Ralph J. Volino

United States Naval Academy, Annapolis, Maryland

Experimental Studies of Low-Pressure Turbine 

Flows and Flow Control

Final Report 

NASA/CR—2012-217656

September 2012

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Ralph J. Volino

United States Naval Academy, Annapolis, Maryland

Experimental Studies of Low-Pressure Turbine 

Flows and Flow Control

Final Report 

NASA/CR—2012-217656

September 2012

Notice for Copyrighted Information 

This manuscript has been authored by an employee of the United States Naval Academy under a NASA Interagency Agreement 

No. C–31011–K with the National Aeronautics and Space Administration. The United States Government has a nonexclusive, 

irrevocable, worldwide license to prepare derivative works, publish or reproduce this manuscript, and allow others to do 

so, for United States Government purposes. Any publisher accepting this manuscript for publication acknowledges that the 

United States Government retains such a license in any published form of this manuscript. All other rights are retained by 

the copyright owner. Attachments are reprinted by permission. 

National Aeronautics and

Space Administration

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Prepared under NASA Interagency Agreement No. C–31011–K 
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Available electronically at http://www.sti.nasa.gov

Trade names and trademarks are used in this report for identifi cation 

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either expressed or implied, by the National Aeronautics and 

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This work was sponsored by the Fundamental Aeronautics Program 

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Level of Review: This material has been technically reviewed by NASA technical management.

Acknowledgments

This project was performed under the NASA Glenn Research Center’s Low Pressure Turbine (LPT) Flow Physics program. 

The contract monitor was David E. Ashpis. His support, useful discussions, input, and suggestions are 

greatly appreciated and contribute to the success of this project. 
Experimental Studies of Low-Pressure Turbine 

Flows and Flow Control 

Final Report 

Summary 

An experimental study of the suction side boundary layer on an airfoil under simulated low pressure 

turbine conditions has been completed. Detailed measurements from the boundary layer have been 

acquired under a wide range of experimental conditions. The data have been saved and analyzed and 

results have been published. Models describing the flow behavior have been produced. The nature of the 

boundary layer separation, transition to turbulence, and reattachment has been documented. Methods for 

controlling the separation, transition and reattachment have been developed, experimentally implemented, 

studied in detail through boundary layer measurements, and reported through publications. Ten baseline 

cases without flow control were documented under high and low freestream turbulence conditions at five 

Reynolds numbers of 25,000, 50,000, 100,000, 200,000 and 300,000 based on passage exit velocity and 

suction surface wetted length. Boundary layer profiles were acquired at eleven streamwise locations 

between the leading and trailing edge of the airfoil, documenting the steamwise component of the 

velocity. Integral quantities such as boundary layer thicknesses and shape factors, as well as skin friction 

coefficients were computed based on the mean velocity results. Turbulence statistics were computed 

including the intermittency for the transitional flow. Large amounts of data were acquired at each 

measurement location in every profile, which allowed the calculation of turbulence spectra at all 

locations. The mean flow quantities, turbulence statistics and spectra have all been reported in the 

publications listed below. Addition two-component velocity profile measurements were acquired at the 

five most downstream measurement locations, where separation, transition, and reattachment occur. Mean 

flow quantities, turbulence statistics and spectra were again documented for the two component 

measurements, and included documentation of the turbulent shear stress and its spectra. 

Passive flow control was investigated under the same 10 experimental conditions with the same 

documentation at the five most downstream measurement locations. Passive control was first achieved 

using two-dimensional bars located at the pressure minimum on the suction side of the airfoil. Cases with 

three different bar thicknesses were fully documented, for a total of 30 new experimental cases. Cases 

with three-dimensional passive devices were documented next. The three-dimensional devices were small 

circular cylinders extending from the airfoil surface. Two different cylinder heights were considered, 

matching two of the two-dimensional bar heights. The cylinders were located at the same streamwise 

position as the two-dimensional bars. Many cases with different cylinder separations were documented. 

Data were acquired at multiple spanwise locations downstream of the three-dimensional devices. Based 

on the transition behavior and the spectral data of the baseline and passive flow control cases, a transition 

model and correlation were developed and successfully compared to other experimental cases from the 

literature. Passive flow control was successful at controlling the separation problem at low Reynolds 

numbers, with varying degrees of success from case to case and varying levels of impact at higher 

Reynolds numbers. 

Active flow control was achieved using a row of oscillating (synthetic) vortex generator jets located 

at the pressure minimum on the suction side of the airfoil. One case with low freestream turbulence and a 

low Reynolds number of 25,000 was studied. All of the data acquisition and documentation noted above 

were done for this case. Data were acquired at multiple spanwise locations to document the threedimensional nature of the flow downstream of the jets along the airfoil surface. Time averaged flow 

NASA/CR—2012-217656 1

Ralph J. Volino 

United States Naval Academy 

Annapolis, Maryland 21402 
results were presented and phase averaged results were also computed to show the response of the 

boundary layer to the unsteady jets. The active flow control successfully eliminated the large separation 

problem for the low Reynolds number case. 

Detailed documentation of the experimental conditions, descriptions of the experiments, experimental 

results and modeling efforts are presented for all the experimental cases in the following publications. 

Reprints are included in the appendix. 

Journal Publications 

1. Volino, R.J. “Separated Flow Transition Under Low-Pressure Turbine Airfoil Conditions: 

Part 1—Mean Flow and Turbulence Statistics,” ASME Journal of Turbomachinery, Vol. 124, 

pp. 545-655, 2002. 

2. Volino, R.J. “Separated Flow Transition Under Low-Pressure Turbine Airfoil Conditions: 

Part 2—Turbulence Spectra,” ASME Journal of Turbomachinery, Vol. 124, pp. 656-664, 

2002. 

3. Volino, R.J. “Passive Flow Control on Low-Pressure Turbine Airfoils,” ASME Journal of 

Turbomachinery, Vol. 125, pp. 754-764, 2003. 

4. Volino, R.J. “Separation Control on Low-Pressure Turbine Airfoils Using Synthetic Vortex 

Generator Jets,” ASME Journal of Turbomachinery, Vol. 125, pp. 765-777, 2003. 

5. Bohl, D.G. and Volino, R.J. “Experiments With Three-Dimensional Passive Flow Control 

Devices on Low-Pressure Turbine Airfoils,” ASME Journal of Turbomachinery, Vol. 128, 

pp. 251-260, 2006. 

6. Suzen, Y.B., Huang, P.G., Ashpis, D.E., Volino, R.J., Corke, T.C., Thomas, F.O., Huang, J., 

Lake, J.P. and King, P.I. “A Computational Fluid Dynamics Study of Transitional Flows in 

Low-Pressure Turbines Under a Wide Range of Operating Conditions,” ASME Journal of 

Turbomachinery, Vol. 129, pp. 527-541, 2007. 

Peer Reviewed Conference Papers (fully reviewed by 3 independent reviewers) 

1. Volino, R.J. “Separated Flow Transition Under Low-Pressure Turbine Airfoil Conditions: 

Part 1—Mean Flow and Turbulence Statistics,” ASME paper GT-2002-30236. Presented at 

the 2002 ASME International Gas Turbine Conference, Amsterdam, The Netherlands. 

2. Volino, R.J. “Separated Flow Transition Under Low-Pressure Turbine Airfoil Conditions: 

Part 2—Turbulence Spectra,” ASME paper GT-2002-30237. Presented at the 2002 ASME 

International Gas Turbine Conference, Amsterdam, The Netherlands. 

3. Volino, R.J. “Passive Flow Control on Low-Pressure Turbine Airfoils,” ASME paper 

GT2003-38728. Presented at the 2003 ASME International Gas Turbine Conference, Atlanta, 

GA. 

4. Volino, R.J. “Separation Control on Low-Pressure Turbine Airfoils Using Synthetic Vortex 

Generator Jets,” ASME paper GT2003-38729. Presented at the 2003 ASME International Gas 

Turbine Conference, Atlanta, GA. 

NASA/CR—2012-217656 2
5. Volino, R.J. and Bohl, D.G. “Separated Flow Transition Mechanism and Prediction with 

High and Low Freestream Turbulence under Low Pressure Turbine Conditions,” ASME 

paper GT2004-53360. Presented at the 2004 ASME International Gas Turbine Conference, 

Vienna, Austria. 

6. Bohl, D.G. and Volino, R.J. “Experiments with Three-Dimensional Passive Flow Control 

Devices on Low-Pressure Turbine Airfoils,” ASME paper GT2005-68969. Presented at the 

2005 ASME International Gas Turbine Conference, Reno, NV. 

Other Conference Papers (peer reviewed) 

1. Suzen, Y.B., Huang, P.G., Volino, R.J., Corke, T.C., Thomas, F.O., Huang, J., Lake, J.P., and 

King, P.I. “A Comprehensive CFD Study of Transitional Flows in Low-Pressure Turbines 

Under a Wide Range of Operating Conditions,” AIAA paper 2003-3591. Presented at the 33rd

AIAA Fluid Dynamics Conference, Orlando, FL. 

2. Volino, R.J. and Bohl, D.G. “Structure of Oscillating Vortex Generator Jets,” Proceedings of 

the Fourth International Symposium on Turbulence and Shear Flow Phenomena, Vol. 2, 

pp. 589-594, 2005. Presented at the Fourth International Symposium on Turbulence and 

Shear Flow Phenomena, Willamsburg, VA. 

Other Conference Presentations With Published Abstracts 

1. Volino, R.J. “Separation Control Using Synthetic Vortex Generator Jets,” Bulletin of the 

American Physical Society, DFD 2002, Vol. 47, No. 10, pp. 82-83. Presented at the 55th 

Annual Meeting of the Division of Fluid Dynamics of the American Physical Society, Dallas, 

TX, 2002. http://flux.aps.org/meetings/YR02/DFD02/baps/abs/S710004.html (as of 

November 21, 2011). 

2. Volino, R.J. “Active and Passive Flow Control on Low Pressure Turbine Airfoils,” 

Minnowbrook IV Workshop on Transition and Unsteady Aspects of Turbomachinery Flows, 

NASA/TM-2004-212913, p. 34. Presented at the Minnowbrook IV Workshop, Blue 

Mountain Lake, NY, 2003. 

3. Volino, R.J. and Bohl, D.G. “Separated Flow Transition Mechanism with Passive Flow 

Control under Low Pressure Turbine Conditions,” Bulletin of the American Physical Society, 

DFD 2003, Vol. 48, No. 10, pp. 75-76. Presented at the 56th Annual meeting of the Division 

of Fluid Dynamics of the American Physical Society, East Rutherford, NJ, 2003. 

http://flux.aps.org/meetings/YR03/DFD03/baps/abs/S310008.html (as of November 21, 

2011). 

4. Volino, R.J. and Bohl, D.G. “Three-Dimensional Flow Structure Resulting from Oscillating 

Vortex Generator Jets,” Bulletin of the American Physical Society, DFD 2004, Vol. 49, 

No. 9, p. 157. Presented at the 57th Annual Meeting of the Division of Fluid Dynamics of the 

American Physical Society, Seattle, WA, 2004. 

http://flux.aps.org/meetings/YR04/DFD04/baps/abs/S910010.html (As of November, 21, 

2011). 

NASA/CR—2012-217656 3

Appendix 

Reprints of Papers and Publications 

Copyrighted papers are reprinted with permission of the American Society of Mechanical Engineers 

(ASME) and the American Institute of Aeronautics and Astronautics (AIAA). 

 

NASA/CR—2012-217656 5
Ralph J. Volino

Mem. ASME,

Department of Mechanical Engineering,

United States Naval Academy,

Annapolis, MD 21402

e-mail: volino@usna.edu

Separated Flow Transition Under

Simulated Low-Pressure Turbine

Airfoil Conditions—Part 1: Mean

Flow and Turbulence Statistics

Boundary layer separation, transition and reattachment have been studied experimentally

under low-pressure turbine airfoil conditions. Cases with Reynolds numbers (Re) ranging

from 25,000 to 300,000 (based on suction surface length and exit velocity) have been

considered at low (0.5%) and high (9% inlet) free-stream turbulence levels. Mean and

fluctuating velocity and intermittency profiles are presented for streamwise locations all

along the airfoil, and turbulent shear stress profiles are provided for the downstream

region where separation and transition occur. Higher Re or free-stream turbulence level

moves transition upstream. Transition is initiated in the shear layer over the separation

bubble and leads to rapid boundary layer reattachment. At the lowest Re, transition did

not occur before the trailing edge, and the boundary layer did not reattach. Turbulent

shear stress levels can remain low in spite of high free-stream turbulence and high fluctuating streamwise velocity in the shear layer. The beginning of a significant rise in the

turbulent shear stress signals the beginning of transition. A slight rise in the turbulent

shear stress near the trailing edge was noted even in those cases which did not undergo

transition or reattachment. The present results provide detailed documentation of the

boundary layer and extend the existing database to lower Re. The present results also

serve as a baseline for an investigation of turbulence spectra in Part 2 of the present

paper, and for ongoing work involving transition and separation control.

@DOI: 10.1115/1.1506938#

Introduction

Modern low-pressure ~LP! turbine airfoils are subject to increasingly stronger pressure gradients as designers impose higher

loading in an effort to improve efficiency and lower cost by reducing the number of airfoils in an engine. If the adverse pressure

gradient on the suction side of these airfoils becomes strong

enough, the boundary layer will separate. Separation bubbles, particularly those which fail to reattach, can result in a significant

loss of lift and a subsequent degradation of engine efficiency ~e.g.,

Hourmouziadis @1#, Mayle @2#, and Sharma et al. @3#!. The problem is particularly relevant in aircraft engines. Airfoils optimized

to produce maximum power under takeoff conditions may still

experience boundary layer separation at cruise conditions, due to

the thinner air and lower Reynolds numbers at altitude. A component efficiency drop of 2% may occur between takeoff and cruise

conditions in large commercial transport engines, and the difference could be as large as 7% in smaller engines operating at

higher altitudes. Component life may also be affected by more

than an order of magnitude ~Hodson @4#!. Because the LP turbine

produces the bulk of the net power in many engines, changes in its

component efficiency can result in nearly equal changes in overall

engine efficiency ~Wisler @5#!. There are several sources for losses

in an engine, including secondary flows, but the suction side

boundary layer has been identified as the primary source of losses

in the LP turbine ~Curtis et al. @6#!. Prediction and control of

suction side separation, without sacrifice of the benefits of higher

loading, is therefore, necessary for improved engine design.

Separation on LP turbine airfoils is complicated by boundary

layer transition. Turbulent boundary layers are much more resistant to separation than laminar boundary layers. A substantial fraction of the boundary layer on both sides of a turbine airfoil may be

transitional ~Mayle @2#!, so accurately predicting transition location is crucial for accurate prediction of separation. Transition

prediction for turbine airfoils is complex and can depend on a

number of factors, including the free-stream turbulence intensity

~FSTI!, streamwise pressure gradient, airfoil curvature, surface

roughness, and the unsteadiness associated with passing wakes

from upstream stages. Several transition mechanisms are possible

under engine conditions. Mayle @2# classified the modes of transition as ‘‘natural transition’’ involving Tollmien-Schlichting

waves; ‘‘bypass’’ transition caused by high free-stream turbulence

or other large disturbances; ‘‘separated flow’’ transition of the

shear layer over a separation bubble; ‘‘periodic-unsteady’’ transition, which might also be called wake-induced transition; and reverse transition. If transition occurs far enough upstream, it can

prevent separation. If transition occurs in the shear layer over a

separation bubble, it will tend to induce boundary layer reattachment. The lower the Reynolds number, the farther downstream

transition will tend to occur, hence the problems associated with

performance at altitude.

Boundary layer transition has been studied extensively, and in

recent years several studies have focused on transition in the LP

turbine. Halstead et al. @7# present a study from a rotating cascade

with multiple stages and FSTI characteristic of engine conditions.

The adverse pressure gradients in this study were not strong

enough to induce separation, however. Solomon @8# subsequently

modified the facility and provides documentation of separation

from a more aggressive airfoil. Gier and Ardey @9# provide another example from a rotating facility. Boundary layers and separation bubbles on flat plates subject to adverse pressure gradients

have been considered in several studies. Recent work has included

the studies of Hatman and Wang @10#, Sohn et al. @11#, Lou and

Hourmouziadis @12#, Volino and Hultgren @13# and Yaras @14#. On

Contributed by the International Gas Turbine Institute and presented at the International Gas Turbine and Aeroengine Congress and Exhibition, Amsterdam, The

Netherlands, June 3–6, 2002. Manuscript received by the IGTI, January 22, 2002.

Paper No. 2002-GT-30236. Review Chair: E. Benvenuti.

Journal of Turbomachinery Copyright © 2002 by ASME OCTOBER 2002, Vol. 124 Õ 645

NASA/CR—2012-217656 6
airfoils, either in cascade or single-passage cascade-simulators,

studies have included Murawski et al. @15#, Qiu and Simon @16#

and Simon et al. @17#. Cascades with moving wakes, simulating

the effect of upstream blade rows, were utilized by Brunner et al.

@18#, Stadtmu ̈ller et al. @19#, Howell et al. @20# and Kaszeta et al.

@21#. Numerical studies have included the work of Dorney et al.

@22#, Chernobrovkin and Lakshminarayana @23#, Huang and

Xiong @24#, and Thermann et al. @25#. A few studies have included

attempts to control transition and separation. Howell et al. @20#

studied modified airfoil shapes, Van Treuren et al. @26# utilized

vortex generators, and Lake et al. @27# considered various passive

devices including dimples. Bons et al. @28# showed considerable

success using both steady and pulsed vortex generator jets. The

preceding list of studies, while long, is by no means inclusive. It is

merely a sample of recent work, biased toward the most recent

studies.

Much has been learned from the work to date, but the nature of

separated flow transition is still not completely clear, and existing

models are still not as robust as needed for accurate prediction.

The present study expands the existing database. The flow through

a single-passage cascade-simulator is documented under both high

and low FSTI conditions at several different Reynolds numbers.

The geometry of the passage corresponds to that of the ‘‘Pak-B’’

airfoil, which is an industry supplied research airfoil that is representative of a modern, aggressive LP turbine design. This geometry was used in several of the studies mentioned above ~@15–

17,21,23,24,27,28#!, and the pressure profile from the suction side

of this airfoil was matched in the flat plate study of Volino and

Hultgren @13#. Previous work has included documentation at Re

~based on suction surface length and exit free-stream velocity! as

low as 50,000 ~e.g., @13,16,21#!. The present work includes the

first complete documentation ~to the author’s knowledge! of cases

with Re as low as 25,000. Also new is documentation of the

turbulent shear stress in the boundary layer under both high and

low FSTI.

Details of the experimental facility and results of the study

follow. The present paper focuses on mean and statistical quantities. Part 2 of this work @29# includes turbulence spectral results,

providing evidence of the important transition mechanisms.

Experiments

Experiments were conducted in a low speed wind tunnel, described by Volino et al. @30#. Briefly, air enters through blowers

and passes through a series of screens, a honeycomb, two settling

chambers, and a three-dimensional contraction before entering the

test section. At the exit of the contraction, the mean velocity is

uniform to within 1%. The FSTI is 0.560.05%. Nearly all of this

free-stream ‘‘turbulence’’ is actually streamwise unsteadiness at

frequencies below 20 Hz and is not associated with turbulent eddies. The rms intensities of the three components of the unsteadiness are 0.7, 0.2, and 0.2% in the streamwise, pitchwise and spanwise directions, respectively. For low FSTI cases, the test section

immediately follows the contraction. For high FSTI, a passive

grid is installed at the contraction exit followed bya1m long

rectangular settling chamber. Details of the grid are available in

Volino et al. @30#. At the inlet to the test section the high FSTI

mean flow and turbulence are spatially uniform to within 3 and

6%, respectively. The free-stream turbulence is nearly isotropic

with rms intensities of 8.8, 8.9, and 8.3% in the streamise, pitchwise and spanwise directions. The integral length scales of these

components are 3, 1.6, and 1.4 cm. The integral scales were computed from the power spectra of each component.

The test section, shown in Fig. 1, consists of the passage between two airfoils. Details are listed in Table 1. Cascade simulators of this type have been used in studies such as Chung and

Simon @31#, more recently in the present facility by Aunapu et al.

