a real gas turbine compressor blade have been used �16�.Centerline averaged roughness height �Ra� has been obtained forsmooth and rough blades in various positions. From Ra values,

Fig. 6 Deviation distribution of smooth blade at 0.3 chorddownstream of the trailing edge

Fig. 7 Loss coefficient distribution of smooth blade at 0.3chord downstream of the trailing edge

Table 2 Comparison of pressure data for smooth blade

Re x /cMaximum

loss

Wakewidth�%�

AveragedYp

Present 400,000 0.3 0.34 20 0.030Leipold et al. �13� 450,000 0.55 0.3 20 0.036

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quivalent sandgrain roughness �ks� has been calculated accordingo the relation proposed by Koch and Smith, Jr. �17�. Based on thes values, the roughness Reynolds number �k+� has also been cal-ulated, and the range of roughness Reynolds numbers for the realompressor is between 20 and 150 �16�. The equations for rough-ess parameters are as follows:

Ra =1

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ks � 6.2Ra �5�

k+ = Recks

c�cf

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here

cf = 2.87 + 1.58 logc

ks−2.5

�7�

Cascade blades have been roughened by coating the blade withmery grain. Figure 8 shows a picture of smooth and roughenedlades. Equivalent sandgrain roughnesses of 12 �m, 180 �m,00 �m, 425 �m, and 850 �m have been tested. Grain sizesave been chosen by matching the roughness Reynolds number tohe real compressor roughness. They can be classified into threeoughness regimes—aerodynamically smooth, transitionallyough, and fully rough—and are listed in Table 3.

4.2 Axial Velocity. Figure 9 shows the axial velocity plottedersus pitch. Pitchwise locations 0 and 1 are as shown in Fig. 3.s the roughness size is increased, the wakes get wider andeeper. The increasing trend of wakes acts mainly on the suctionide. Figure 10 shows the axial velocity ratio plotted versusoughness. Axial velocity ratio decreases with increasing rough-ess almost linearly. Relative to the smooth case, the axial veloc-ty ratio decreases by 0.1%, 2.1%, 2.5%, and 5.4% for the testedoughness ks /c of 0.00090, 0.00150, 0.00213, and 0.00425.

4.3 Deviation. Figure 11 shows deviation plotted versusitch. Again, locations 0 and 1 are as shown in Fig. 3. As rough-

Fig. 8 Smooth and rough blades

Table 3 Roughness characteristics of test blade

ks��m� ks /c k+ Roughness regime

12 0.00006 2 Aerodynamically smooth180 0.00090 19 Transitionally rough300 0.00150 34 Transitionally rough425 0.00213 51 Transitionally rough850 0.00425 110 Fully rough

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ness increases, the maximum deviation locus shifts toward thesuction side. Also, the mean deviation increases with increasingroughness. The increasing deviation means that the flow is turnedless than that in the smooth condition. Figure 12 shows the mass-averaged deviation plotted versus roughness. The mass-averageddeviation increases from 6° for the smooth blade to 10° for ks /c of0.00425. Relatively, the deviation increases by 24%, 38%, 51%,and 70% compared with the smooth case. These values have beencompared with Mal’tsev and Shakhov’s correlation �12�.Mal’tsev’s correlation underpredicts deviation by approximately2°. The reason for this discrepancy could be due to the differencein blade camber—28° in Mal’tsev and Shakhov �12� and 22.5° inthe present case. However, the increasing trend of deviation forthe roughened blade is similar to the result shown in Mal’tsev andShakhov’s paper.

4.4 Loss. Figure 13 shows the loss coefficient plotted versusthe pitchwise location. The width and magnitude of the high lossregion get bigger with increasing roughness. Also, most of the

Fig. 9 Axial velocity distribution at 0.3 chord downstream ofthe trailing edge

Fig. 10 Axial velocity ratio versus roughness



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hanges can be seen near the suction side. This trend is also simi-ar to that found in a high-speed turbine cascade by Hummel et al.18�. The increase in loss on the suction side is likely due to thearlier transition and thickened wake region �11�. By using theseoss data, mass-averaged loss coefficient can be calculated, andig. 14 shows the mass-averaged loss coefficient versus rough-ess. Quantitatively, for the tested roughnesses, the loss coeffi-ient increases by 12%, 44%, 132%, and 217% compared with themooth case. Furthermore, the loss coefficient increases more rap-dly in the transitionally rough region and suggests the existencef a critical roughness. In the fully rough and aerodynamicallymooth regions, the changes in loss are less sensitive tooughness.

The wake thickening and consequent increase in loss due tooughness can be explained by a combination of the followingeasons. The first two reasons have to do with boundary layerevelopment on the blade surface, and the third reason has to doith wake evolution. The first reason is the earlier transition

ig. 11 Deviation distribution at 0.3 chord downstream of therailing edge

Fig. 12 Mass-averaged deviation versus roughness

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caused by roughness. Roberts and Yaras �19� reported an earliertransition due to the roughness on a flat plate at a Reynolds num-ber of 350,000, and Schreiber et al. �20� also presented an inducedtransition due to the roughness in a compressor cascade. There-fore, a similar phenomenon can be expected for the compressorcascade used in this investigation. Second, roughness also affectsthe evolution of the turbulent boundary layer. According to Brzeket al. �21�, roughness increases turbulent fluctuation and skin fric-tion, and high fluctuation induces a more rapid development of theturbulent boundary layer. Thus, turbulent boundary layers becomeeven thicker with roughness, and a thicker and bigger wake isformed at the blade trailing edge. Third, turbulence intensity in theboundary layer is increased toward the blade trailing edge due toroughness �13� and enhances mixing in the wake downstream.Thus, the width of and loss in the wake are both increased, andsimilar trends have been found by Zhang and Ligrani �7�. Figure15 shows a schematic �not to scale� of the first two effects on theblade boundary layer thickness.