@32#, and with the PakB geometry by Qiu and Simon @16# and

Kaszeta et al. @21#. A large span to chord ratio of 4.3 was chosen

to insure two-dimensional flow at the spanwise centerline of the

airfoils, where all measurements were made. Upstream of each

airfoil are flaps, which control the amount of bleed air allowed to

escape from the passage. These are adjusted to produce the correct

leading edge flow and pressure gradient along the airfoils. A tailboard on the pressure side of the passage also aids in setting the

pressure gradient.

Single passage test sections have several advantages. For a

given wind tunnel with fixed maximum flow rate, the single passage can be considerably larger than a passage in a multi-blade

facility. The larger size and simpler geometry can also result in

better probe access. Previous studies ~e.g., @31,32#! demonstrated

that the full flow field, including the three-dimensional secondary

flows near the endwalls, in a single passage can be set to match

that in a corresponding multi-blade cascade. The present test section also has some advantages over flat plate test sections. First,

the airfoil curvature is matched. Second, with an adverse pressure

gradient, suction is often needed to prevent separation on the wall

opposite a flat test plate ~e.g., Volino and Hultgren @13#!. A cascade simulator does not require suction due to the favorable pressure gradient on the pressure side of the passage.

Single passages also have disadvantages. It is, of course, impossible to establish periodicity. Stage losses cannot be directly

determined since there is flow only on one side of each airfoil and

the downstream wake is, therefore, unlike that in a multi-blade

facility. This limitation, however, is not prohibitive for the present

study. Primary concern is with boundary layer separation and transition, which occur in the passage. Although the downstream

wake may be different, the flow in the passage does match that of

a multi-blade facility.

Ten different cases have been documented including high and

low FSTI cases at five Reynolds numbers (Re525,000, 50,000,

100,000, 200,000, and 300,000!. The Reynolds number range is

representative of conditions from cruise to takeoff. The FSTI levels in an engine may vary considerably, but the values in the

present work are believed to span the range of most interest. Solomon @8# surveyed several studies that included wake effects and

Fig. 1 Schematic of the test section

Table 1 Test section parameters

646 Õ Vol. 124, OCTOBER 2002 Transactions of the ASME

NASA/CR—2012-217656 7
found FSTI values ranging from 1 to 5% between wakes and from

3 to 23% within wakes. Wakes can affect transition and separation

in the boundary layer in three ways. First, the FSTI rises during a

wake passage compared to the between-wake value. Second, a

calmed region follows wake induced transition. The calmed flow

is nonturbulent, but unlike a steady nonturbulent flow, it can be

very resistant to separation. Finally, independent of the FSTI effect, each wake includes a mean velocity deficit, resulting in temporal deceleration and acceleration as the wake passes. Lou and

Hourmouziadis @12# separated this temporal effect from the wake

turbulence effect, using downstream control to create an oscillating velocity in their test section. In the present study, only steady

flow is considered. While the significant effects of temporal acceleration and calming are not present, the high and low FSTI

cases of the present study do allow a means for evaluating the

effect of wake turbulence level.

Measurements. Pressure surveys were made for each case

using a pressure transducer ~0–870 Pa range Validyne transducer!

and a Scanivalve. Stagnation pressure was measured with a pitot

tube upstream of the passage inlet, and eleven pressure taps were

located on each airfoil along their spanwise centerlines. Locations

of the taps on the suction side are listed in Table 2 along with

measured local FSTI components, and the ReK product at these

stations based on a nonseparating, inviscid solution. The pressure

distribution on the upstream portion of the suction side always

closely matched the inviscid solution for flow over the airfoil.

This allowed the use of the measured static pressure at the third

pressure tap on the suction side, along with the inviscid flow

solution for the passage and the upstream stagnation pressure, to

determine the nominal passage exit velocity, which was used to

normalize the measured pressure distributions. More conventionally the measured inlet velocity and the inlet and exit flow angles

are used to compute the exit velocity. Because the velocity at the

third tap is 1.9 times that at the passage inlet, and therefore easier

to measure, using the third tap velocity reduced the bias uncertainty in the pressure coefficients, particularly at the lower Reynolds numbers. The uncertainty in the suction side pressure coefficients was 7% at the lowest Re, and below 4% in other cases.

Most of this uncertainty was due to bias error. Stochastic error

was minimized by averaging pressure transducer readings over a

10-s period.

Velocity profiles were measured at eleven streamwise stations

along the suction side at the locations given in Table 2. Profiles

were measured near but not at the spanwise centerline of the airfoil to insure that the pressure taps did not interfere with the

velocity measurements. Profiles were acquired with a hot-wire

anemometer ~TSI model IFA100! and a single sensor boundary

layer probe ~TSI model 1218-T1.5!. The sensor diameter is 3.8

mm, and the active length is 1.27 mm. At each measurement location, data were acquired for 26 seconds at a 20 kHz sampling

rate (219 samples!. All raw data were saved. The high sampling

rate provides an essentially continuous signal, which is needed for

intermittency and spectral post-processing. The long sampling

time results in low uncertainty in both statistical and spectral

quantities. Data were acquired at 60 wall normal locations in each

profile, extending from the wall to the free-stream, with most

points concentrated in the near wall region. The closest point was

0.1 mm from the wall, which corresponds to y/Ls50.0004 and

between 0.01 and 0.2 boundary layer thicknesses. Flow direction

in a separation bubble cannot be determined with a single-sensor

hotwire, but velocity magnitude can be measured and was found

to be essentially zero within the bubbles of the present cases.

Determining the direction was not, therefore, considered essential.

At locations where the boundary layer was attached, local wall

shear stress was computed from the near wall profile using the

technique of Volino and Simon @33#. Uncertainties in the mean

velocity are 3–5%, except in the very near wall region (y1,5)

where near-wall corrections ~Wills @34#! were applied to the mean

velocity. Uncertainties in the momentum and displacement thicknesses computed from the mean profiles are 10%. Uncertainty in

the shape factor, H, and the wall shear stress are both 8%.

The uncertainty in the fluctuating streamwise velocity is below

10%, except in the very near wall region, where spatial averaging

effects become important in some cases. Ligrani and Bradshaw

@35,36# showed that spatial averaging over the length of a hotwire sensor can result in low apparent u8. Their experiments were

done in a fully-turbulent boundary layer with Reu52600. The spatial averaging effects become important when the sensor length is

longer than the width of the smaller near wall streaks in a turbulent boundary layer. Ligrani and Bradshaw @35,36# found that the

spatial averaging effects become small when the dimensionless

sensor length, l ut /n, is less than about 25. The error also becomes smaller as the sensor is moved away from the wall. This is

expected since the average size of the turbulent eddies should

increase with distance from the wall. The Ligrani and Bradshaw

@35,36# results suggest that when the distance from the wall, y, is

larger than the sensor length, l , that spatial averaging errors are

under 10% even for large l ut /n. Closer to the wall they showed

errors in u8 as large as 30% when l ut /n560.

In the present study, l ut /n remains below 25 in all cases with

Re,200,000. Spatial averaging is not, therefore, expected to be a

problem, even near the wall. For the Re5200,000 cases, l ut /n is

above 25 at Station 11 of the low FSTI case and at Stations 9–11

of the high FSTI case, reaching values as high as 60. Spatial

averaging should not be significant for y.1 mm (y/Ls.0.004),

but may cause errors as high as 30% closer to the wall. It is not

certain that the errors are this large, however. The momentum

thickness Reynolds numbers in the present cases are all below

700, which is significantly below the Reu52600 value of the Ligrani and Bradshaw @35,36# study. This may indicate less developed turbulence in the present study, which could imply fewer

small-scale eddies and lower averaging errors. For the Re

5300,000 cases, l ut /n reaches values as high as 90 at Stations

10 and 11 of the low FSTI case and Stations 9-11 of the high FSTI

case. As in the Re5200,000 cases, errors should be small when

y.1 mm, but may be larger closer to the wall.

A boundary layer cross-wire probe ~TSI model 1243-T1.5! was

used to measure profiles of the wall normal velocity and turbulent

shear stress at Stations 7–11 for each case. The upstream boundary layer was too thin for cross-wire measurements. Data were

acquired at 25 locations in each profile, beginning 1 mm from the

wall and extending to the free-stream. Sampling rates and times

were the same as for the single sensor probe. Uncertainty in the

turbulent shear stress is 10%.

Measurements with the cross-wire probe are subject to spatial

averaging errors due to the length of the sensors ~1.27 mm active

length! and the spacing between the two sensors ~1 mm!. Applying the results of Ligrani and Bradshaw @35,36# and Ligrani et al.

@37# to the present cases, spatial averaging may be significant at

Table 2 Measurement stations locations, local acceleration

„inviscid soln...., and measured local free-stream turbulence

Journal of Turbomachinery OCTOBER 2002, Vol. 124 Õ 647

NASA/CR—2012-217656 8
locations very near the wall, particularly for the high Reynolds

number cases. At locations farther from the wall than 1 mm, however, the errors should become small and within the 10% uncertainty estimate given in the foregoing. Hence, no measurements

were made at y locations below 1 mm. Because all measurements

were at y.1 mm and a boundary layer type probe was used,

probe blockage effects were not expected to be significant.

The intermittency, g, is the fraction of time the flow is turbulent

within the transition region, and was determined at each measurement location using the technique described in Volino et al. @30#

with an uncertainty of 10%. Turbulent flow in the boundary layer

is defined here as flow which includes a range of large and small

scales, turbulence production, and dissipation. Using this definition, a boundary layer may be characterized by significant fluctuations, but still be nonturbulent if these fluctuations are induced

by an external source which does not cause near-wall turbulence

production. Such is often the case under high FSTI conditions.

Free-stream eddies ‘‘buffet’’ the boundary layer, inducing nonturbulent boundary layer fluctuations. Buffeting may occur through

pressure fluctuations. Boundary layer fluid is pushed in the wall

normal direction across the mean gradient in the streamwise velocity, resulting in significant u8 fluctuations. This type of motion

was termed ‘‘inactive’’ by Bradshaw @38# since it does not result

in momentum transport, in spite of potentially high u8 levels. In

addition to buffeting, some free-stream eddies may penetrate into

the boundary layer and cause some mixing. As described in Volino

@39#, the eddies which have the greatest effect, whether through

buffeting or penetrating the boundary layer, will be the larger,

energy containing eddies. The nonturbulent boundary layer subject to these external effects will be characterized by largeamplitude, low-frequency fluctuations. Transition to turbulence is

characterized not so much by large increases in u8 levels, which

may remain essentially constant, but by the appearance of higher

frequencies superimposed on the low frequencies. The higher frequencies signal the generation of turbulence in the near wall region. Volino et al. @30# provide examples of an intermittent flow

switching in time between disturbed nonturbulent and turbulent

states.

The presence or absence of high frequencies in a signal is used

to distinguish between turbulent and nonturbulent flow, using the

algorithm presented by Volino et al. @30#. The algorithm is similar

to others found in the literature. Briefly, the time derivative of a

signal is computed and compared to a threshold. Rapid ~highfrequency! fluctuations result in high derivatives. When the derivative is larger than the threshold, the flow is declared instantaneously turbulent and the intermittency function is assigned a

value of 1. When the derivative is below the threshold, the intermittency is assigned a value of 0. The time average of the function

is the intermittency, g. Volino et al. @30# showed that intermittency

can be computed based on u8 or u8v8 signals with essentially the

same result. Results based on u8 are presented in this paper.

Results

Pressure Profiles. Pressure coefficients for all ten cases are

shown in Fig. 2. At the upstream stations on the suction side, there

is good agreement between the data at all Reynolds numbers and

the inviscid flow solution. Separation appears to occur at s/Ls of

about 0.6 in all cases. These results agree with those of Volino and

Hultgren @13#, who also observed that the separation location did

not depend strongly on the Reynolds number or FSTI. They are in

contrast to other studies, such as Qiu and Simon @16#, which

showed that the separation location depended more strongly on

Re. Reattachment depends strongly on both Re and FSTI. With

low FSTI ~Fig. 2~a!!, the boundary layer appears to be separated

in all cases at Stations 8 and 9. It reattaches by Station 10 for the

Re5200,000 and 300,000 cases, reattaches by Station 11 for the

Re5100,000 case, and does not reattach at all for the Re

525,000 and 50,000 cases. Reattachment for the high FSTI cases

~Fig. 2~b!! occurs upstream of the low FSTI locations. The boundary layer appears to be separated in all cases at Station 8, but has

already reattached by Station 9 in the Re5200,000 and 300,000

cases. Reattachment has occurred by Station 10 for the Re

5100,000 case, and appears to be beginning at Station 11 for the

Re550,000 case. The boundary layer does not appear to reattach

when Re525,000, in spite of the high FSTI. Results for the low

FSTI cases are very similar to the flat plate results of Volino and

Hultgren @13#. The present high FSTI results appear to show about

a 10% larger separation region than the flat plate cases of @13#.

The differences are small and of the order of the resolution of the

measurement stations. Any differences between the studies are

presumably due to differences in the free-stream turbulence. Although the high FSTI in both studies was about 8%, the inlet

free-stream turbulence was more anisotropic in Volino and Hultgren @13#, and the integral length scales of the free-stream turbulence in @13# were about double those in the present study. The

larger length scale presumably caused earlier transition in the

Volino and Hultgren @13# study, resulting in a slightly shorter

separation bubble. Comparison to the high FSTI cases of Simon

et al. @17# shows reattachment about 14% farther upstream in @17#

than in the present study. With low FSTI, Simon et al. @17# did not

observe reattachment at all when Re5100,000, while it was observed in the present study. For the low FSTI Re5200,000 case,

they indicate reattachment about 6% farther upstream than the

present study. Although the streamwise pressure gradients were

Fig. 2 Cp profiles: „a... low FSTI, „b... high FSTI

648 Õ Vol. 124, OCTOBER 2002 Transactions of the ASME

NASA/CR—2012-217656 9
nominally the same in the present study and Simon et al. @17#,

small differences in the pressure gradients along with differences

in the intensity and length scales of the free-stream turbulence

were apparently responsible for the differences in reattachment

location.

Upstream Boundary Layer. The local free-stream velocities

at Station 1–5 for all 10 cases closely followed the equation

U`

Ue

51.48S s

LS

D

0.214

(1)

which corresponds to the free-stream velocity distribution for a

Falkner-Skan wedge flow. Figure 3 shows that the 50 mean velocity profiles from all 10 cases at these stations collapse onto the

same Falkner-Skan profile. Skin friction coefficients, shown in

Fig. 4, which were computed using the near wall profiles, also

follow the Falkner-Skan solution. There is no significant difference between the low and high FSTI cases. Agreement with the

flat plate data of Volino and Hultgren @13# is good. For the low

FSTI cases, it is not surprising that the laminar boundary layer

closely follows the expected laminar solution. Under the same

high FSTI inlet conditions, however, Volino et al. @30# showed

that a non-turbulent boundary layer may be strongly influenced by the free-stream turbulence and exhibit large deviation from laminar behavior. In the present study, the acceleration parameter, K, is

in some of the cases over 10 times larger than in the Volino et al.

@30# study, and the boundary layer thickness is as little as 1/5 that

in @30#. These differences apparently limit the free-stream effect

on the upstream boundary layer, resulting in the observed laminarlike behavior.

Figure 5 shows profiles of the rms fluctuating streamwise velocity, u8, for the 50 upstream profiles. The low FSTI u8 is mainly

streamwise unsteadiness that scales with the local U` . Values are

low everywhere, increasing slightly from the free-stream value to

a peak at y/u53 and then dropping to zero at the wall. For the

high FSTI cases the free-stream u8 level does not change significantly within the test section and scales with Ue . Since the

boundary layer fluctuations are caused by the free-stream fluctuations, the u8 profiles collapse when normalized on Ue . The collapse is not perfect since the free-stream u8 does drop somewhat

as the eddies are strained in the accelerating flow. Qualitatively

the behavior is the same as in the low FSTI cases, with high

values in the free-stream rising to a peak at y/u53 and dropping

to zero near the wall. The peak in u8 has lower magnitude and is

farther from the wall than would be expected in a turbulent

boundary layer. In all cases the boundary layer is clearly laminarFig. 3 Mean velocity profiles from Station 1–5, all cases

Fig. 4 Skin friction coefficients from Station 1–6, all cases

Fig. 5 Boundary layer u8 profiles from Stations 1–5—„a... low

FSTI cases, „b... high FSTI cases

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NASA/CR—2012-217656 10
like in spite of the high u8 level. This is shown in both the mean

velocity profiles ~Fig. 3! and the local skin friction coefficients

~Fig. 4! which follow laminar flow solutions.

Low FSTI Transition. Downstream of Station 6, the pressure

gradient becomes adverse and separation and transition occur.

Figures 6 through 10 show profiles at Stations 6–11 of U, u8, the

turbulent shear stress, 2u8v8, and the intermittency, g, for the

five low FSTI cases. The v8 profiles are qualitatively very similar

to the 2u8v8 profiles in all cases and are not shown. In the Re

525,000 case ~Fig. 6!, the mean velocity profile has just separated

at Station 7, and the separation bubble grows continuously larger

at the downstream stations. There is no reattachment, but rather a

massive separation with a burst bubble at the trailing edge. The

intermittency is nearly zero everywhere, indicating that the shear

layer remains essentially laminar. Although the flow is laminar, u8

is nonzero. The u8 peak grows as the flow moves downstream,

and its location is concurrent with the inflection point in the mean

profile. The 2u8v8 values remain near zero through Station 10,

but then rise at Station 11 with a peak in the shear layer at the

same location as the u8 peak. The magnitude of this peak is extremely low; the eddy viscosity at the peak is only about 1/30th of

the molecular kinematic viscosity. Although not significant in

terms of eddy transport, this peak may signify the beginning of

transition.

Figure 7 shows the profiles for the Re550,000 case. Results are

very similar to the Re525,000 case of Fig. 6. The boundary layer

does not reattach. The 2u8v8 profile at Station 11 again exhibits

a low level peak. The dimensionless value of this peak is about

three times that of the peak in Fig. 6, indicating that the Re

550,000 case may be closer to transitioning.

Results for the Re5100,000 case are shown in Fig. 8. The mean

velocity profiles show that the boundary layer is on the verge of

separating at Station 7, but is still attached. It has separated by

Station 8, and the separation bubble grows through Station 10,

although it does not become as thick as in the lower Re cases. At

Station 10, the mean velocity near the wall rises slightly above

zero, indicating the beginning of reattachment. At Station 11 the

boundary layer is clearly reattached. The intermittency is near

zero through Station 10, and then suddenly increases to 1 at Station 11, indicating fully turbulent flow as the boundary layer reattaches. The peak in the intermittency is well away from the wall,

indicating that transition begins in the shear layer over the separation bubble. The u8 profiles exhibit a peak in the shear layer at

Stations 8 and 9, similar to the behavior at the lower Re. At

Station 10 there is an increase in u8 near the wall as reattachment

begins. The turbulent shear stress profile rises above zero at Station 10 with a dimensionless value that is an order of magnitude

larger than the peak shown in Fig. 7 for the Re550,000 case. At

Station 11 the magnitude of the peak has increased by another

order of magnitude and the boundary layer is clearly turbulent.

The peak in 2u8v8 is well away from the wall, indicating that

while the boundary layer is turbulent and reattached, it has not yet

recovered to fully developed turbulent conditions.

Figure 9 shows the profiles for the Re5200,000 case. The mean

velocity profiles indicate that separation does not occur until near

Station 8. There is a clear separation bubble at Station 9, and the

boundary layer is reattached by Station 10. By Station 11 the

mean profile appears to have recovered to a fully developed turbulent shape. The intermittency jumps from near 0 at Station 9 to

1 at Station 10, indicating a rapid transition and reattachment. The

magnitude of u8 increases similarly, from a small peak near

the inflection point of the mean profile at Station 9 to high

values throughout the boundary layer at Station 10. The turbulent

shear stress profiles show the same sudden increase between

Stations 9 and 10, and the peak is still away from the wall at

Station 11, indicating that recovery from separation may not be

fully complete.

Fig. 6 Station 6–11 profiles for low FSTI, ReÄ25,000 case—„a...

mean velocity, „b... u8, „c... Àu8v8, „d... intermittency Fig. 7 Station 6–11 profiles for low FSTI, ReÄ50,000 case—„a...

mean velocity, „b... u8, „c... Àu8v8, „d... intermittency

Fig. 8 Station 6–11 profiles for low FSTI, ReÄ100,000 case—

„a... mean velocity, „b... u8, „c... Àu8v8, „d... intermittency

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NASA/CR—2012-217656 11
The Re5300,000 case profiles are shown in Fig. 10. Separation

again occurs near Station 8. By Station 9 the separation bubble

has clearly grown and the nonzero mean velocities near the wall

indicate that the boundary layer is on the verge of reattachment.

The u8 values increase greatly between Stations 8 and 9. The u8

profile has a peak in the shear layer over the separation bubble and

a second peak near the wall, which is indicative of reattachment.

The intermittency profile at Station 9 shows this same double

peak. The boundary layer is reattached and fully turbulent by Station 10. The turbulent shear stress profile rises to a small but

discernable nonzero level at Station 9, corresponding to the beginning of reattachment and the rise of the intermittency. By Station

11, 2u8v85ut

2 near the wall, which would be expected for a fully

developed, attached turbulent boundary layer.

The velocity data of Figs. 6–10 agree with the pressure profiles

of Fig. 2~a!. The separation locations agree, although the velocity

profiles provide better resolution and indicate that separation does

move downstream somewhat as Reynolds number increases. The

reattachment locations indicated by the pressure profiles correspond to locations where the velocity profiles have clearly reattached. Incipient reattachment is visible in the mean velocity one

station upstream of full reattachment in some cases. In all cases

the magnitude of u8 begins to increase in the shear layer after

separation. At first this increase in u8 occurs without a corresponding increase in 2u8v8, which remains near zero, and the

shear layer remains laminar. As the flow continues downstream,

low but nonzero 2u8v8 values eventually appear in the shear

layer, and at the higher Reynolds numbers this is quickly followed

by a rapid rise of 2u8v8, sudden transition to turbulence, and

almost immediate reattachment of the boundary layer. At the two

lowest Reynolds numbers, the initial rise in 2u8v8 was detected,

but it occurred so far downstream that transition and reattachment

never occurred. The present results are consistent with those of

Hatman and Wang @10#, Lou and Hourmouziadis @12#, and Volino

and Hultgren @13#, who also considered low FSTI separated flow

transition. They also reported rapid transition and attributed it to

the breakdown of a Kelvin-Helmholtz-type instability of the shear

layer. The transition mechanism in the present study will be discussed in more detail in Part 2 @29#.

High FSTI Transition. Profiles for the high FSTI cases are

shown in Figs. 11–15. For the Re525,000 case ~Fig. 11!, the

mean velocity profiles show that the boundary layer has separated

by Station 6. The boundary layer appears on the verge of reattaching at Stations 10 and 11, but is not clearly reattached. The u8

level rises rapidly after separation, with a peak in the shear layer

at each station. The free-stream buffets the shear layer, forcing

fluid across a large mean velocity gradient, dU/dy, which causes

high u8 levels. The same effect is present in the attached boundary

layer upstream ~Fig. 5b!, but is damped somewhat by the wall.

Free-stream buffeting and high u8 do not necessarily imply turbulent transport, and the turbulent shear stress remains low through

Station 9. The 2u8v8 level rises to high levels at Stations 10 and

11, but the peak is in the shear layer and drops to zero at the wall.

Perhaps at this very low Reynolds number, even significant transport in the shear layer is insufficient to promote full reattachment

of the boundary layer. Turbulent reattachment may be an intermittent phenomenon, related to and much like transition. At the lower

Re it may occur over an extended distance. The intermittency

indicates that the flow remains nonturbulent, in spite of the high

levels of 2u8v8 at the downstream stations. The intermittency

function, as defined above, only declares the flow turbulent when

Fig. 9 Station 6–11 profiles for low FSTI, ReÄ200,000 case—

„a... mean velocity, „b... u8, „c... Àu8v8, „d... intermittency

Fig. 10 Station 6–11 profiles for low FSTI, ReÄ300,000 case—

„a... mean velocity, „b... u8, „c... Àu8v8, „d... intermittency

Fig. 11 Station 6–11 profiles for high FSTI, ReÄ25,000 case—

„a... mean velocity, „b... u8, „c... Àu8v8, „d... intermittency, „e... v8

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NASA/CR—2012-217656 12
the velocity fluctuations include a full range of both large and

small scales. The apparent mismatch between the 2u8v8 and g

profiles of Fig. 11 may indicate that the fluctuations which cause

the turbulent shear stress initially do not include this range of

scales. Turbulent shear stress spectra are presented in Part 2 @29#.

Profiles of the wall normal fluctuating velocity, v8, are also shown

in Fig. 11. At Stations 7–9, high free-stream values drop to zero at

the wall, with no peak corresponding to the peak in u8. The freestream buffeting effect on v8 is damped by the wall, a phenomenon also observed in attached, nonturbulent boundary layers under high FSTI conditions ~Volino et al. @30#!. A peak emerges in

v8 in the shear layer at Station 10 and 11, corresponding completely with the rise in 2u8v8 at these stations. The link between

v8 and 2u8v8 was clear at all Re, making it unnecessary to

present both v8 and 2u8v8 for the remaining cases.

The Re550,000 case of Fig. 12 is very similar to the Re

525,000 case. Reattachment is clearer, however, at Station 11,

and the intermittency is nonzero at this station. Figure 13 shows

the Re5100,000 case. In this case the intermittency indicates that

transition has begun by Station 9, which corresponds to an initial

rise in 2u8v8. The mean profile shows that the boundary layer is

reattached at the last two stations. The high 2u8v8 peaks away

from the wall indicate that the boundary layer has not fully recovered from the separation at Station 11.

It is not clear that separation occurs in the Re5200,000 case

~Fig. 14!, but the velocity mean profile at Station 8 has an inflection point and appears to be close to separating. By Station 9 the

boundary layer is clearly attached, the intermittency indicates

transition is underway, u8 is high even near the wall, and 2u8v8

has risen to a turbulent level. At Station 11, 2u8v8 reaches a

maximum equal to ut

2 near the wall, indicating a fully developed

attached turbulent flow.

The Re5300,000 results of Fig. 15 are very similar to those at

Re5200,000. Transition begins slightly earlier at Re5300,000,

with the intermittency greater than zero at Station 8. The thin

boundary layer at Re5300,000 results in peaks in 2u8v8 at Stations 9 and 10 that are too close to the wall to resolve with the

cross-wire probe.

In general, transition in the high FSTI cases began upstream of

the locations in the corresponding low FSTI cases, and the transition region length was longer with high FSTI. This agrees with

the observations of Volino and Hultgren @13#, who also observed

Fig. 12 Station 6–11 profiles for high FSTI, ReÄ50,000 case—

„a... mean velocity, „b... u8, „c... Àu8v8, „d... intermittency

Fig. 13 Station 6–11 profiles for high FSTI, ReÄ100,000

case—„a... mean velocity, „b... u8, „c... Àu8v8, „d... intermittency

Fig. 14 Station 6–11 profiles for high FSTI, ReÄ200,000

case—„a... mean velocity, „b... u8, „c... Àu8v8, „d... intermittency

Fig. 15 Station 6–11 profiles for high FSTI, ReÄ300,000

case—„a... mean velocity, „b... u8, „c... Àu8v8, „d... intermittency

652 Õ Vol. 124, OCTOBER 2002 Transactions of the ASME

NASA/CR—2012-217656 13
that transition was less abrupt with high FSTI. Boundary layer u8

levels are much higher with high FSTI, but the turbulent shear

stress magnitude remains low until transition begins and does not

appear to depend strongly on FSTI.

Shape Factor and Momentum Thickness. As discussed

above, stage losses cannot be determined quantitatively using a

single passage test section, but is possible to compute the momentum thickness of the suction side boundary layer at the trailing

edge. As explained by Howell et al. @20#, this momentum thickness is proportional to the suction side profile loss when the

boundary layer shape factor and passage exit angle remain constant. For those cases in which the boundary layer reattaches and

recovers to a fully developed turbulent shape, H is approximately

equal to 1.4. In these cases the suction side profile loss is likely

the dominant loss mechanism ~Howell et al. @20#!. In those cases

in which the boundary layer does not fully reattach, or reattaches

near the trailing edge, the shape factor will be very large due to a

large displacement thickness. The momentum thickness may be

relatively small compared to the displacement thickness in these

cases, since the wall shear is essentially zero and u will not grow

significantly in the free-shear layer. Large losses would then be

expected in the wake, downstream of the airfoil.

Figure 16 shows the shape factor at Station 11 (s/Ls50.94) as

a function of Re. In the Re5200,000 and 300,000 cases, the shape

factor is approximately 1.4, indicating that the boundary layer is

reattached and that the momentum thickness is a good indicator of

overall losses. For the lower Re cases, H is significantly higher,

particularly in the low FSTI cases. The lower H in the high FSTI

cases indicates that high FSTI helps to keep the separation bubble

thinner by promoting more mixing in the shear layer over the

bubble and by inducing earlier transition and reattachment. Figure

17 shows the Station 11 momentum thickness as a function of Re.

High FSTI helps keep the separation bubble thinner, as shown

above in the mean velocity profiles of Figs. 6–15, which tends to

result in lower u and lower losses when the boundary layer reattaches. High FSTI also promotes increased mixing, however,

which tends to increase u. For the Re5300,000 cases, Fig. 17

indicates that the second effect is more significant and u is higher

for the high FSTI case. Transition and reattachment occur sufficiently far upstream in the low FSTI case that the effect of the

high FSTI in promoting even earlier reattachment is not enough to

counter the enhanced mixing effects. This suggests that small,

controllable separation bubbles may be acceptable or even desirable in some cases, as proposed by Hourmouziadis @1# for controlled diffusion blading. For the Re5200,000 and 100,000 cases,

the earlier reattachment caused by high FSTI is more significant

and u is lower for the high FSTI cases. This result is consistent

with the observation in several studies that unsteady wakes

from upstream airfoils result in lower losses. Enhanced mixing in

the shear layer explains the higher u for the high FSTI cases at

Re525,000. These momentum thicknesses do not relate directly

to losses since the shear layer does not reattach in either of the

Re525,000 cases.

Comparison to Correlations. Hatman and Wang @10# discuss three modes of separated flow transition. Based on their

criteria, the present cases all fall into their laminar-separation

long-bubble category. The data support this; separation occurred

before transition. Hatman and Wang @10#, Mayle @2# and others

suggest that the Thwaites @40# criteria, Reu

2

K520.082, is a good

predictor for laminar separation. A laminar, attached flow solution would put Reu

2

K520.082 between Stations 6 and 7 in all

cases of the present study. The presence of the separation bubble

changes the local acceleration, however, which tends to move the

location where Reu

2

K520.082. Volino and Hultgren @13# found

the Thwaites criteria to be a good predictor of separation and it

appears to work well for the present study as well. Exact prediction of the separation point is not straightforward, however, due to

the interdependence of local K values and the separation bubble

location.

Prediction of transition and reattachment is more difficult. Hatman and Wang @10# present a transition correlation based on low

FSTI data which predicts that transition should not occur in any of

the present cases. Clearly, however, transition and reattachment do

occur. The 2u8v8 profiles indicate that transition is imminent

even in the low FSTI, low Re cases. Volino and Hultgren @13#

drew comparisons to correlations from Mayle @2# and Davis et al.

@41# with mixed results. Comparisons to the present data are similarly mixed. The Mayle correlations predict the distance from the

separation point to the onset of transition based on Reu at the

separation location. He presents a correlation for short separation

bubble length and a correlation for long bubble length, which is

3.3 times the short bubble length. The Davis et al. @41# correlation

also predicts the distance from separation to transition onset, but

as a function of the FSTI. Table 3 presents the distance from

separation to the start of transition, normalized on the suction

surface length, for all cases of the present experiments and as

predicted by the Mayle @2# and Davis et al. @41# correlations. The

finite spacing between the measurement stations results in uncertainty in Reu at separation and in the exact locations of separation

and transition, so a range of values is given for each quantity in

Fig. 16 Shape factor at Station 11 Fig. 17 Momentum thickness at Station 11

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NASA/CR—2012-217656 14
the table. The results of the present cases lie between the Mayle

long and short bubble correlations to within the resolution of the

measurement locations. The Davis et al. correlation tends to predict too long a distance for the low FSTI cases and too short a

distance for the high FSTI cases.

Existing correlations appear to give reasonable rough estimates

of separated flow transition in some cases, but they are not particularly accurate or robust predictors. The general agreement between the similar cases of the present study, Volino and Hultgren

@13#, and Simon et al. @17# suggests that prediction of separated

flow transition should be possible to some extent. The differences

between the results of these studies, noted above, suggest that

very accurate prediction of the flow may prove difficult and

strongly dependent on small differences in boundary conditions. It

is doubtful that a simple, robust correlation can be developed to

incorporate all relevant boundary condition effects and provide

very accurate predictions. It is also questionable whether the

boundary conditions could be specified accurately enough for actual engine conditions. Perhaps they can, with addition research,

however, and it may be possible to improve predictions with advanced computational schemes. The difficulty of predicting transition suggests that it may be advantageous to develop flow control schemes to force transition to occur at desired locations rather

than try to predict it under existing conditions.

Conclusions

Separated flow transition has been documented for cases with

Reynolds numbers ranging from 25,000 to 300,000 at both high

and low FSTI. The following conclusions can be drawn from the

results.

1 The start of boundary layer reattachment occurs near the start

of transition, and both depend strongly on Re and FSTI.

2 High FSTI results in competing effects with regard to losses.

It causes enhanced mixing, which tends to promote boundary

layer growth and increase losses. At the same time, it promotes

earlier transition and reattachment, which reduces boundary layer

thickness and losses. At the highest Re, reattachment occurred

shortly after separation regardless of FSTI level, and high FSTI

resulted in higher losses. At the intermediate Re, high FSTI reduced losses. At the lowest Re, the boundary layer did not reattach

even with high FSTI, so losses would be high regardless of FSTI

level.

3 The turbulent shear stress level can remain near zero in spite

of high FSTI and high u8 in the boundary layer. The beginning of

a rise in 2u8v8 signals the beginning of transition. In the lowest

Re cases, transition did not occur and the boundary layer did not

reattach, but the beginning of a rise in 2u8v8 was observed near

the trailing edge. This hints that it may be possible to induce

transition even at very low Re.

4 The present results agree roughly with similar studies from

the literature and existing correlations but there are significant

differences. Attempts to control transition and force its location

may prove more fruitful than prediction of unmodified flow. The

present study provides an extensive, detailed baseline data set for

ongoing flow control experiments.

Acknowledgments

This work was sponsored by the NASA Glenn Research Center.

The grant monitor is Dr. David Ashpis. Additional matching support was provided through a U.S. Naval Academy Recognition

Grant. Mr. Dale Boyer of the Technical Support Department at the

Naval Academy fabricated the airfoils for the test section.

Nomenclature

Cf 5 skin friction coefficient

Cp 5 2(PT2P)/rUe

2

, pressure coefficient

FSTI 5 free-stream turbulence intensity

H 5 d*/u, shape factor

K 5 (n/U`

2 )(dU` /ds), acceleration parameter

Ls 5 suction surface length

l 5 hot-wire sensor length

P 5 pressure

PT 5 upstream stagnation pressure

Re 5 UeLs /n, exit Reynolds number

Res 5 U`s/n, local Reynolds number

Reu 5 momentum thickness Reynolds number

s 5 streamwise coordinate, distance from leading edge

U 5 mean streamwise velocity

U` 5 local free-stream velocity

Ue 5 nominal exit free-stream velocity, based on inviscid

solution

u8 5 rms streamwise fluctuating velocity

ut 5 At w /r, friction velocity

2u8v8 5 time averaged turbulent shear stress

v8 5 rms wall normal fluctuating velocity

y 5 cross-stream coordinate, distance from wall

y1 5 yut /n, distance from wall in wall coordinates

d* 5 displacement thickness

g 5 intermittency, fraction of time flow is turbulent

n 5 kinematic viscosity

r 5 density

t w 5 wall shear stress

u 5 momentum thickness

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a Transitional Boundary Layer Under High Free-Stream Turbulence Conditions,’’ ASME Paper 2001-GT-0192.

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Blade/Endwall Junction of a Gas Turbine: Visualization in a Large-Scale Cascade Simulator,’’ ASME Paper 90-WA/HT-4.

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of Turbulence in the Viscous Sublayer Using Subminiature Hot-Wire Probes,’’

Exp. Fluids, 5, pp. 407–417.

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262–268.

@38# Bradshaw, P., 1994, ‘‘Turbulence: the Chief Outstanding Difficulty of Our

Subject,’’ Exp. Fluids, 16, pp. 203–216.

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Separation Bubbles on Infinite Swept Wings,’’ AIAA Paper 85-1685.

Journal of Turbomachinery OCTOBER 2002, Vol. 124 Õ 655

NASA/CR—2012-217656 16
Ralph J. Volino

Mem. ASME,

Department of Mechanical Engineering,

United States Naval Academy,

Annapolis, MD 21402

e-mail: volino@usna.edu

Separated Flow Transition Under

Simulated Low-Pressure Turbine

Airfoil Conditions—Part 2:

Turbulence Spectra

Spectral analysis was used to investigate boundary layer separation, transition and reattachment under low-pressure turbine airfoil conditions. Cases with Reynolds numbers

ranging from 25,000 to 300,000 (based on suction surface length and exit velocity) have

been considered at low (0.5%) and high (9% inlet) free-stream turbulence levels. Spectra

of the fluctuating streamwise velocity and the turbulent shear stress are presented. The

spectra for the low free-stream turbulence cases are characterized by sharp peaks. The

high free-stream turbulence case spectra exhibit more broadband peaks, but these peaks

are centered at the same frequencies observed in the corresponding low turbulence cases.

The frequencies of the peaks suggest that a Tollmien-Schlichting instability mechanism

drives transition, even in the high turbulence cases. The turbulent shear stress spectra

proved particularly valuable for detection of the early growth of the instability. The

predictable nature of the instability may prove useful for future flow control work.

@DOI: 10.1115/1.1506939#

Introduction

In Part 1 of the present study @1#, the significance of boundary

layer separation, transition, and reattachment to the flow over

modern low-/pressure turbine airfoils was discussed. Measured

mean velocity and statistical turbulence quantities were presented

for cases with high and low free-stream turbulence intensity

~FSTI! and Reynolds numbers ~based on suction surface length

and exit velocity! ranging from 25,000 to 300,000. The separation

point tended to move downstream somewhat as Reynolds number

increased. Transition and reattachment locations moved upstream

significantly as Re or FSTI were increased.

While the statistical quantities presented in Part 1 @1# provide a

quantitative description of what happens under different Re and

FSTI conditions, they do not explain the transition mechanism. To

better explain the transition process and accurately predict or control it, an understanding of the physics which cause the results

observed in Part 1 @1# is needed. The present paper uses spectral

analysis to investigate separated flow transition.

Mayle @2# classified the modes of transition as ‘‘natural transition,’’ ‘‘bypass’’ transition; ‘‘separated flow’’ transition of the

shear layer over a separation bubble; ‘‘periodic-unsteady’’ transition, which might also be called wake-induced bypass transition;

and reverse transition. Under low FSTI, zero streamwise pressure

gradient conditions, natural transition is expected. This type of

transition has been extensively documented and can be predicted

with linear stability analysis. As described by Schlichting @3#,

when the displacement thickness Reynolds number exceeds a

critical value, the boundary layer becomes unstable to small disturbances, which begin to grow as Tollmien-Schlichting ~TS!

waves. These waves eventually become three dimensional and

result in turbulent spots. Under high FSTI, zero pressure gradient

conditions, large disturbances can cause a bypass of the linear

growth stages of transition, resulting in the sudden appearance of

turbulent spots. At intermediate FSTI, elements of both bypass

and natural transition may be observed. Sohn and Reshotko @4#,

for example, presented data for a 1% FSTI case, showing both

spectral peaks at possible TS frequencies and broadband unsteadiness more typical of bypass transition.

In some cases, transition is observed even though linear stability theory predicts that the boundary layer should not develop TS

waves. Volino @5#, for example, considered a favorable pressure

gradient case with high FSTI that clearly underwent transition.

The boundary layer thickness remained low in this case due to the

acceleration, resulting in Red* below the critical limit for linear

instability.

Separated flow transition could potentially include elements of

either natural or bypass transition. In separated flow cases the

pressure gradient is adverse, resulting in a boundary layer or shear

layer that typically is unstable to TS waves. High FSTI, however,

might be the dominant factor in a separated shear layer, overwhelming the effect of any TS waves and producing bypass transition. Hughes and Walker @6# list several studies with FSTI below

0.9% in which TS waves were detected in adverse pressure gradient cases. They also note that Halstead et al. @7# did not detect

TS waves in the flow through a rotating cascade with more representative, higher FSTI. Solomon and Walker @8#, however, provide evidence of TS waves under conditions similar to those of

Halstead et al. @7#. Hughes and Walker @6# considered a flow with

wakes, in which the FSTI between wakes ranged from less than

1% to about 3%, and the FSTI in the wakes was about 8%. They

provide clear evidence of TS waves.