Fig. 13 Loss coefficient distribution at 0.3 chord downstreamof the trailing edge

Fig. 14 Mass-averaged loss coefficient versus roughness

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ConclusionsFrom the experimental results, the following conclusions can beade.

1. As the surface roughness of the compressor cascade bladeincreases, the axial velocity is decreased, and deviation andloss are increased.

2. With increasing roughness ks /c from 0.0006 to 0.0090,0.00150, 0.00213, and 0.00425, the axial velocity ratio de-creases by 0.1%, 2.1%, 2.5%, and 5.4%.

3. With increasing roughness ks /c from 0.0006 to 0.0090,0.00150, 0.00213, and 0.00425, the deviation increases by24%, 38, 51%, and 70%.

4. With increasing roughness ks /c from 0.0006 to 0.0090,0.00150, 0.00213, and 0.00425, the loss increases by 12%,44%, 132%, and 217%.

5. Loss increases more rapidly with a transitionally rough sur-face.

6. Among axial velocity ratio, deviation, and loss, loss is theparameter most sensitive to changes in blade roughness.

cknowledgmentThe authors gratefully acknowledge the financial support from

he BK21 Program and the Micro Thermal System Research Cen-er, both of which are funded by the Korean Government, thenstitute for Advanced Machinery and Design of Seoul Nationalniversity, and Korea Midland Power Co., Ltd. �Project No.-2005-0-016�.

omenclaturec � chord length

H � spanS � pitch� � solidity �chord/pitch�� stagger anglek1 � blade inlet anglek2 � blade exit angleCp � static pressure coefficientP � static pressure on the blade

P0 � total pressure of the downstreamP1 � static pressure of the mainstream

P01 � total pressure of the mainstreamYp � loss coefficientk1 � cascade blade inlet anglek2 � cascade blade exit angle�1 � flow inlet angle of cascade

Fig. 15 Schematic of thickened wake due to roughness

�2 � flow exit angle of cascade

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� � deviationRa � centerline averaged roughnessyi � roughness height from centerlinecf � friction coefficientk+ � roughness Reynolds numberks � equivalent sandgrain roughness

Rec � chord Reynolds number

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Analytic Approach to Predict the Fouling Phenomena in the Axial Compressorof the Industrial Gas Turbine,” Proc. Inst. Mech. Eng., Part A, 219�3�, pp.203–212.

�2� Taylor, R. P., 1990, “Surface Roughness Measurements on Gas TurbineBlades,” ASME J. Turbomach., 112�2�, pp. 175–180.

�3� Bogard, D. G., Schmidt, D. L., and Tabbita, M., 1998, “Characterization andLaboratory Simulation of Turbine Airfoil Surface Roughness and AssociatedHeated Transfer,” ASME J. Turbomach., 120�2�, pp. 337–342.

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�6� Bammert, K., and Sandstede, H., 1980, “Measurement of the Boundary LayerDevelopment Along a Turbine Blade With Rough Surfaces,” ASME J. Eng.Power, 102�4�, pp. 978–983.

�7� Zhang, Q., and Ligrani, P. M., 2006, “Aerodynamic Losses of a CamberedTurbine Vane: Influences of Surface Roughness and Freestream TurbulenceIntensity,” ASME J. Turbomach., 128�3�, pp. 536–546.

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�9� Pailhas, G., Touvet, Y., and Aupoix, B., 2008, “Effects of Reynolds Numberand Adverse Pressure Gradient on a Turbulent Boundary Layer Developing ona Rough Surface,” J. Turbul., 9�1�, pp. 1–24.

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�13� Leipold, R., Boese, M., and Fottner, L., 2000, “The Influence of TechnicalSurface Roughness Caused by Precision Forging on the Flow Around a HighlyLoaded Compressor Cascade,” ASME J. Turbomach., 122�3�, pp. 416–425.

�14� Sohn, J. H., 2007, “Influence of Blade Surface Roughness on Flow Character-istics in a Linear Compressor Cascade,” MS thesis, Department of Mechanicaland Aerospace Engineering, Seoul National University, Seoul.

�15� Saravanamuttoo, H. I. H., Rogers, G. F. C., and Cohen, H., 2001, Gas TurbineTheory, 5th ed., Prentice-Hall, Harlow, p. 25.

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�17� Koch, C. C., and Smith, L. H., Jr., 1976, “Loss Sources and Magnitudes inAxial-Flow Compressors,” ASME J. Eng. Power, 98�3�, pp. 411–424.

�18� Hummel, F., Lotzerich, M., Cardamone, P., and Fottner, L., 2005, “SurfaceRoughness Effects on Turbine Blade Aerodynamics,” ASME J. Turbomach.,127�3�, pp. 453–461.

�19� Roberts, S. K., and Yaras, M. I., 2004, “Boundary-Layer Transition OverRough Surfaces With Elevated Free-Stream Turbulence,” Proceedings ofASME Turbo Expo 2004, Vienna, Austria, Paper No. GT2004-53668, Vol. 4,pp. 105–118.

�20� Schreiber, H.-A., Steinert, W., and Küsters, B., 2002, “Effects of ReynoldsNumber and Free-Stream Turbulence on Boundary Layer Transition in a Com-pressor Cascade,” ASME J. Turbomach., 124�1�, pp. 1–9.

�21� Brzek, B. G., Cal, R. B., Johansson, G., and Castillo, L., 2008, “TransitionallyRough Zero Pressure Gradient Turbulent Boundary Layers,” Exp. Fluids,
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