Hatman and Wang @9#, Volino and Hultgren @10#, and Lou and

Hourmouziadis @11# all considered low FSTI, adverse pressure

gradient flows and observed transition in the shear layer over

separation bubbles. Spectral data in all three studies showed clear

evidence of an instability along with harmonics. It was expected

that this instability was very similar to the Kelvin-Helmholtz instabilities observed in free shear layers, although the unstable frequencies were somewhat different than expected for free shear

layers since the separation bubbles were bounded by the wall on

one side. Volino and Hultgren @10# also considered high FSTI

cases and observed broadband unsteadiness in the spectra of the

streamwise fluctuating velocity, u8. Spikes at discreet frequencies,

which were observed in the low FSTI cases, were not present.

They stated that transition in the high FSTI cases appeared to be

through a bypass mode.

Contributed by the International Gas Turbine Institute and presented at the International Gas Turbine and Aeroengine Congress and Exhibition, Amsterdam, The

Netherlands, June 3–6, 2002. Manuscript received by the IGTI, January 22, 2002.

Paper No. 2002-GT-30237. Review Chair: E. Benvenuti.

656 Õ Vol. 124, OCTOBER 2002 Copyright © 2002 by ASME Transactions of the ASME

NASA/CR—2012-217656 17
Clearly there is some disagreement regarding the transition

mechanism in separated boundary layers, particularly under high

FSTI conditions. Some of these differences may stem from physical differences in the boundary conditions between the various

studies. The present study addresses the issue through spectral

analysis of flows over a range of Reynolds numbers at both high

and low FSTI. Included in the analysis are spectra of the turbulent

shear stress, which were not considered in previous studies.

Experiments

The experimental facility and the cases considered are described in detail in Part 1 @1#. A low-speed wind tunnel supplies

air to a single-passage cascade-simulator with geometry and flow

angles matching those for the industry supplied Pak-B airfoil. For

the low FSTI cases, the background turbulence level for the wind

tunnel is nominally 0.5%, and consists primarily of low frequency

unsteadiness. A passive grid is used to generate a high inlet FSTI

of 8.7%.

Velocity data were acquired at 11 streamwise measurement stations along the spanwise centerline of the suction side of the passage. Station locations are given in Table 1. At each station, instantaneous streamwise velocity was measured at 60 locations as a

single sensor hot-wire probe was traversed from the airfoil surface

to the free-stream. The voltage from the hot-wire was offset and

amplified by a factor of 10 and low pass filtered at 10 kHz using

signal conditioners ~TSI model 157!. At each location, data were

acquired for 26 s at a 20 kHz sampling rate (219 samples!. The

high sampling rate provides an essentially continuous signal,

which is needed for spectral processing. The long sampling time

results in low uncertainty in both statistical and spectral quantities.

Two component velocity measurements were made at Stations

7–11 with a cross-wire probe. The upstream boundary layer was

too thin for cross-wire measurements. Data were acquired at 25

locations in each profile, beginning 1 mm from the wall and extending to the free-stream. Sampling rates and times were the

same as with the single sensor probe. Power spectra of u8, the

wall normal fluctuating velocity, v8, and the turbulent shear

stress, 2u8v8, were computed for the data from all measurement

locations. Uncertainties in the u8, v8 and 2u8v8 spectra are all

10%. Frequencies are resolved from 4.88 to 10 kHz in 4.88 Hz

increments using a 4096 point Fast Fourier Transform to compute

the spectra. As a check, the spectra were integrated with respect to

frequency and found to equal the corresponding time averaged

Reynolds stresses.

Data sets were acquired for cases at high and low FSTI with

exit Reynolds numbers of 25,000, 50,000, 100,000, 200,000 and

300,000. The upstream boundary layer through Station 6 remained

laminar in all of these cases, with the mean velocity following a

Falkner-Skan wedge flow solution. Downstream of Station 6 the

pressure gradient becomes adverse. Separation occurred between

Stations 6 and 8, with the separation point moving downstream as

Re increased. Reattachment did not occur in the high or low FSTI,

Re525,000 cases or the low FSTI, Re550,000 case. In all other

cases the boundary layer did reattach, and the reattachment point

moved upstream as Re or FSTI was increased. Details are available in Part 1 @1#. The u8 level rose in the shear layer in each case

after the boundary layer separated, but the turbulent shear stress

remained near zero until transition began. Transition began in the

shear layer and quickly led to boundary layer reattachment. In the

Re525,000 cases, low but non-zero turbulent shear stress was

observed at the most downstream stations, indicating that although the shear layer was still non-turbulent and separated, it

was showing signs of the start of transition.

As discussed in Part 1 @1#, the finite length of the hot-wire

sensors ~1.27 mm! will result in some spatial averaging and could

result in attenuation of the measured fluctuating velocity components. Based on the results of Ligrani and Bradshaw @12,13#, it

was explained in Part 1 that the errors in the rms fluctuating quantities are within the 10% uncertainty estimates in the majority of

cases in the present study. Exceptions occur for the Re5200,000

cases at Station 11 of the low FSTI case and Stations 9–11 of the

high FSTI case. For the Re5300,000 cases, larger errors are expected for Stations 10–11 of the low FSTI case and Stations 9–11

of the high FSTI case. For these cases, errors may be as large as

30% near the wall, but should be under 10% at y locations greater

than 1 mm. As explained by Ligrani et al. @14#, measurements

with the cross-wire probe are subject to potentially larger errors

due to the finite spacing ~1 mm! between the two sensors. These

errors will be largest near the wall, but become smaller than the

10% uncertainty for y locations above 1 mm. For this reason,

cross-wire measurements were only made for y.1 mm.

The spectra in the present paper are presented to show the energy content of the fluctuating quantities as a function of frequency. The average errors in these spectra, therefore, should be

the same as those given in the foregoing for the corresponding

rms quantities. The errors will not be uniform with respect to

frequency across the spectra, however. As explained by Ligrani

and Bradshaw @13#, spatial averaging effects will be most severe

for the smallest scales ~highest frequencies! in the flow. Applying

the spectral results of Ligrani and Bradshaw @13# to the present

study, errors due to spatial averaging at y locations above 1 mm

will rise above 10% at frequencies above 150, 300, 600, 1200, and

1800 Hz for the Re525,000, 50,000, 100,000, 200,000, and

300,000 cases respectively. Below these frequencies the errors

should be under 10%. These frequencies are all above the frequencies of the spectral peaks in the results presented below, so the

peaks should not be significantly attenuated. For the Re525,000

and 50,000 cases, all significant energy in the spectra is below the

frequencies given above, so there is no significant attenuation of

the results at any frequency. For the Re5100,000 case, only the

high frequency ‘‘tail’’ of the spectra will be subject to significant

error. Magnitudes at these high frequencies may be as much as

30% low. Similarly for the Re5200,000 and 300,000 cases, errors

will only rise above 10% at frequencies about twice those of the

spectral peaks. Magnitudes at the highest frequencies may be as

much as 40% low in these cases.

Frequencies of Interest

Transition in the shear layer over the separation bubble could

be initiated by Tollmien-Schlichting waves originating in the

boundary layer upstream of separation. Upstream of Station 6, the

pressure gradient is favorable and the boundary layer should not

develop TS waves at any frequencies. At Station 6 the flow begins

to decelerate and immediately becomes unstable. Walker @15# provides the following equation for the frequency of maximum amplification rate for TS waves:

2pn f /U`

2 53.2 Red *

23/2 (1)

The TS frequencies predicted by Eq. ~1! are listed in Table 2.

Values are given for Stations 6 and 7 for the cases in which the

boundary layer is still attached at these stations.

Another possible path to transition is breakdown of the shear

layer though a Kelvin-Helmholtz-type instability. The frequency

of the instability should scale with the velocity change across the

shear layer and inversely with the shear layer thickness. Since the

velocity in the separation bubble is nearly zero, the velocity difference across the shear layer equals the local free-stream velocity. The shear layer thickness, d S , is determined from the mean

Table 1 Measurement stations

Journal of Turbomachinery OCTOBER 2002, Vol. 124 Õ 657

NASA/CR—2012-217656 18
velocity profiles presented in Part 1 @1#. Table 2 lists the quantity

U` /d S for each station where the boundary layer is separated.

Results

Free-Stream Spectra. The u8 and v8 free-stream spectra are

shown in Fig. 1 for each station of the low FSTI, Re5300,000

case. Frequency is plotted on a log scale versus frequency times

power spectral density on a linear scale. In these coordinates the

area under the curve in any frequency band is proportional to the

contribution to the quantity of interest in that band. The u8 spectra

are dominated by low amplitude unsteadiness at frequencies below 20 Hz. The u8 spectra for the lower Re cases ~not shown!

have proportionately lower amplitudes, but the frequency range

remains the same. The frequencies associated with turbulent eddies would be expected to scale with the free-stream velocity.

Since the frequencies in the present cases remain constant as Re is

changed, the unsteadiness in the wind tunnel is most likely not

associated with turbulent eddies. The v8 spectra show the same

low frequencies as u8 but at 1/10th the magnitude. A second lower

peak is centered at about 100 Hz. This peak is also present with

the same magnitude in u8, and is visible in an expanded version

of Fig. 1~a!. The frequency of this second peak scales with the

free-stream velocity and is proportionately lower for the lower Re

cases. Its magnitude is very low, and by itself would correspond to

an rms turbulence level of about 0.01%. It is likely the residual

turbulence remaining after the wind tunnel screens.

The u8 and v8 free-stream spectra for the high FSTI Re

5300,000 case are shown in Fig. 2. The magnitude of u8 is about

25 times larger than in the low FSTI case of Fig. 1~a!. At the

upstream stations there is a peak at about 30 Hz. This peak decays

due to streamwise straining of the flow as it is accelerated through

Station 6. In the adverse pressure gradient region downstream

of Station 6, a broadband peak emerges centered at about 700 Hz.

In the lower Re cases, the u8 spectra is qualitatively the same, but

Fig. 1 Free-stream spectra for low FSTI, ReÄ300,000 case—

„a... u8, „b... v8

Fig. 2 Free-stream spectra for high FSTI, ReÄ300,000 case—

„a... u8, „b... v8

Table 2 Most unstable Tollmien-Schlichting frequencies prior

to separation and U ` Õd S values in shear layer

658 Õ Vol. 124, OCTOBER 2002 Transactions of the ASME

NASA/CR—2012-217656 19
the magnitudes scale with U`

2 and the frequencies scale with

U` . The v8 spectra of Fig. 2~b! all appear similar with a broadband peak centered at 60 Hz. As with u8, the magnitudes and

frequencies of the v8 spectra in the lower Re cases scale with the

free-stream velocity, and appear qualitatively similar to those of

Fig. 2~b!.

Upstream Boundary Layer. Contours of the boundary layer

u8 spectra from Station 4 of the low FSTI Re5100,000 case

are shown in Fig. 3. Frequency in Hz is shown on the horizontal

axis on a log scale, and distance from the wall normalized on

the suction surface length is on the vertical axis on a linear scale.

The frequency is left dimensional since there is no single appropriate normalization for all regions of the flow. The contours

in Fig. 3 show the dimensionless magnitude of the spectra as f

•PSD(u8

2)/U`

2 . A slice through the data of Fig. 3 at a fixed

distance from the wall would produce a spectrum in the coordinates of Fig. 2, except with dimensionless magnitude. The number

~1.46e-05!, which appears in the center of the figure, indicates the

magnitude of the outermost contour and the contour spacing.

Hence, the outermost contour value is 1.4631025, the next contour value is 2.9231025, the next is 4.3831026, etc. The same

format is used in all the figures which follow. Figure 3 shows a

peak near the wall centered at 50 Hz and unsteadiness below 20

Hz extending from the free-stream to near the wall. The results

shown in Fig. 3 are typical of the behavior at Stations 1–6 in all

the low FSTI cases. The 50 Hz peak was also visible in the Re

5200,000 and 300,000 cases, but not at the two lower Re. This

peak did not appear to have any significant effect on the downstream boundary layer.

Figure 4 shows the spectra from Station 4 of the high FSTI

Re5300,000 case in the coordinates of Fig. 3. The Fig. 4 spectra

are typical of the spectra at Stations 1–6 of all the high FSTI

cases. There is a near wall peak at 100 Hz, which is 1.7 times the

frequency of the free-stream v8 frequency peak shown in Fig.

2~b!. The distance of the peak from the wall is larger in the lower

Re cases, scaling with the boundary layer thickness. The frequency of the peak scales with Re, and is consistently 1.7 times

the dominant v8 frequency in the free-stream. This indicates that

the boundary layer unsteadiness is induced by free-stream buffeting. Volino @5# discusses free-stream buffeting of boundary layers

in more detail.

Low FSTI Transition. Upstream of Station 6, the boundary

layer unsteadiness is attributable to low amplitude streamwise unsteadiness in the low FSTI cases, and free-stream buffeting in the

high FSTI cases. Buffeting, as explained in Part 1 @1#, refers to the

effect of free-stream pressure fluctuations on the boundary layer

as fluid is pushed in the wall normal direction across the gradient

in the mean streamwise velocity. Downstream of Station 6, the

pressure gradient becomes adverse, and the spectra become more

interesting. Figure 5 shows the u8 spectra contours for Stations

7–11 of all the low FSTI cases. The coordinates of each subplot

are the same as those of Fig. 3. In the Re525,000 case ~top row of

Fig. 5!, the contours at Station 7 show the low frequency unsteadiness observed at the upstream stations. Between Stations 7 and 10,

the magnitude of the peak increases by an order of magnitude.

The peak at each station moves away from the wall. Comparison

to the mean velocity profiles presented in Part 1 @1# shows that the

location of the peak corresponds, not surprisingly, to the shear

layer over the separation bubble. The Re550,000 case ~row 2!

shows similar behavior through Station 10, but a second peak

emerges at about 78 Hz at Station 11. This second peak is indicative of transition.

In the Re5100,000 case ~row 3 of Fig. 5!, a sharp peak appears

at Station 10 at 273 Hz. The contour spacing increases by two

orders of magnitude between Stations 9 and 10, indicating a similar increase in the magnitude of the peak. The peak is so sharp and

so large that it appears as a line in the contour plot. Figure 6

shows this Re5100,000, Station 10 data in the same coordinates

as Fig. 5, but three dimensionally. The low frequency unsteadiness

visible at Station 9 is still present, but since its magnitude is only

1/50th that of the 273 Hz peak, it is barely visible in Fig. 6. The

sharp peak in the shear layer is typical of all the low FSTI cases.

In the Re5200,000 case ~row 4 of Fig. 5!, a small peak appears

in the shear layer at 698 Hz at Station 9. By Station 10 this peak

has increased in size by two orders of magnitude and is becoming

more broadband as the boundary layer reattaches and becomes

turbulent. Although not clear in Fig. 5, the sharp spectral peak

remains in the center of this broadband turbulence and is clear in

the format of Fig. 6. Similar behavior is apparent in the Re

5300,000 case, but the spectral peak and subsequent turbulence

are centered at 922 Hz.

Contours of the normalized turbulent shear stress spectra, f

•PSD(2u8v8)/U`

2 , for the low FSTI cases are shown in Fig. 7.

The format is the same as in Fig. 5. The shear layer is laminar at

Stations 7 and 8 in all cases, and the turbulent shear stress is near

zero. The Re525,000 case shows a low magnitude, 15 Hz peak at

Station 10. The peak was obscured in the u8 spectra of Fig. 5 due

to the presence of other streamwise unsteadiness at similar frequencies. The magnitude of the 15 Hz shear stress peak increases

Fig. 3 Contours of fPSD„u8

2..., low FSTI, ReÄ100,000, Station

4; outer contour magnitude and contour spacing indicated by

numerical value in field of figure

Fig. 4 Contours of fPSD„u8

2..., high FSTI, ReÄ300,000, Station

4; outer contour magnitude and contour spacing indicated by

numerical value in field of figure

Journal of Turbomachinery OCTOBER 2002, Vol. 124 Õ 659

NASA/CR—2012-217656 20
by a factor of 15 between Station 10 and 11, but its amplitude is

still quite low. Close inspection of the u8 spectra ~Fig. 5! at Station 11 shows a double peak, with one peak at 15 Hz. The mean

profiles of Part 1 @1# indicate that transition may be imminent, but

the shear layer is still laminar at Station 11.

The Re550,000 case shows similar behavior to that at Re

525,000. A turbulent shear stress peak appears at Station 10 and

is 37 times larger by Station 11. This peak did not become visible

until Station 11 in the u8 spectra. Similarly, in the Re5100,000

case a sharp turbulent shear stress peak emerges at Station 9, one

station upstream of its appearance in u8. The shear stress peak

appears at Station 9 in the Re5200,000 and 300,000 cases. In all

cases, the peak appears at the same frequency in u8 and 2u8v8,

but tends to become visible earlier in the 2u8v8 spectra due to

the lower magnitude of the low-frequency ‘‘noise’’ in 2u8v8.

The frequencies of the spectral peaks are listed for all cases in

Table 3. Also indicated in the table are the stations at which these

frequencies were determined. The frequencies of the peaks did not

change significantly in the streamwise direction. Because the

peaks do not appear until after the boundary layer has separated,

there is some reason to believe that a Kelvin-Helmholtz-type inFig. 5 Contours of f"PSD„u8

2...ÕU `

2 , low FSTI cases; station number indicated above each column; by row—„a... Re

Ä25,000, „b... ReÄ50,000, „c... ReÄ100,000, „d... ReÄ200,000, „e... ReÄ300,000; outer contour magnitude and contour spacing

indicated by numerical value in field of each subplot, 5 contours shown in each subplot

Fig. 6 Spectra of f"PSD„u8

2...ÕU `

2 , low FSTI, ReÄ100,000,

Station 10

Table 3 Measured peak frequencies in spectra

660 Õ Vol. 124, OCTOBER 2002 Transactions of the ASME

NASA/CR—2012-217656 21
stability may be involved. For the low FSTI cases, comparison of

the spectral peak frequencies to the U` /d S values in Table 2

shows that the spectral frequencies and U` /d S values both increase with Re, as expected. The ratio of the measured frequency

to U` /d S is not a constant, however, ranging from about 0.04 to

about 0.08. Comparison of the measured frequencies to the TS

frequencies of Table 2 shows better correlation. At the three highest Re, the measured frequencies agree with the TS frequencies at

Station 7 ~the last station before separation! to within 20%. The

boundary layer is already separated at Station 7 in the Re

525,000 and 50,000 cases, so the frequency comparison must be

made at Station 6. Agreement with the TS frequencies is not as

good in these cases as at the high Re. Hughes and Walker @6# note

that since the most unstable TS frequency changes with streamwise position, the frequency observed in the shear layer need not

equal the TS value at any single upstream position.

High FSTI Transition. Contours of the u8 spectra for the

high FSTI cases are shown in Fig. 8. In the Re525,000 case ~top

row!, there is low frequency unsteadiness at the same frequencies

observed upstream ~Fig. 4!. As in the low FSTI cases, the peak u8

location is in the shear layer over the separation bubble. In the

Re550,000 case, the outermost contour at Station 10 extends to

higher frequencies than at the upstream stations, and by Station 11

a new peak has emerged at 60 Hz, extending from the shear layer

down to the wall. The magnitude of this new peak is about equal

to the magnitude of the original, low-frequency peak, which is

also still visible. Similar behavior is clear for the Re5100,000

case. Higher frequencies begin to emerge at Station 9, and a large,

clear peak centered at 260 Hz is visible at Stations 10 and 11. In

the Re5200,000 and 300,000 cases, some signs of higher frequencies are already visible at Station 8, and a double peak is clear at

Station 9. At Stations 10 and 11, the higher frequency peak overwhelms the lower frequencies.

Comparing Fig. 8 to Fig. 5, the contour levels in the high FSTI

case are about 2 orders of magnitude higher than in the low FSTI

case at Stations 7 and 8. This is expected and due to the higher u8

caused by free-stream buffeting in the high FSTI case. Farther

downstream, however, after the higher frequencies emerge, the

contour levels are higher for the low FSTI cases. This is somewhat misleading, as the rms u8 levels are actually very similar in

the high and low FSTI cases at these stations. The lower contour

levels in the high FSTI case result because the u8 fluctuations are

distributed over a wider frequency band. The spectral peaks of

Fig. 8 are much broader than the corresponding peaks of the low

FSTI cases shown in Figs. 5–7. Figure 9 shows the u8 spectra at

Station 10 of the high FSTI, Re5100,000 case. Comparing to Fig.

6, the high FSTI case exhibits much more low frequency activity

due to free-stream buffeting, and the peak is clearly broader in

frequency than the 273 Hz spike of the low FSTI case.

Figure 10 shows contours of the turbulent shear stress spectra

for the high FSTI cases. Values at Stations 7 and 8 are low, in

spite of the high FSTI, indicating that much of the u8 unsteadiness

observed in Fig. 8 does not involve turbulent transport. Distinct,

broadband peaks emerge by Station 9 in all cases. As in the low

Fig. 7 Contours of f"PSD„Àu8v8...ÕU `

2 , low FSTI cases; see Fig. 5 caption for further explanation

Journal of Turbomachinery OCTOBER 2002, Vol. 124 Õ 661

NASA/CR—2012-217656 22
FSTI cases, these peaks are visible in 2u8v8 farther upstream

than they are in u8. The peak becomes visible in 2u8v8 soon

after it forms, but the u8 peak must grow to become larger than

the free-stream induced fluctuations before it is discernable.

The broad peaks in the high FSTI cases might suggest that

transition occurs through a bypass mode. This was the conclusion

of Volino and Hultgren @10#. Closer inspection, however, reveals

strong similarity to the low FSTI cases. The frequencies of the

peaks in the high FSTI cases are listed in Table 3. The broad

nature of the high FSTI case peaks, as shown in Fig. 9, result in an

uncertainty of about 20% in the frequency values in Table 3. With

this uncertainty, the frequency at each Re is essentially the same

as in the corresponding low FSTI case. Comparison to the TS

frequencies in Table 2 show the same good agreement observed in

the low FSTI cases. Close inspection of Fig. 10 reveals that at the

four highest Re, the peak in 2u8v8 begins to appear, with very

low magnitude, at Stations 7 and 8. The boundary layer is still thin

at these stations, so part of the peak lies closer to the wall than can

be measured with the cross-wire probe. The closest measurement

to the wall with the cross-wire was at y51 mm, which corresponds to y/d 99.5 between 0.2 and 0.65 at the stations in question.

What is visible, however, is at the same frequencies as present

downstream. Since the boundary layer is still attached at Station 7,

it is doubtful these fluctuations could be induced by a KelvinHelmholtz-type instability. Instead, it appears that a TS instability

plays a role, even in the high FSTI cases. This agrees with the

findings of Hughes and Walker @6#, who used instantaneous wall

shear measurements to identify wave packets both within and between wakes in an unsteady flow. Although it appears that TS

waves play a role in the high FSTI cases, it should be noted that it

is also possible that transition occurs though a bypass mode, and

Fig. 8 Contours of f"PSD„u8

2...ÕU `

2 , high FSTI cases; see Fig. 5 caption for further explanation

Fig. 9 Spectra of f"PSD„u8

2...ÕU `

2 , high FSTI, ReÄ100,000,

Station 10

662 Õ Vol. 124, OCTOBER 2002 Transactions of the ASME

NASA/CR—2012-217656 23
that the spectral peaks are due not to TS waves, but to the turbulence within turbulent spots. Perhaps both TS and bypass modes

play a role, as suggested by Mayle @2#.

Conclusions

Turbulence spectra provide valuable insights into the transition

mechanism. Clear sharp peaks were observed in the spectra of the

low FSTI cases, at frequencies indicating a probable TS instability

mechanism for the breakdown of the shear layer over the separation bubble to turbulence. The spectra were more broadband for

the high FSTI cases, but the peaks of these spectra occurred at

the same frequencies as in the corresponding low FSTI cases,

suggesting a possible similar transition mechanism at high and

low FSTI. The turbulent shear stress spectra were valuable for

detection of instabilities upstream of the location where peaks

became discernable in the u8 spectra. In the high FSTI cases,

low magnitude peaks were detected in 2u8v8 upstream of separation, further supporting the argument for a TS transition. Larger

initial disturbances in the high FSTI cases resulted in detectable

2u8v8 farther upstream than in the low FSTI cases. Since the

adverse pressure gradient boundary layer is unstable to disturbances over a broad range of frequencies and the free-stream turbulence contains a range of scales, broad peaks emerge in the

boundary layer spectra of the high FSTI cases. These peaks contrast with the sharp peaks of the low FSTI cases, which result

from the growth of small disturbances at the most unstable frequencies. The presence and predictability of the shear layer instability may prove useful in future attempts to induce transition for

separation control.

Acknowledgments

This work was sponsored by the NASA Glenn Research Center.

The grant monitor is Dr. David Ashpis. Additional matching support was provided through a U.S. Naval Academy Recognition

Grant.

Nomenclature

FSTI 5 free-stream turbulence intensity

f 5 frequency in Hz

Ls 5 suction surface length

PSD 5 power spectral density of u8

2 or 2u8v8

Re 5 UeLs /n, exit Reynolds number

Red* 5 U`d*/n displacement thickness Reynolds number

s 5 streamwise coordinate, distance from leading edge

U` 5 local free-stream velocity

Ue 5 nominal exit free-stream velocity

u8 5 streamwise fluctuating velocity

2u8v8 5 turbulent shear stress

v8 5 wall normal fluctuating velocity

y 5 cross-stream coordinate, distance from wall

d* 5 displacement thickness

Fig. 10 Contours of f"PSD„Àu8v8...ÕU `

2 , high FSTI cases; see Fig. 5 caption for further explanation

Journal of Turbomachinery OCTOBER 2002, Vol. 124 Õ 663

NASA/CR—2012-217656 24
d S 5 shear layer thickness

d 99.5 5 99.5% boundary layer thickness

n 5 kinematic viscosity

References

@1# Volino, R. J., 2002, ‘‘Separated Flow Transition Under Low Pressure Turbine

Airfoil Conditions: Part 1 - Mean Flow and Turbulence Statistics,’’ ASME J.

Turbomach., 124, pp. 645–655.

@2# Mayle, R. E., 1991. ‘‘The Role of Laminar-Turbulent Transition in Gas Turbine Engines,’’ ASME J. Turbomach., 116, pp. 509–537.

@3# Schlichting, H., 1979, Boundary Layer Theory, 7th ed., McGraw Hill, New

York, NY.

@4# Sohn, K. H., and Reshotko, E., 1991, ‘‘Experimental Study of Boundary Layer

Transition with Elevated Freestream Turbulence on a Heated Flat Plate,’’

NASA CR 187068.

@5# Volino, R. J., 2002, ‘‘An Investigation of the Scales in Transitional Boundary

Layers Under High Free-Stream Turbulence Conditions,’’ ASME Paper GT2002-30233.

@6# Hughes, J. D., and Walker, G. J., 2001, ‘‘Natural Transition Phenomena on an

Axial Compressor Blade,’’ ASME J. Turbomach., 123, pp. 392–401.

@7# Halstead, D. E., Wisler, D. C., Okiishi, T. H., Walker, G. J., Hodson, H. P., and

Shin, H.-W., 1997, ‘‘Boundary Layer Development in Axial Compressors and

Turbines: Part 3 of 4—LP Turbines,’’ ASME J. Turbomach., 119, pp. 225–237.

@8# Solomon, W. J., and Walker, G. J., 1995, ‘‘Incidence Effects on Wake-Induced

Transition on an Axial Compressor Blade,’’ Proc. 12th Int. Symposium on Air

Breathing Engines, Melbourne, Australia, pp. 954–964.

@9# Hatman, A., and Wang, T., 1999, ‘‘A Prediction Model for Separated Flow

Transition,’’ ASME J. Turbomach., 121, pp. 594–602.

@10# Volino, R. J., and Hultgren, L. S., 2001, ‘‘Measurements in Separated and

Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions,’’ ASME J. Turbomach., 123, pp. 189–197.

@11# Lou, W., and Hourmouziadis, J., 2000, ‘‘Separation Bubbles Under Steady and

Periodic-Unsteady Main Flow Conditions,’’ ASME J. Turbomach., 122, pp.

634–643.

@12# Ligrani, P. M., and Bradshaw, P., 1987, ‘‘Spatial Resolution and Measurement

of Turbulence in the Viscous Sublayer Using Subminiature Hot-Wire Probes,’’

Exp. Fluids, 5, pp. 407–417.

@13# Ligrani, P. M., and Bradshaw, P., 1987, ‘‘Subminiature Hot-Wire Sensors:

Development and Use,’’ J. Phys. E, 20, pp. 323–332.

@14# Ligrani, P. M., Westphal, R. V., and Lemos, F. R., 1989, ‘‘Fabrication and

Testing of Subminiature Multi-Sensor Hot-Wire Probes,’’ J. Phys. E, 22, pp.

262–268.

@15# Walker, G. J., 1989, ‘‘Transitional Flow on Axial Turbomachine Blading,’’

AIAA J., 27, pp. 595–602.

664 Õ Vol. 124, OCTOBER 2002 Transactions of the ASME

NASA/CR—2012-217656 25
Ralph J. Volino

Department of Mechanical Engineering,

United State Naval Academy,

Annapolis, MD 21402

e-mail: volino@usna.edu

Passive Flow Control on

Low-Pressure Turbine Airfoils

Two-dimensional rectangular bars have been used in an experimental study to control

boundary layer transition and reattachment under low-pressure turbine conditions. Cases

with Reynolds numbers (Re) ranging from 25,000 to 300,000 (based on suction surface

length and exit velocity) have been considered at low (0.5%) and high (8.5% inlet)

free-stream turbulence levels. Three different bars were considered, with heights ranging

from 0.2% to 0.7% of suction surface length. Mean and fluctuating velocity and intermittency profiles are presented and compared to results of baseline cases from a previous

study. Bar performance depends on the bar height and the location of the bar trailing

edge. Bars located near the suction surface velocity maximum are most effective. Large

bars trip the boundary layer to turbulent and prevent separation, but create unnecessarily

high losses. Somewhat smaller bars had no immediate detectable effect on the boundary

layer, but introduced small disturbances that caused transition and reattachment to move

upstream from their locations in the corresponding baseline case. The smaller bars were

effective under both high and low free-stream turbulence conditions, indicating that the

high free-stream turbulence transition is not simply a bypass transition induced by the

free stream. Losses appear to be minimized when a small separation bubble is present, so

long as reattachment begins far enough upstream for the boundary layer to recover from

the separation. Correlations for determining optimal bar height are presented. The bars

appear to provide a simple and effective means of passive flow control. Bars that are large

enough to induce reattachment at low Re, however, cause higher losses at the highest Re.

Some compromise would, therefore, be needed when choosing a bar height for best overall performance. @DOI: 10.1115/1.1626685#

Introduction

Modern low-pressure turbine ~LPT! airfoils are subject to increasingly stronger pressure gradients as designers impose higher

loading in an effort to improve efficiency and lower cost by reducing the number of airfoils in an engine. If the adverse pressure

gradient on the suction side of these airfoils becomes strong

enough, the boundary layer will separate. Separation bubbles, particularly those that fail to reattach, can result in a significant loss

of lift and a subsequent degradation of engine efficiency ~e.g.,

Hourmouziadis @1#, Mayle @2#, and Sharma et al. @3#!. The problem is particularly relevant in aircraft engines. Airfoils optimized

to produce maximum power under takeoff conditions may still

experience boundary layer separation at cruise conditions, due to

the thinner air and lower Reynolds numbers at altitude. A component efficiency drop of 2% may occur between takeoff and cruise

conditions in large commercial transport engines, and the difference could be as large as 7% in smaller engines operating at

higher altitudes. Component life may also be affected by more

than an order of magnitude ~Hodson @4#!. Because the LPT produces the bulk of the net power in many engines, changes in its

component efficiency can result in nearly equal changes in overall

engine efficiency ~Wisler @5#!. There are several sources for losses

in an engine, including secondary flows, but the suction side

boundary layer has been identified as the primary source of losses

in the LPT ~Curtis et al. @6#!. Prediction and control of suction

side separation, without sacrifice of the benefits of higher loading,

are therefore, necessary for improved engine design.

Separation on LPT airfoils is complicated by boundary layer

transition. Turbulent boundary layers are much more resistant to

separation than laminar boundary layers. If transition occurs far

enough upstream, it can prevent separation. If transition occurs in

the shear layer over a separation bubble, it will tend to induce

boundary layer reattachment. The lower the Reynolds number, the

farther downstream transition will tend to occur and hence the

problems associated with performance at altitude.

Separated flow transition has been studied extensively, and in

recent years several studies have focused on transition in the LPT.

Volino @7# provides a review of much of that work. Separation can

be affected through naturally occurring phenomena in an engine

and through deliberate attempts at flow control. Several studies

have shown that high free-stream turbulence intensity ~FSTI!

tends to cause the transition to move upstream, resulting in a

smaller separation bubble. Reducing the separation bubble size

tends to result in thinner boundary layers after reattachment,

thereby reducing losses. Moving the transition upstream, however,

results in a longer turbulent region on the airfoil, which tends to

increase losses. Volino @7# showed that the net result of these

competing effects depends on the Reynolds number. High FSTI

tends to reduce losses at low Re. At high Re, where separation

bubbles are relatively small even with low FSTI, high FSTI results in higher losses. At very low Re, boundary layers may fail to

reattach even with high FSTI ~e.g., Volino @7# and Van Treuren

et al. @8#!. Unsteadiness caused by wakes generated upstream of

an airfoil has been shown in several studies ~e.g., Howell et al.

@9#! to reduce the extent of separation bubbles and reduce losses.

As with elevated FSTI, wake unsteadiness is most effective at

reducing losses at lower Re, where the steady flow separation

bubbles are largest. Stadtmu ̈ller et al. @10# found that at high Re,

losses were higher with wakes than in steady flow.

Existing results suggest that separation bubbles should be kept

small, but without producing an unnecessarily long turbulent region. Hourmouziadis @1# discussed ‘‘controlled diffusion blading,’’ in which an airfoil is designed so that a small separation

bubble is present. The bubble itself is not thick enough to produce

high losses, and its presence allows a shorter turbulent region near

the trailing edge. This idea is discussed below in conjunction with

the present results.

Capitalizing on the beneficial effects of unsteady wakes, Howell et al. @9# and Brunner et al. @11# studied airfoils modified for

Contributed by the International Gas Turbine Institute and presented at the International Gas Turbine and Aeroengine Congress and Exhibition, Atlanta, GA, June

16–19, 2003. Manuscript received by the IGTI Dec. 2002; final revision Mar. 2003.

Paper No. 2003-GT-38728. Review Chair: H. R. Simmons.

754 Õ Vol. 125, OCTOBER 2003 Copyright © 2003 by ASME Transactions of the ASME

NASA/CR—2012-217656 26
higher lift. Losses increased with airfoil loading, as adverse pressure gradients became stronger and separation bubbles became

larger. With wake passing, however, the magnitude of the loss

increase was in some cases relatively small compared to the increase in lift. Aft loaded airfoils tended to have lower losses, since

separation and transition occurred closer to the trailing edge, resulting in a shorter turbulent region.

While high FSTI and wakes help to mitigate separated flow

problems, they clearly do not solve all problems, as evidenced by

the known efficiency drop in modern engines at altitude. Howell

et al. @9# indicated that their highly loaded airfoils might be close

to a limit, and that higher loading could cause unacceptable separation problems even in the presence of wakes. Looking beyond

FSTI and wakes, other types of flow control could prove useful.

The literature contains numerous examples of separation control.

Most have been applied to external flows over aircraft, but a few

studies have considered passive devices added to LPT airfoils.

Van Treuren et al. @8# utilized vortex generators on the suction

surface of an LPT airfoil. The vortex generators caused reattachment at Re550,000 ~all Re in the present paper are based on exit

velocity and suction surface length!. Losses appeared to be

slightly lower with the vortex generators. The vortex generators

were not effective at Re525,000, and the boundary layer did not

reattach even with 8% FSTI. Van Treuren et al. @8# did not consider higher Re. In another study, Lake et al. @12# used various

passive devices including dimples and boundary layer trips in an

LPT cascade. They considered cases with Re above 100,000. Murawski and Vafai @13# added extensions to the trailing edges of the

airfoils in their cascade. These extensions tended to move the

separation location downstream. At low Re, they reduced the

length of the separation bubble and reduced losses. At high Re,

losses increased. Byerley et al. @14# used ‘‘Gurney flaps’’ to control separation. These devices were trips, near the trailing edge on

the pressure side of the airfoils. They helped to keep the boundary

layer attached on the suction side, but also increased losses in the

cascade. Active separation control has also been employed. Bons

et al. @15,16# used steady and pulsed vortex generator jets to successfully control separation under LPT conditions.

The studies listed above indicate that separation control should

be possible under LPT conditions. Existing results are, however,

limited both in the range of Reynolds numbers considered in each

study and in the types of data acquired. More experiments are

needed with various types of devices to expand the experimental

data base. Detailed measurements will also help in the explanation

of the physical mechanisms by which various devices affect the

flow.

Passive flow control is considered in the present work. Thin

bars of rectangular cross section are placed on the suction surface

of an LPT airfoil near the suction surface velocity peak. Experiments were conducted in a single-passage cascade simulator, described in Volino @7#. The geometry of the passage corresponds to

that of the ‘‘Pak-B’’ airfoil, which is an industry supplied research

airfoil that is representative of a modern, aggressive LPT design.

Volino @7# documented cases in the present facility without flow

control. These serve as baseline cases for the present study.

Experiments

Experiments were conducted in a low speed wind tunnel, described by Volino et al. @17#. Briefly, air enters through blowers

and passes through a honeycomb, a series of screens, two settling

chambers, and a three-dimensional contraction before entering the

test section. At the exit of the contraction, the mean velocity is

uniform to within 1%. The FSTI is 0.5%60.05%. Nearly all of

this free-stream ‘‘turbulence’’ is actually streamwise unsteadiness

at frequencies below 20 Hz and is not associated with turbulent

eddies. The rms intensities of the three components of the unsteadiness are 0.7%, 0.2%, and 0.2% in the streamwise, pitchwise,

and spanwise directions, respectively. For low-FSTI cases, the test

section immediately follows the contraction. For high FSTI, a

passive grid is installed at the contraction exit followed by a 1-mlong rectangular settling chamber. At the inlet to the test section

the high-FSTI mean flow and turbulence are spatially uniform to

within 3% and 6%, respectively. The free-stream turbulence is

nearly isotropic with rms intensities of 8.8%, 8.9%, and 8.3% in

the streamise, pitchwise, and spanwise directions. The integral

length scales of these components are 3 cm, 1.6 cm and 1.4 cm.

The integral scales were computed from the power spectra of each

component.

The test section, shown in Fig. 1, consists of the passage between two airfoils. Details are listed in Table 1 and more information is available in Ref. @7#. A large span-to-chord ratio of 4.3

was chosen to ensure two-dimensional flow at the spanwise centerline of the airfoils, where all measurements were made. Upstream of each airfoil are flaps, which control the amount of bleed

air allowed to escape from the passage. The flaps, along with a

tailboard on the pressure side of the passage, are adjusted to produce the correct leading edge flow and pressure gradient along the

airfoils. The flow in the passage matches that in a multiblade

cascade.

Experimental conditions match those of the ten baseline cases

of Volino @7#, who considered high- and low-FSTI cases at five

Reynolds numbers ~Re525,000, 50,000, 100,000, 200,000, and

300,000!. The Reynolds number range is representative of condiFig. 1 Schematic of the test section

Table 1 Test section parameters

Axial

chord

~mm!

True

chord

~mm!

Pitch

~mm!

Span

~mm!

Suction

side, Ls

~mm!

Inlet

flow

angle

Exit

flow

angle

153.6 170.4 136.0 660.4 228.6 35° 60°

Journal of Turbomachinery OCTOBER 2003, Vol. 125 Õ 755

NASA/CR—2012-217656 27
tions from cruise to takeoff. The FSTI levels in an engine may

vary considerably, but the values in the present work are believed

to span the range of most interest.

Prior to the detailed experiments of the present study, various

devices were used in preliminary attempts at flow control. The

devices included trip wires of various diameters, rectangular bars

of various widths and thicknesses, and delta wing vortex generators of various heights, spacing, and angles with respect to the

flow. All of these devices were tried at several streamwise locations along the suction surface. Documentation included streamwise pressure profiles and velocity profiles acquired near the trailing edge. Large devices of any type eliminated separation ~as

indicated by the pressure profiles!, but caused large increases in

losses ~as indicated by large increases in momentum deficit at the

trailing edge!. As the size of the devices was reduced, it was found

that all devices that were just large enough to induce boundary

layer reattachment at Re525,000 caused about the same increase

in losses at higher Re. This was somewhat unexpected, as it was

thought that the delta wings vortex generators might present less

blockage, and more effectively promote mixing and inhibit separation than the trips or bars. Reasons for this unexpected finding

are discussed with the results below. Since no device appeared to

have a clear advantage, rectangular bars were chosen for further

study because of their simplicity. The bars were of uniform rectangular cross section and extended along the airfoil span, as

shown in Fig. 2. It should be noted that the tests described above

were not exhaustive, and do not preclude the possible usefulness

of vortex generators or other types of devices.

The results of the preliminary tests with the bars indicated that

the streamwise width of a bar and the location of its leading edge

were unimportant. The bar height and the location of its trailing

edge were critical. Hence, it appeared that the backward-facing

step at the trailing edge was most important for flow control. Bars

were most effective when the trailing edge was near the location

of the suction surface velocity peak. If the trailing edge was much

farther downstream, it was located under the separation bubble

and was ineffective. If the trailing edge was upstream in the favorable pressure gradient region, the stabilizing effect of the accelerating flow appeared to lessen the bar’s effectiveness.

In the present study, rectangular bars were fabricated from multiple layers of vinyl tape. The trailing edge of the bar was located

at s/Ls50.51, near the suction surface velocity peak. All bars

were 6 mm wide in the streamwise direction. Bar heights of 0.4

mm, 0.8 mm, and 1.6 mm were used. The bar heights were all less

than 1% of Ls . They compare to local boundary layer thickness at

the bar location of about 3.8 mm, 2.7 mm, 2.0 mm, 1.4 mm, and

1.2 mm in the baseline Re525,000 through 300,000 cases, respectively. For each bar height, all 10 cases of the baseline study were

redocumented, for a total of 30 new experimental cases.

Measurements. Pressure surveys were made for each case

using a pressure transducer ~0–870 Pa range Validyne transducer!

and a Scanivalve. Stagnation pressure was measured with a pitot

tube upstream of the passage inlet, and 11 pressure taps were

located on each airfoil along their spanwise centerlines. Locations

of the taps on the suction side are listed in Table 2 along with

measured local FSTI components, and the Re K product at these

stations based on a nonseparating, inviscid solution. The uncertainty in the suction side pressure coefficients was 7% at the lowest Re, and below 4% in other cases. Most of this uncertainty was

due to bias error. Stochastic error was minimized by averaging

pressure transducer readings over a 10-s period.

Velocity profiles on the suction surface were measured at

streamwise stations corresponding to pressure taps 7–11, as given

Fig. 2 Scale drawing of suction side airfoil showing location

of bar

Table 2 Measurement stations locations, local acceleration „inviscid solution..., and measured

local free-stream turbulence

Station s/Ls Re K

Low

FSTI

u8/U`

~%!

Low

FSTI

v8/U`

~%!

High

FSTI

u8/U`

~%!

High

FSTI

v8/U`

~%!

1 0.111 1.58 0.44 5.2

2 0.194 1.20 0.39 4.6

3 0.278 0.86 0.37 4.0

4 0.361 0.75 0.38 3.5

5 0.444 0.62 0.39 3.2

6 0.528 20.02 0.41 2.8

7 0.611 20.81 0.47 0.05 2.9 5.9

8 0.694 20.95 0.47 0.12 3.0 6.2

9 0.777 20.58 0.48 0.14 3.4 6.6

10 0.861 20.53 0.54 0.11 3.8 6.8

11 0.944 20.18 0.51 0.11 4.0 6.8

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NASA/CR—2012-217656 28
in Table 2. These stations are downstream of the bar. Profiles at

stations 1–6 are fully documented for the baseline cases in Refs.

@7,18#, and show that the upstream boundary layer closely follows

a laminar solution, even in the high-FSTI cases. Profiles were

measured near but not at the spanwise centerline of the airfoil to

insure that the pressure taps did not interfere with the velocity

measurements. Profiles were acquired with a hot-wire anemometer ~AA Lab Systems model AN-1003! and a single-sensor

boundary layer probe ~TSI model 1218-T1.5!. The sensor diameter is 3.8 mm, and the active length is 1.27 mm. At each measurement location, data were acquired for 26 s at a 20-kHz sampling rate (219 samples!. All raw data were saved. The high

sampling rate provides an essentially continuous signal, which is

needed for intermittency and spectral postprocessing. The long

sampling time results in low uncertainty in both statistical and

spectral quantities. Data were acquired at 60 wall normal locations

in each profile, extending from the wall to the free stream, with

most points concentrated in the near-wall region. The closest point

was within 0.1 mm of the wall, which corresponds to y/Ls

50.0004 and between 0.01 and 0.2 boundary layer thicknesses.

Flow direction in a separation bubble cannot be determined with a

single-sensor hot-wire, but velocity magnitude can be measured

and was found to be essentially zero within the bubbles of the

present cases. Determining the direction was not, therefore, considered essential. Uncertainties in the mean velocity are 3–5%

except in the very near wall region where near-wall corrections

~Wills @19#! were applied to the mean velocity. Uncertainties in

the momentum and displacement thicknesses computed from the

mean profiles are 10%. Uncertainty in the shape factor, H, is 8%.

The uncertainty in the fluctuating streamwise velocity is below

10%, except in the very-near-wall region, where spatial averaging

effects, due to the finite length of the hot-wire sensor, become

important in some cases. For the present cases, as explained by

Volino @7# based on the work of Ligrani and Bradshaw @20,21#,

spatial averaging should not be significant for the Re525,000 and

50,000 cases, even near the wall. For the higher Re cases, spatial

averaging should not be significant for y.1 mm (y/Ls.0.004),

but may cause errors as high as 30% closer to the wall. It is not

certain that the errors are this large, however. The estimates are

based on the results of Ligrani and Bradshaw @20,21#, who considered a boundary layer with Reu52600. The momentum thickness Reynolds numbers in the present cases are all below 1300.

This may indicate less developed turbulence in the present study,

which could imply fewer small-scale eddies and lower averaging

errors.

The intermittency, g, is the fraction of time the flow is turbulent

within the transition region. It was determined at each measurement location based on the instantaneous streamwise velocity signal, using the technique described by Volino et al. @17#. The uncertainty in g is 10%. As explained by Volino et al. @17#, turbulent

flow is defined here to include a range of large- and small-scale

eddies, turbulence production, and dissipation. A boundary layer

may be characterized by significant u8 fluctuations but still be

nonturbulent if these fluctuations are induced by an external

source that does not also cause near-wall turbulence production.

Such is often the case under high-FSTI conditions. Free-stream

eddies buffet the boundary layer, inducing nonturbulent boundary

layer fluctuations but very little momentum transport. Transition

to turbulence is characterized not so much by large increases in u8

levels, which may remain essentially constant, but by the appearance of higher frequencies. The higher frequencies signal the generation of turbulence in the near-wall region and are used to distinguish between turbulent and nonturbulent flow. Further

discussion is available in Ref. @17#.

Results

Velocity and Pressure Profiles. Pressure coefficients for the

low-FSTI, Re525,000 cases are shown in Fig. 3. Also shown is

the inviscid solution for the present geometry. In all cases there is

good agreement with the inviscid solution on the pressure side.

On the suction side, the baseline case shows good agreement with

the inviscid solution in the favorable pressure gradient region, but

a large separation bubble in the adverse pressure gradient region.

Separation is indicated by the nearly constant Cp values, which

are well above the invisid solution. The Cp values remain high to

the trailing edge, showing no sign of reattachment. With the 0.4-

mm-thick bar, there is an increase in Cp over the baseline value at

s/Ls50.53. The pressure tap at this location is immediately downstream of the bar, and the flow over the tap is probably affected by

the close proximity of the bar. The Cp values in this case remain

high to the trailing edge, indicating that the boundary layer does

not reattach. The same is true for the 0.8-mm bar case. With the

1.6-mm bar, Cp drops below the baseline values near the end of

the favorable pressure gradient region. The larger bar is apparently

enough of an obstruction to slow the near-wall flow upstream of

the bar. Downstream of the 1.6-mm bar, Cp values are high, as in

the other cases, but at the most downstream pressure tap Cp drops

to near the inviscid solution value, indicating boundary layer reattachment.

The velocity profiles for the low-FSTI, Re525,000 cases are

shown in Fig. 4. The top row of the figure shows dimensionless

mean velocity profiles at stations 7–11. The baseline case shows a

boundary layer near separation at station 7, a small separation

bubble at station 8, and an increasingly larger bubble at stations

9–11. The mean profiles of the 0.4-mm and 0.8-mm bar cases are

Fig. 3 Cp profiles, low FSTI, ReÄ25,000 cases

Fig. 4 Station 7–11 profiles, low FSTI, ReÄ25,000 cases: „a...

mean velocity, „b... u8, „c... intermittency

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NASA/CR—2012-217656 29
virtually indistinguishable from the baseline case. The 1.6-mm bar

case shows a clear separation bubble at station 7, suggesting that

the boundary layer has separated from the trailing edge of the bar.

This bubble continues to grow, and at station 8 the separation

bubble is larger with the 1.6-mm bar than in the other cases. At

station 9 all cases appear similar. This agreement is really a crossing, as the shear layer in the 1.6-mm bar case is on the verge of

reattaching, while the separation bubble is growing in the other

cases. At station 10, the near-wall velocity in the 1.6-mm bar case

has begun to rise, indicating incipient reattachment. By station 11,

the boundary layer has clearly reattached in the 1.6-mm bar case,

although the mean profile has not recovered to a fully developed

turbulent shape. Dimensionless u8 profiles are shown in the second row of Fig. 4. As with the mean profiles, the baseline case and

the 0.4-mm and 0.8-mm bar cases are indistinguishable through

station 10. All show a small u8 peak growing in the shear layer

over the separation bubble. As explained by Volino @7,18#, this

peak is not indicative of transition. It is caused primarily by lowfrequency fluctuations that are amplified when they act across the

region of high mean velocity gradient in the shear layer. At station

11, there is a slight increase in u8 in the 0.8-mm bar case over the

baseline case, and the high values extend into the near-wall region. These near-wall fluctuations suggest the beginning of transition and reattachment, but they are not large enough in this case

to significantly affect the mean profile. The 1.6-mm bar case

shows a larger u8 peak than the other cases at stations 7 and 8.

The peak is in the shear layer and is similar to the peaks at stations

9 and 10 of the other cases. It does not indicate transition. The

peak becomes larger at station 9, and extends into the near-wall

region, which is a sign of incipient transition. By station 10 u8 is

much larger with a clear double peak. This rise in u8 corresponds

to the beginning of reattachment observed in the mean profile. The

third row of Fig. 4 shows the local intermittency. Its is zero in all

cases through station 9, but begins to rise at stations 10 and 11 of

the 1.6-mm bar case. The intermittency peak is in the shear layer,

indicating that this is where transition begins. Intermittency only

reaches about 13% at station 11. This is consistent with the mean

velocity profile, which shows the boundary layer is reattached but

not yet a fully developed turbulent profile. As the turbulence is

intermittent, it is likely that the boundary layer is only intermittently reattached. The high u8 peak at station 11 is also consistent

with a transitional boundary layer. As a boundary layer becomes

fully turbulent, the dimensionless u8 peak will decrease in magnitude to about 0.1, and move close to the wall.

The behavior in the 1.6-mm bar case is interesting. The bar was

not large enough to immediately trip the boundary layer to turbulent, but it did move the separation point upstream. This caused

the transition to move upstream, and led to at least a partial reattachment by the trailing edge, which did not occur in the other

cases. The 0.8-mm bar trip case is also very interesting. The bar in

this case was so small that it had no immediate measurable affect

on the mean or u8 profiles. Well downstream at station 11, however, the effect of this bar became visible in the u8 profile. Apparently this bar introduced a very small disturbance in the flow,

which was too small to detect at first, but grew as it moved downstream.

The velocity profiles of Fig. 4 and the pressure profiles of Fig.

3 are in good agreement. Both show transition and reattachment at

the same locations, and the measured static pressures agree with

the local free-stream velocities of Fig. 4. The agreement between

the pressure and velocity results was apparent in all cases. For

brevity, the pressure profiles are not presented in the cases which

follow.

Figure 5 shows the velocity profiles for the high-FSTI, Re

525,000 cases. The format is that same as in Fig. 4. As in the

low-FSTI case, the baseline, 0.4-mm trip and 0.8-mm trip cases

are nearly indistinguishable at the upstream stations. By station

11, some differences are apparent in the mean profiles for these

cases. The separation bubble is less distinct in the cases with the

bars, but the boundary layer still does not appear fully reattached.

Intermittency rises slightly above zero at station 11 of the 0.8-mm

bar case, while remaining essentially at zero with the smaller bar

and in the baseline case. The u8 profiles show a large peak in the

shear layer, which grows in the streamwise direction. As shown in

Volino @7#, this peak is caused by the action of the high FSTI on

the shear layer, and does not indicate significant momentum transport. As in the low-FSTI cases of Fig. 4, the 1.6-mm bar case

shows significant differences from the other cases in Fig. 5. The

u8 peak is significantly higher in this case at stations 7 and 8. At

station 9, u8 values are higher in the near-wall region, the intermittency rises above zero, and the mean profile appears to be

reattached. At stations 10 and 11 the intermittency continues to

increase. The mean profile adjusts toward a more turbulent shape

between stations 9 and 11. The u8 peak decreases somewhat by

station 11, but still shows the relatively high values of a transitional boundary layer, rather than the somewhat lower values of a

fully turbulent boundary layer. As in the low-FSTI, Re525,000

case, the 1.6-mm bar is not large enough to immediately trip the

boundary layer to turbulent, but it causes transition to move upstream and leads to a reattachment that did not occur in the baseline or smaller bar cases.

The velocity profiles of the low-FSTI, Re550,000 case are

shown in Fig. 6. The effects of the bars are clear. At station 7, the

1.6-mm bar has caused a relatively large separation bubble compared to the other cases and a small u8 peak in the shear layer

over this bubble. The smaller bar cases are indistinguishable from

the baseline case, with mean profiles only on the verge of separation and u8 near zero. By station 8, the 1.6-mm bar case has

Fig. 5 Station 7—profiles, high FSTI, ReÄ25,000 cases: „a...

mean velocity, „b... u8, „c... intermittency

Fig. 6 Station 7–11 profiles, low FSTI, ReÄ50,000 cases: „a...

mean velocity, „b... u8, „c... intermittency

758 Õ Vol. 125, OCTOBER 2003 Transactions of the ASME

NASA/CR—2012-217656 30
undergone a sudden transition, with g51, high u8 levels in both

the shear layer and near-wall regions, and the beginning of reattachment as indicated by nonzero mean velocity near the wall.

The other cases are indistinguishable, exhibiting a small separation bubble and showing no sign of transition or reattachment. At

station 9, the boundary layer is clearly reattached in the 1.6-mm

bar case, and u8 values are beginning to rise in the 0.8-mm bar

case. At station 10 the intermittency indicates fully turbulent flow

and the boundary layer has reattached in the 0.8-mm bar case. The

0.4-mm bar case is still separated with g50, but u8 has begun to

rise near the wall. By station 11 the 0.4-mm bar case has become

turbulent and the boundary layer has started to reattach. The baseline case remains nonturbulent with a large separation bubble at

station 11. As observed in the Re525,000 cases, the 1.6-mm bar is

not large enough to immediately trip the boundary layer to turbulent, but it does move separation upstream, which causes transition and reattachment to move significantly upstream. The smaller

bars appear to have no immediate effect on the boundary layer,

but they must introduce small disturbances that grow in the

streamwise direction and have a significant effect in moving the

transition and reattachment upstream. The 0.8-mm bar must introduce a larger disturbance than the 0.4-mm bar, since transition and

reattachment occur one station farther upstream with the 0.8-mm

bar.

Figure 7 shows the high-FSTI, Re550,000 cases. With the

larger bars, transition and reattachment move upstream. With the

1.6-mm bar, the intermittency is already nonzero by station 7, and

the boundary layer is fully turbulent and attached by station 8.

With the 0.8-mm bar, g rises above zero at station 8 and is near

fully turbulent by station 10. The mean profile appears to indicate

reattachment by station 9. The intermittency rises above zero in

the 0.4-mm bar case at station 9, and continues to rise at stations

10 and 11. The mean profile shows reattachment at station 10. In

the baseline case, the intermittency begins to rise at station 10,

and the boundary layer is reattached at station 11. At station 11 the

mean profiles are indistinguishable in the cases with bars, and

fuller than in the baseline case. In all cases, the transition begins

upstream of the location in the corresponding low-FSTI case of

Fig. 6, but the transition length is longer. Volino and Hultgren @22#

also observed that transition begins farther upstream with highFSTI, but is more abrupt in low-FSTI cases.

Figure 8 shows the low-FSTI, Re5100,000 cases. The 1.6-mm

bar immediately trips the boundary layer to turbulent and eliminates the separation bubble. The 0.8-mm bar causes a small u8

peak above the baseline values at station 7. The intermittency

jumps from 0 to 1 between stations 7 and 8, and the separation

bubble is effectively eliminated. As in the lower-Re cases, the

0.4-mm bar has no visible effect at station 7, and the mean and u8

profiles are indistinguishable from the baseline case. The boundary layer separates, but by station 9 the shear layer has become

fully turbulent and begun to reattach. In the baseline case, u8 does

not begin to show elevated near wall values until station 10, and

transition and reattachment occur at station 11. The mean profiles

at station 11 show the fullest profile and thinnest boundary layer

in the 0.4-mm bar case. The larger bars result in thicker boundary

layers. The mean profile in the baseline case has not yet recovered

to a fully turbulent shape. As will be discussed below, the thinner

attached boundary layer in the 0.4-mm bar case suggests that this

case will have lower losses than the other cases.

The high-FSTI, Re5100,000 cases are shown in Fig. 9. As in

the low-FSTI case, the 1.6-mm bar trips the boundary layer to

turbulent and eliminates separation at this Re. The intermittency is

nonzero at station 7 of the 0.8-mm bar case, and it continues to

rise through station 10, indicating an extended transition zone.

Because the transition begins so far upstream, the separation

bubble is eliminated and transition occurs in an attached boundary

layer. With the 0.4-mm bar, the intermittency indicates that the

transition does not begin until station 8, so a small separation

bubble forms, as in the baseline case. The boundary layer is reattached by station 9, however, and the transition is nearly complete

by station 10. In the baseline case, the transition begins at station

9, and the boundary layer is reattached at station 10. Examining

the mean profiles, the 1.6-mm bar causes an immediate thickening

of the boundary layer, and the separation bubble in the baseline

case also causes a thicker boundary layer. By station 11, the mean

profiles for these two cases agree closely. The boundary layers are

thinner in the cases with the smaller bars.

Fig. 7 Station 7–11 profiles, high FSTI, ReÄ50,000 cases: „a...

mean velocity, „b... u8, „c... intermittency

Fig. 8 Station 7–11 profiles, low FSTI, ReÄ100,000 cases: „a...

mean velocity, „b... u8, „c... intermittency

Fig. 9 Station 7–11 profiles, high FSTI, ReÄ100,000 cases: „a...

mean velocity, „b... u8, „c... intermittency

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NASA/CR—2012-217656 31
In the low-FSTI, Re5200,000 cases of Fig. 10, the intermittency profiles show that the 0.8-mm and 1.6-mm bars immediately

trip the boundary layer to turbulent and eliminate the separation

bubble. The 1.6-mm bar, which is the same thickness as the

boundary layer at station 7 of the baseline case, results in a substantially thicker boundary layer than in all of the other cases. The

0.4-mm bar case shows a small separation bubble at station 8, but

is fully turbulent and reattached by station 9. The baseline case

exhibits a clear separation bubble at station 9, and is fully turbulent and reattached by station 10. The mean profiles at station 11

show that the growth of the bubble in the baseline case results in

a thicker boundary layer than in the 0.8- and 0.4-mm bar cases.

Figure 11 shows the high-FSTI, Re5200,000 cases. As in the

low-FSTI cases of Fig. 10, the 0.8- and 1.6-mm bars trip the

boundary layer to turbulent. The transition has already started, as

indicated by the nonzero intermittency, in the 0.4-mm bar case at

station 7. In all of these cases, there is no separation. In the baseline case, g does not rise above zero until station 9, and there may

be a small separation bubble at station 8. At station 11, the mean,

u8, and intermittency profiles of the baseline, 0.4-mm, and

0.8-mm bar cases are all in good agreement, while the 1.6-mm bar

case exhibits a noticeably thicker boundary layer.

The low FSTI, Re5300,000 cases are shown in Fig. 12. As in

the Re5200,000 cases, the 0.8-mm and 1.6-mm bars trip the

boundary layer to turbulent. The 0.4-mm bar appears to have no

effect at station 7, where the mean and u8 profiles agree with the

baseline case and the intermittency is zero. By station 8, however,

the boundary layer in the 0.4-mm bar case has become fully turbulent, while in the baseline case it is still laminar and has separated. By station 9 the shear layer in the baseline case is transitional, and it is fully turbulent and reattached by Station 10. At

Station 11 the mean profiles for the four cases are all different,

with the 0.4-mm bar and baseline cases having the thinnest

boundary layers.

Figure 13 shows the high-FSTI, Re5300,000 cases. As in the

low-FSTI cases, the 0.8-mm and 1.6-mm bars trip the boundary

layer to fully turbulent, and the 0.4-mm bar causes the transition

to start by station 7 and finish by station 8. Transition has started

at station 8 of the baseline case and is complete near station 10. In

all of these cases the transition begins far enough upstream to

prevent separation. The mean profiles at all stations show that the

boundary layer is thinnest in the baseline case and that the thickness increases with the bar size. With the 0.4-mm and 0.8-mm

bars the boundary layer is only slightly thicker than in the baseline

case, but it is substantially thicker in the 1.6-mm bar case.

Some consistent trends run through the data from all cases. If a

bar is large enough, it will immediately trip the boundary layer to

fully turbulent and prevent separation. As Reynolds number increases, the boundary layer thickness decreases as does the thickness of the bar required for tripping. If a bar is small enough, it

initially appears to have no effect on the boundary layer. The

boundary layer appears to proceed over the bar with no measurable change in the mean velocity or u8 from the corresponding

baseline case. The bars must, however, introduce some small disturbance into the boundary layer. The boundary layer is unstable

against small disturbances in the adverse pressure gradient region,

so the small disturbances grow and eventually cause transition.

Larger bars must impart larger ~albeit sometimes still undetectFig. 10 Station 7–11 profiles, low FSTI, ReÄ200,000 cases: „a...

mean velocity, „b... u8, „c... intermittency

Fig. 11 Station 7–11 profiles, high FSTI, ReÄ200,000 cases:

„a... mean velocity, „b... u8, „c... intermittency

Fig. 12 Station 7–11 profiles, low FSTI, ReÄ300,000 cases: „a...

mean velocity, „b... u8, „c... intermittency

Fig. 13 Station 7–11 profiles, high FSTI, ReÄ300,000 cases:

„a... mean velocity, „b... u8, „c... intermittency

760 Õ Vol. 125, OCTOBER 2003 Transactions of the ASME

NASA/CR—2012-217656 32
able! perturbations than the smaller bars, resulting in transition

locations that move upstream as bar size is increased. The optimal

bar size depends on the Reynolds number and is discussed further

below.

The present results shed some light on the transition mechanism

under both high- and low-FSTI conditions. Volino @18# examined

spectra of the fluctuating velocity in the boundary layers and shear

layers of the baseline cases. He observed sharp peaks in the spectra of the low-FSTI cases at frequencies that matched the most

unstable frequencies for Tollmien-Schlicthing ~TS! waves in the

boundary layer just upstream of separation. He therefore concluded that the transition in the shear layer might be through a TS

mechanism in these cases. In the high-FSTI cases, Volino @18#

observed broadband peaks in the spectra and the relatively long

transition regions noted above. Volino and Hultgren @22# made

similar observations, and concluded that the high-FSTI separated

flow transition was through a bypass mode, very similar to the

high-FSTI transition in an attached boundary layer. Volino @18#,

however, noted that the broadband peaks in the high-FSTI case

spectra were centered at the same frequencies as in the low-FSTI

cases, suggesting a similar transition mechanism under high- and

low-FSTI conditions. He concluded that disturbances that began

to grow in the boundary layer prior to separation were causing a

TS-type transition in the shear layer over the separation bubble in

both the high- and low-FSTI cases. High-FSTI has a strong effect

in moving the transition upstream, but the bars in the present cases

had an equally strong or stronger effect in both the high- and

low-FSTI cases. This confirms that the free-stream turbulence is

not solely responsible for bypass transition in the high-FSTI cases.

The magnitude of the disturbances induced by the bars is too

small to be quantified based on the mean or u8 results presented

above. Perhaps more can be learned from boundary layer spectra.

Analyses of spectra based on u8 fluctuations for the present cases

show some interesting but inconclusive results. Volino @18# found

that u8 spectra are often characterized by low-frequency fluctuations that are induced by the free stream and have no direct effect

on transition. In the early stages of transition, these low-frequency

fluctuations can hide the very-low-amplitude fluctuations important for transition. Volino @18# found that spectra of the turbulent

shear stress are less affected by the low-frequency unsteadiness

and can provide a better means for detecting the early stages of

transition. Acquisition and analysis of turbulent shear stress data

for the cases of the present study may prove useful for explaining

and quantifying the transition mechanism.

Shape Factor and Momentum Thickness. The shape factor

and momentum thickness are useful parameters for evaluating the

state of the boundary layer with respect to separation, transition,

and losses. They provide a means for summarizing the information presented in the velocity profiles of Figs. 4 –13. In the present

cases, the boundary layer has a shape factor H of about 2.4 at the

end of the favorable pressure gradient region. This is the expected

value for this laminar, accelerated boundary layer. If the boundary

layer separates, the displacement thickness grows rapidly, while

the momentum thickness remains nearly constant. The result is a

very high shape factor. If the boundary layer reattaches, the displacement thickness drops, and the momentum thickness begins to

grow. The boundary layer eventually recovers to a fully turbulent

shape, with a shape factor of about 1.6 in the present cases.

Stage losses in a multiblade turbine cascade can be determined

through measurement of the momentum deficit in the wake downstream of the blade row. With the single-passage facility of the

present study, wake measurements are not meaningful, since there

is flow on only one side of the airfoils on each side of the passage.

If a boundary layer separates and does not fully reattach, or reattaches near the trailing edge, momentum thickness will be relatively low at the trailing edge, and high losses will be generated in

the wake downstream of the passage. While the losses in such a

case cannot be quantified in the present study, it is safe to assume

that they would be unacceptably high, and that there would be an

unacceptable loss of lift from the airfoils. For those cases in which

the boundary layer reattaches and recovers to a fully developed

turbulent shape, the suction side profile loss is likely the dominant

loss mechanism in the passage ~Howell et al. @9#!. As explained

by Howell et al. @9#, for a given shape factor and passage exit

angle, the momentum thickness of the suction side boundary layer

at the trailing edge is proportional to the suction side profile loss.

Figure 14 provides an example of the development of the shape

factor and momentum thickness, using the low-FSTI, Re

5100,000 cases. In the baseline case, H increases from 2.4 to a

high value of about 5 as the boundary layer separates. Reattachment occurs near the trailing edge, and H drops to about 2.2,

which is still above the turbulent value of 1.6, indicating that

recovery from the separation is not complete. In the 0.4-mm bar

case, the boundary layer separates and H reaches a value of 3.7.

The boundary layer then reattaches, and H gradually drops to a

fully turbulent value by the trailing edge. With the 0.8-mm bar,

transition occurs far enough upstream to prevent separation, and H

drops continuously from a laminar value to a turbulent value as

transition occurs. In the 1.6-mm bar case the boundary layer is

tripped to turbulence, and H quickly reaches its turbulent value.

The 1.6-mm bar causes an immediate thickening of the boundary

layer, and the momentum thickness remains higher than in the

other cases at all streamwise locations. The 0.4- and 0.8-mm bars

appear to have no immediate effect on u. When transition and

reattachment occur, however, u begins to rise. When reattachment

occurs in the baseline case, it causes u to increase to a higher

value than in the 0.4- and 0.8-mm bar cases. Near the trailing edge

u is lowest in the 0.4-mm bar case. This would presumably be the

case with the lowest profile losses. The larger bars force the transition to occur farther upstream than necessary, resulting in a

longer turbulent region and higher losses. In the baseline case the

separation bubble becomes relatively thick, resulting in a thick

boundary layer after reattachment. The 0.4-mm bar case provides

a good example of the controlled diffusion described by Hourmouziadis @1#. With the 0.4-mm bar, the separation bubble is relatively thin, and the turbulent region is relatively short, resulting in

lower losses.

Figure 15 shows the shape factor and momentum thickness at

station 11 (s/Ls50.94), near the trailing edge, for all the lowFSTI cases. In the Re525,000 cases, the shape factor indicates

that the boundary layer only reattaches in the 1.6-mm bar case.

The shape factor in this case is still above the expected turbulent

value, indicating that recovery from the separation in not complete. Comparison of momentum thicknesses is not meaningful at

Fig. 14 Shape factor and momentum thickness versus

streamwise location, low FSTI, ReÄ100,000: „a... H, „b... u

Journal of Turbomachinery OCTOBER 2003, Vol. 125 Õ 761

NASA/CR—2012-217656 33
this Re. For the Re550,000 cases, the 0.8- and 1.6-mm bars cause

reattachment, while recovery from the separation is only partially

complete in the 0.4-mm bar case and the shear layer remains

separated in the baseline case. The 0.8- and 1.6-mm bar cases are,

therefore, preferable at this Re, and both have about the same

momentum thickness at station 11. At Re5100,000, already described in Fig. 14, the 0.4-mm bar produces the lowest losses. The

0.4-mm bar case is also best at Re5200,000, with slightly lower

losses than with the 0.8-mm bar or in the baseline case. At Re

5300,000, the transition occurs sufficiently far upstream in the

baseline case to keep the separation bubble small and produce

lower losses than in any of the cases with bars.

The station 11 shape factors and momentum thicknesses for the

high-FSTI cases are shown in Fig. 16. As in the low-FSTI cases,

only the 1.6-mm bar is large enough to force reattachment at

Re525,000, and it does not even quite result in full recovery to a

turbulent profile. At Re550,000, the shape factor shows that all of

the bars cause reattachment, while the boundary layer in the baseline case has reattached but not fully recovered from the separation. The 0.4- and 0.8-mm bar cases have lower momentum thickness than the 1.6-mm bar case. At the higher Re, reattachment is

complete in all cases, and the 1.6-mm bar cases have significantly

higher losses than the other cases. At Re5100,000, the 0.4- and

0.8-mm bar cases and the baseline case all have about the same

losses. As Re increases to 200,000 and 300,000, the baseline case

emerges as the case with lowest losses, in agreement with the

low-FSTI cases of Fig. 15.

The optimal bar height clearly varies with the Reynolds number. As Re increases, the boundary layer becomes thinner and

more prone to transition, so a smaller bar is needed. At Re

525,000, the 1.6-mm bar is needed, and a larger bar would be

desirable to force a more complete reattachment. At Re550,000,

the 0.8-mm bar is best, since it is large enough to cause reattachment at low FSTI, but produces lower losses than the thicker bar

at high FSTI. At Re5100,000 and 200,000, the 0.4-mm bar is

best, since it is large enough to force complete reattachment at

low FSTI, and results in equal or slightly lower losses than the

baseline or 0.8-mm bar cases. At Re5300,000, the baseline case is

best, although the losses are only slightly lower than those of the

0.4-mm bar case. If a bar is used for passive flow control, a single

bar thickness must be chosen for optimal overall performance.

The best size will depend on the operating range of the engine. If

the operating range is large, a compromise between improved performance at cruise and higher losses at takeoff may be needed.

Correlation of Results. The size of a bar necessary to trip a

boundary layer to turbulence can be predicted using the following

correlation from Gibbings @23#.

Red5Udd/n.600, (1)

where d is the bar thickness and Ud is the velocity in the untripped

boundary layer at y5d at the streamwise location of the bar.

Equation ~1! predicts that bar thicknesses of 4.7 mm, 2.3 mm, 1.3

mm, 0.68 mm, and 0.50 mm would be needed to immediately trip

the boundary layer to turbulent in the Re525,000 through

300,000 cases, respectively. In agreement with this prediction, the

results above show that the boundary layer was only tripped in the

Re5200,000 and 300,000 cases with the 0.8-mm and 1.6-mm

bars, and in the Re5100,000 cases with the 1.6-mm bar. Since an

optimal bar does not immediately trip the boundary layer; it will

be thinner than indicated by Eq. ~1!.

The most effective bars in the present cases appear to be those

that cause reattachment to begin between stations 8 and 9, at s/Ls

of about 0.74. When reattachment begins by this location, there is

sufficient distance downstream for the reattachment and recovery

from the separation to be completed before the trailing edge. The

beginning of reattachment and the start of transition are related

and occur at approximately the same location. There are a few

correlations in the literature for prediction of the distance from

separation to transition onset. In general they are not very robust,

but some give reasonable estimates. Mayle @2# provides the following correlations:

Rest5300 Reus

0.7 ~short bubble!, (2)

Rest51000 Reus

0.7 ~long bubble!. (3)

Equations ~2! and ~3! apply to short and long separation bubbles,

respectively. Volino @7# found that the present baseline case results

lie between the predictions of Eqs. ~2! and ~3! tending toward the

long bubble correlation at low FSTI and about midway between

the two correlations at high FSTI. Although they differ by a factor

of 3, Eqs. ~2! and ~3! provide at least a rough estimate of the

reattachment location.

The following correlation provides an estimate of the effect of

bar height on reattachment location. The equation is based on a

curve fit of the present data:

~st2s p!m /~st2s p!b5@110.23~d/u p!

1.56#21, (4)

where (st2s p)b is the distance from the suction side velocity

maximum (s/Ls50.53) to the location of the beginning of reattachment in the baseline case, and (st2s p)m is this distance with

a bar in place. The present data along with Eq. ~4! are shown in

Fig. 15 Station 11 shape factor and momentum thickness versus Re, low-FSTI cases: „a... H, „b... u

Fig. 16 Station 11 shape factor and momentum thickness versus Re, high-FSTI cases: „a... H, „b... u

762 Õ Vol. 125, OCTOBER 2003 Transactions of the ASME

NASA/CR—2012-217656 34
Fig. 17. The finite spacing of the streamwise measurement stations

results in some uncertainty in the transition start location, as indicated by the error bars in Fig. 17. Volino @7# showed that the

boundary layer behavior upstream of the separation is predictable

and laminar. To predict reattachment in a case with a bar, the

laminar solution could be used to predict the separation location

and the momentum thickness before separation. A correlation such

as Eqs. ~2! or ~3! could then be used to estimate the distance to

transition and reattachment in the baseline flow. This would give

stb . Equation ~4! could then be used to predict stm for a given bar

thickness. Alternatively, the desired stm could be specified and

used with Eq. ~4! to predict the optimal bar thickness. Setting

stm5169 mm (stm /Ls50.74) and using the measured ~when

available! or predicted values for stb and u p for the baseline cases,

optimal bar heights have been predicted for the present cases and

are presented in Table 3.

Discussion. The above correlations are based only on the

present data set, so it is doubtful that they are universally applicable. Still, they provide a start in the assessment of passive flow

control devices. With more experiments with different airfoils, it

may be possible to refine correlations such as Eq. ~4!, to make

them more generally applicable.

Alternative passive devices such as vortex generators or

dimples should also be considered. There is no guarantee, however, that these devices will provide improvement over the present

bars. Dimples have been used in some applications such as internal blade cooling, to provide enhanced heat transfer with lower

pressure drop than boundary layer trips. Dimples enhance heat

transfer by promoting turbulence and mixing, which would also

tend to promote boundary layer attachment. In the present application, however, we do not seek to enhance turbulence or heat

transfer, or even to fully eliminate the separation bubble. Since the

smaller bars in the present experiments did not increase losses or

boundary layer thickness, or have any other immediate measurable effect on the boundary layer, it is not clear that any other

device will be superior.

While the present bars have proven effective, it is clear from

Table 3 and Figs. 15 and 16 that the optimal bar height varies with

Re and FSTI. If applied passively, a single bar height would be

selected for the entire operating range of the engine. If the operating range is large, a compromise will be necessary between

improved performance at low Re and higher losses at high Re.

Passive flow control has the distinct advantage of being relatively

simple to implement in practice, but active flow control may provide a means for optimizing performance over a wider range of

conditions. Unsteady active control also provides possibilities for

further flexibility and improved performance not available with

passive devices. Further consideration of active control is presented by Volino @24#.

The present cases all involve steady inlet flow. In engine flows,

the periodic wakes from upstream airfoils will make the flow unsteady. While steady flow experiments are necessary for building

understanding of the flow and flow control devices, experiments

should eventually be performed in flows with wakes.

Conclusions

1. Rectangular bars have been successfully employed as flow

control devices on the suction side of a low-pressure turbine airfoil. Boundary layer reattachment was forced even in very-lowReynolds-number cases.

2. Optimal bars are not large enough to immediately trip the

boundary layer to turbulent or prevent separation, but rather induce very small disturbances that at first are essentially undetectable, but eventually promote transition in the shear layer at a

downstream location.

3. Bars were effective under both high- and low-FSTI conditions, indicating that the high-FSTI transition is not simply a bypass transition induced by the free stream.

4. The optimal location for reattachment results in a relatively

short turbulent region, but occurs sufficiently far upstream to prevent a large separation bubble and ensure complete recovery from

the separation before the trailing edge. A bar height can be selected to induce reattachment at the desired location.

5. The optimal bar height varies with the Reynolds number and

free-stream turbulence level. Bars that were large enough to induce reattachment at the lowest Re produced significantly higher

losses at the higher Re. If a wide range of Reynolds numbers are

encountered in practice, some compromise between improved performance at low Re and higher losses at high Re will be necessary

in the choice of an overall best bar height.

Fig. 17 Correlation of transition and reattachment start location to bar height; bars indicate range of possible values resulting from finite station spacing

Table 3 Baseline case boundary layer thickness at bar location „sÕLsÄ0.53..., and predicted

bar heights for tripping „dtrip... and for incipient reattachment „dopt... at sÕLsÄ0.74 „all values in

mm...

Re31023

Low FSTI High FSTI

d 99.5p u p dtrip dopt d 99.5p u p dtrip dopt

25 3.8 0.48 4.7 2.4 3.9 0.48 4.7 1.5

50 2.7 0.34 2.3 1.2 2.8 0.33 2.3 0.6

100 1.9 0.23 1.3 0.4 2.2 0.24 1.3 0.2

200 1.3 0.16 0.7 0.2 1.6 0.16 0.7 0

300 1.2 0.14 0.5 0.1 1.4 0.14 0.5 0

Journal of Turbomachinery OCTOBER 2003, Vol. 125 Õ 763

NASA/CR—2012-217656 35
Acknowledgments

This work was sponsored by the NASA Glenn Research Center.

The grant monitor is Dr. David Ashpis. Additional matching support was provided through a U.S. Naval Academy Recognition

Grant.

Nomenclature

Cp 5 2(PT2P)/rUe

2

, pressure coefficient

d 5 Bar height

FSTI 5 Free-stream turbulence intensity

H 5 d */u, shape factor

K 5 (n/U`

2 )(dU` /ds), acceleration parameter

Ls 5 Suction surface length

P 5 Pressure

PT 5 Upstream stagnation pressure

Re 5 UeLs /n, exit Reynolds number

Red 5 Udd/n, Reynolds number based on bar height

Rest 5 U`(st2ss)/n, separation to transition distance Reynolds number

Reu 5 Momentum thickness Reynolds number

s 5 Streamwise coordinate, distance from leading edge

U 5 Mean streamwise velocity

U` 5 Local free-stream velocity

Ud 5 Mean velocity at bar height in baseline boundary

layer

Ue 5 Nominal exit free-stream velocity, based on the inviscid solution

u8 5 rms streamwise fluctuating velocity

y 5 Cross-stream coordinate, distance from wall

d 99.5 5 99.5% boundary layer thickness

d* 5 Displacement thickness

g 5 Intermittency, fraction of time flow is turbulent

n 5 Kinematic viscosity

r 5 Density

u 5 Momentum thickness

Subscripts

b 5 Baseline flow

m 5 Modified flow

p 5 Suction surface pressure minimum, velocity maximum

s 5 Separation location

t 5 Transition start location

References

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Center, Cleveland, OH, pp. 53–64.

@6# Curtis, E. M., Hodson, H. P., Banieghbal, M. R., Denton, J. D., Howell, R. J.,

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@7# Volino, R. J., 2002, ‘‘Separated Flow Transition Under Simulated LowPressure Turbine Airfoil Conditions: Part 1—Mean Flow and Turbulence Statistics,’’ ASME J. Turbomach., 124, pp. 645–655.

@8# Van Treuren, K. W., Simon, T., von Koller, M., Byerley, A. R., Baughn, J. W.,

and Rivir, R., 2002, ‘‘Measurements in a Turbine Cascade Flow Under Ultra

Low Reynolds Number Conditions,’’ ASME J. Turbomach., 124, pp. 100–106.

@9# Howell, R. J., Ramesh, O. N., Hodson, H. P., Harvey, N. W., and Schulte, V.,

2001, ‘‘High Lift and Aft-Loaded Profiles for Low-Pressure Turbines,’’ ASME

J. Turbomach., 123, pp. 181–188.

@10# Stadtmu ̈ller, P., Fottner, L., and Fiala, A., 2000, ‘‘Experimental and Numerical

Investigation of Wake-Induced Transition on a Highly Loaded LP Turbine at

Low Reynolds Numbers,’’ ASME paper no. 2000-GT-0269.

@11# Brunner, S., Fottner, L., and Schiffer, H.-P., 2000, ‘‘Comparison of Two Highly

Loaded Low Pressure Turbine Cascades Under the Influence of Wake-Induced

Transition,’’ ASME paper no. 2000-GT-268.

@12# Lake, J. P., King, P. I., and Rivir, R. B., 2000, ‘‘Low Reynolds Number Loss

Reduction on Turbine Blades With Dimples and V-Grooves,’’ AIAA paper no.

00-738.

@13# Murawski, C. G., and Vafai, K., 1999, ‘‘Effect of Variable Axial Chord on a

Low-Pressure Turbine Blade,’’ J. Propul. Power, 15, pp. 667–674.

@14# Byerley, A. R., Sto ̈rmer, O., Baughn, J. W., Simon, T. W., Van Treuren, K. W.,

and List, J., 2002, ‘‘Using Gurney Flaps to Control Laminar Separation on

Linear Cascade Blades,’’ ASME J. Turbomach., 125, pp. 114–120.

@15# Bons, J. P., Sondergaard, R., and Rivir, R. B., 2001, ‘‘Turbine Separation

Control Using Pulsed Vortex Generator Jets,’’ ASME J. Turbomach., 123, pp.

198–206.

@16# Bons, J. P., Sondergaard, R., and Rivir, R. B., 2002, ‘‘The Fluid Dynamics of

LPT Blade Separation Control Using Pulsed Jets,’’ ASME J. Turbomach., 124,

pp. 77–85.

@17# Volino, R. J., Schultz, M. P., and Pratt, C. M., 2001, ‘‘Conditional Sampling in

a Transitional Boundary Layer Under High Free-Stream Turbulence Conditions,’’ ASME J. Fluids Eng., 125, pp. 28–37.

@18# Volino, R. J., 2002, ‘‘Separated Flow Transition Under Simulated LowPressure Turbine Airfoil Conditions: Part 2—Turbulence Spectra,’’ ASME J.

Turbomach., 124, pp. 656–664.

@19# Wills, J. A. B., 1962, ‘‘The Correction of Hot-Wire Readings for Proximity to

a Solid Boundary,’’ J. Fluid Mech., 12, pp. 65–92.

@20# Ligrani, P. M., and Bradshaw, P., 1987, ‘‘Spatial Resolution and Measurement

of Turbulence in the Viscous Sublayer Using Subminiature Hot-Wire Probes,’’

Exp. Fluids, 5, pp. 407–417.

@21# Ligrani, P. M., and Bradshaw, P., 1987, ‘‘Subminiature Hot-Wire Sensors:

Development and Use,’’ J. Phys. E, 20, pp. 323–332.

@22# Volino, R. J., and Hultgren, L. S., 2001, ‘‘Measurements in Separated and

Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions,’’ ASME J. Turbomach., 123, pp. 189–197.

@23# Gibbings, J. C., 1959, ‘‘On Boundary-Layer Transition Wires,’’ Aeronautical

Research Council, Current Papers 462.

@24# Volino, R. J., 2003, ‘‘Separation Control on Low-Pressure Turbine Airfoils

Using Synthetic Vortex Generator Jets,’’ ASME J. Turbomach., 125, pp. 765–

777.

764 Õ Vol. 125, OCTOBER 2003 Transactions of the ASME

NASA/CR—2012-217656 36
Ralph J. Volino

Department of Mechanical Engineering,

United States Naval Academy,

Annapolis, MD 21402

e-mail: volino@usna.edu

Separation Control on

Low-Pressure Turbine Airfoils

Using Synthetic Vortex Generator

Jets

Oscillating vortex generator jets have been used to control boundary layer separation

from the suction side of a low-pressure turbine airfoil. A low Reynolds number (Re

525,000) case with low free-stream turbulence has been investigated with detailed measurements including profiles of mean and fluctuating velocity and turbulent shear stress.

Ensemble averaged profiles are computed for times within the jet pulsing cycle, and

integral parameters and local skin friction coefficients are computed from these profiles.

The jets are injected into the mainflow at a compound angle through a spanwise row of

holes in the suction surface. Preliminary tests showed that the jets were effective over a

wide range of frequencies and amplitudes. Detailed tests were conducted with a maximum

blowing ratio of 4.7 and a dimensionless oscillation frequency of 0.65. The outward pulse

from the jets in each oscillation cycle causes a disturbance to move down the airfoil

surface. The leading and trailing edge celerities for the disturbance match those expected

for a turbulent spot. The disturbance is followed by a calmed region. Following the

calmed region, the boundary layer does separate, but the separation bubble remains very

thin. Results are compared to an uncontrolled baseline case in which the boundary layer

separated and did not reattach, and a case controlled passively with a rectangular bar on

the suction surface. The comparison indicates that losses will be substantially lower with

the jets than in the baseline or passively controlled cases. @DOI: 10.1115/1.1626686#

Introduction

Modern low-pressure turbine ~LPT! airfoils are subject to increasingly stronger pressure gradients as designers impose higher

loading in an effort to improve efficiency and lower cost by reducing the number of airfoils in an engine. If the adverse pressure

gradient on the suction side of these airfoils becomes strong

enough, the boundary layer will separate. Separation bubbles, particularly those which fail to reattach, can result in a significant

loss of lift and a subsequent degradation of engine efficiency ~e.g.,

Hourmouziadis @1#, Mayle @2#, and Sharma et al. @3#!. The problem is particularly relevant in aircraft engines. Airfoils optimized

to produce maximum power under takeoff conditions may still

experience boundary layer separation at cruise conditions, due to

the thinner air and lower Reynolds numbers at altitude. A component efficiency drop of 2% may occur between takeoff and cruise

conditions in large commercial transport engines, and the difference could be as large as 7% in smaller engines operating at

higher altitudes.

Separation on LPT airfoils is complicated by boundary layer

transition, which can prevent separation if it occurs far enough

upstream, or induce boundary layer reattachment if it occurs in the

shear layer over a separation bubble. At lower Reynolds numbers

transition will tend to occur farther downstream, hence the problems associated with performance at altitude.

Separated flow transition has been studied extensively, and in

recent years several studies have focused on transition in the LPT.

Volino @4# provides a review of much of that work. Separation can

be affected through naturally occurring phenomena such as high

free-stream turbulence intensity ~FSTI! or the unsteadiness caused

by wakes generated upstream of an airfoil. Further discussion of

these effects is available in Volino @5#. While high FSTI and

wakes help to mitigate separated flow problems, they clearly do

not solve all problems, as evidenced by the known efficiency drop

in modern engines at altitude. Howell et al. @6#, for example, studied airfoils modified for higher lift, noting that their highly loaded

airfoils might be close to a limit, and that even higher loading

could cause unacceptable separation problems even in the presence of wakes. Looking beyond free-stream turbulence and

wakes, other types of separation control could prove useful. Gadel-Hak @7# provides a recent review. Techniques include boundary

layer tripping, vortex generation, suction, and injection of fluid

normal to the wall to either increase the boundary layer momentum or promote turbulence.

While the general literature is extensive, only a few studies

have considered separation control under LPT conditions. Some

have utilized passive techniques. Lake et al. @8# considered

dimples and boundary layer trips. Van Treuren et al. @9# considered vortex generators. Volino @5# used rectangular bars to impose

disturbances in a boundary layer and move transition upstream.

Passive flow control is appealing for its simplicity and the relative

ease with which it might be implemented in gas turbine environments. It has its limitations, however. Volino @5# found that passive devices can successfully control separation even at the lowest

Reynolds number of interest, but that these devices caused substantial increases in losses at higher Re. This is an important limitation for aircraft engines, where the Re range between takeoff

and cruise is large. An active device could be turned off at high

Re. Static passive devices are also unable to take advantage of the

unsteadiness caused by wake passing. An active device might be

timed to turn on and off in response to wake passing events.

Unsteady devices might also take advantage of the calmed region

following a transient turbulent event.

The literature contains several examples of active separation

control. Lee et al. @10# used blowing in supersonic engine inlets to

prevent or control separation. Sturm et al. @11# reported on blowContributed by the International Gas Turbine Institute and presented at the International Gas Turbine and Aeroengine Congress and Exhibition, Atlanta, GA, June

16–19, 2003. Manuscript received by the IGTI December 2002; final revision March

2003. Paper No. 2003-GT-38729. Review Chair: H. R. Simmons.

Journal of Turbomachinery Copyright © 2003 by ASME OCTOBER 2003, Vol. 125 Õ 765

NASA/CR—2012-217656 37
ing in a compressor cascade. Johnston and Nishi @12# used vortexgenerator-jets ~VGJ’s! to control separation in turbulent boundary

layers. This method utilizes blowing from ‘‘small, skewed, and

pitched holes’’ to create streamwise vortices similar to those created by solid vortex generators. Any jet injected into a flow will

tend to produce some turbulence, and the turbulent mixing will

tend to bring some high momentum fluid into the near wall region

and inhibit separation. Streamwise vortices bring additional high

momentum fluid into the near wall region. The most effective

VGJ’s enter the boundary layer at a relatively shallow pitch angle

~typically 30–45°! relative to the wall and a high skew angle

~45–90°! relative to the main flow. Compton and Johnston @13#

showed that the co-rotating vortices produced by VGJ’s are stronger and more effective for separation control than the counterrotating vortices which form downstream of a normal jet. McManus et al. @14# and Raghunathan et al. @15# used pulsed VGJ’s.

Sinha and Pal @16# used acoustic excitation to perturb an unsteady

separating flow. Jacobson and Reynolds @17# used piezoelectrically driven cantilevers to influence the near wall turbulence

structure on a flat plate. They noted that the devices could be used

in separation control. Miau et al. @18# used an oscillating fence to

promote reattachment downstream of a backward facing step.

Sinha et al. @19# used a driven flexible wall transducer to detect

pressure fluctuations and then produce near wall vortices upstream

of separation. Whitehead et al. @20# used a film transducer to produce airfoil vibrations and reduce separation at high angles of

attack. Oscillatory blowing has been used in several studies to

control separation on airfoils. Amitay and Glezer @21# provide one

recent example. Oscillatory jets are often referred to as ‘‘synthetic

jets’’ since they have no net mass flow. They are typically directed

normal to a surface, meaning that they probably do not produce

such strong streamwise vortices as VGJ’s.

Only a few active control studies have been conducted under

LPT conditions. Huang et al. @22# and Hultgren and Ashpis @23#

employed high voltage electrodes to produce glow discharge

plasma in a boundary layer to control separation. Bons et al.

@24,25# used steady and pulsed VGJ’s to successfully control

separation on LPT airfoils. They used the ‘‘Pak-B’’ airfoil, which

is an industry supplied research airfoil that is prone to separation

problems at low Re. It has been used in numerous studies, as

noted by Volino @4#. Bons et al. @24# used spanwise rows of VGJ’s

at several streamwise locations on the suction surface of the airfoil, and found that a row near the suction surface velocity maximum ~pressure minimum! was most effective. The VGJ holes

were oriented at 30° to the surface and 90° to the main flow. All

holes were oriented in the same direction, to produce co-rotating

vortices. Reynolds numbers as low as 60,000 ~based on suction

surface length and exit velocity! were considered. Bons et al. @25#

found that both steady and pulsed jets were effective in controlling separation. The pulsed jets were fully effective even when the

dimensionless pulsing frequency F1 was as low as 0.1, where F1

is a ratio of the transit time for flow between the VGJ hole and the

trailing edge to the time interval between pulses. Ensemble averaged velocity profiles showed a long relaxation or ‘‘calmed’’ period following each jet pulse. During this calm period the boundary layer remained attached long after the turbulence generated by

the pulse had moved downstream. Calmed regions have been observed following turbulent spots in transitional boundary layers

~e.g., Gostelow et al. @26# and Schulte and Hodson @27#!. The

mean velocity profiles in the calmed region gradually relax from

the turbulent shape associated with the turbulent spot they follow,

to a laminar ~and in some cases separated! profile shape. The

calmed boundary layer is very resistant to separation, much like a

turbulent boundary layer, but it is very laminarlike in terms of its

fluctuation levels and low losses. The pulsed jets were more effective than continuous jets, even when the pulsed jet duty cycle

was as low as 1%. This was believed to indicate that the starting

vortex formed at the beginning of each jet pulse was responsible

for most of the flow control in the pulsed jet cases.

Synthetic jets hold an advantage over continuous or pulsed jets

in that they require no net mass flow. In the LPT environment, this

means that no compressor bleed air is required. Use of bleed air

for flow control or cooling comes at a cost in efficiency, although

the small amount of air required for the pulsed jets of Bons et al.

@25# might not be prohibitive if bleed air were already routed to

the airfoils for cooling. Synthetic jets would not be useful for a

cooled airfoil since ingestion of hot gas into the airfoil would be

harmful. For uncooled LPT airfoils, however, the airfoil temperature will match the main flow temperature, and ingestion of hot

gas should be acceptable. Routing of bleed air to uncooled airfoils

for flow control may present a prohibitive addition of complexity

and weight. With synthetic jets this problem could be avoided. In

the present study, the oscillating flow of synthetic jets and the

compound angle injection of vortex generator jets are combined to

produce synthetic VGJ’s. This is believed to be the first application of synthetic VGJ’s. They are used to control the flow over a

Pak-B airfoil. A survey of the literature indicates that the jet locations and angles chosen by Bons et al. @25# were likely optimal, so

their geometry has been copied in the present study.

There are many parameters which could be varied in a synthetic

jet study, including Reynolds number, FSTI, jet geometry, jet location, jet velocity, jet oscillation frequency, and jet waveform, to

name a few. These are all potentially important parameters and

should eventually be studied. The scope of the present study is

more focused. A single experimental case is completely documented with detailed measurements including time resolved mean

and fluctuating velocity and turbulent shear stress throughout the

flow field. The goals of the study are to build an understanding of

the physics of how synthetic VGJ’s control separation and to generate questions for future parametric studies which may lead to

optimized flow control for a broad range of flow conditions.

Volino @4,28# studied unmodified flow over the Pak-B airfoil at

Reynolds numbers ranging from 25,000 to 300,000 under both

high and low FSTI. In nearly all cases the boundary layer separated from the suction side of the airfoil. At all but the lowest

Reynolds numbers it reattached before the trailing edge. The most

severely separated case was the low FSTI, Re525,000 case. This

case has therefore been chosen as the test case for the present

application of synthetic VGJ’s. The unmodified case from Volino

@4# is used as a baseline case for comparison to the new results.

Also used for comparison is a case from Volino @5# in which a

passive bar was employed to force reattachment. The bar was

located at s/Ls50.51, extending along the airfoil span. Its streamwise width was 6.35 mm and its height was 1.6 mm. The suction

surface length was 228.6 mm. Bars of various heights were tested.

The 1.6-mm bar was the smallest bar to cause reattachment at

Re525,000. Volino @5# found that the most effective bars in terms

of minimizing losses were not large enough to immediately trip

the boundary layer to turbulent. Rather, they induced small disturbances which grew and caused transition and reattachment downstream of a small separation bubble.

Experiment

Experiments were conducted in a low speed wind tunnel, described by Volino et al. @29#. Briefly, air enters through blowers

and passes through a honeycomb, a series of screens, two settling

chambers, and a three-dimensional contraction before entering the

test section. At the exit of the contraction, the mean velocity is

uniform to within 1%. The FSTI is 0.5%60.05%. Nearly all of

this free-stream ‘‘turbulence’’ is actually streamwise unsteadiness

at frequencies below 20 Hz and is not associated with turbulent

eddies. The rms intensities of the three components of the unsteadiness are 0.7%, 0.2%, and 0.2% in the streamwise, pitchwise

and spanwise directions, respectively. The test section immediately follows the contraction.

The test section, shown in Fig. 1, consists of the passage between two airfoils. Details are listed in Table 1, and more information is available in Volino @4#. A large span-to-chord ratio of

766 Õ Vol. 125, OCTOBER 2003 Transactions of the ASME

NASA/CR—2012-217656 38
4.3 was chosen to insure two-dimensional flow at the spanwise

centerline of the airfoils, where all measurements were made. Upstream of each airfoil are flaps, which control the amount of bleed

air allowed to escape from the passage. The flaps, along with a

tailboard on the pressure side of the passage, are adjusted to produce the correct leading edge flow and pressure gradient along the

airfoils. The flow in the passage matches that in a multiblade

cascade. The single passage configuration allows for a large scale

passage and better probe access than possible with a multiblade

cascade in the same size wind tunnel. The wake downstream of

the passage is not representative of a multiblade facility, however,

since there is flow only on one side of each airfoil. Downstream

effects that could influence the upstream flow in the passage are

also potentially missed. Experimental conditions match those of

the low FSTI, Re525,000 baseline case of Volino @4# and the

passive bar case of Volino @5#.

The synthetic VGJ’s were produced from a cavity within the

suction side airfoil. The airfoils are machined from high density

foam, which has a consistency much like hard wood. The surface

of each airfoil was sanded smooth, painted, and sanded again to

provide a smooth surface. A 1.27-cm-diameter hole was drilled

through the airfoil span at about mid-chord, as shown in Fig. 2, to

form a plenum. One end of the plenum is plugged, and the narrow

end of a funnel is inserted in the other. A 20.3-cm-diameter loudspeaker ~100-W subwoofer! is attached at the wide end of the

funnel. The funnel is sealed to the speaker and to the airfoil with

silicone RTV to prevent air leakage. The speaker is driven with a

200-W audio amplifier, which is in turn powered with a 12-V dc

power supply and driven by a function generator. For the present

study the function generator was set to output a sine wave. The

amplitude of the signal from the function generator and the gain

of the amplifier were adjusted to provide the desired input voltage

to the speaker. Holes for the VGJ’s were drilled into the suction

surface in a spanwise line at s/Ls50.514. The holes are 0.8 mm in

diameter ~0.35% of Ls) and are spaced 8.5 mm apart ~3.7% of

Ls). The holes are drilled at a 90° skew angle with respect to the

main flow and a 30° pitch with respect to the surface, as shown in

Fig. 2. Each hole extends from the suction surface into the cavity

in the core of the airfoil. The length to diameter ratio of the holes

is 7.5.

Measurements. Pressure surveys were made using a pressure

transducer ~0–870-Pa range Validyne transducer! and a Scanivalve. Stagnation pressure was measured with a pitot tube upstream of the passage inlet, and eleven pressure taps were located

on each airfoil along their spanwise centerlines. Locations of the

taps on the suction side are listed in Table 2 along with measured

local FSTI components and the acceleration parameter K at these

stations based on a nonseparating, inviscid solution. The uncertainty in the suction side pressure coefficients was 7%. Most of

this uncertainty was due to bias error. Stochastic error was minimized by averaging pressure transducer readings over a 10-s period.

Velocity profiles on the suction surface were measured at

streamwise stations corresponding to pressure taps 7–11, as given

in Table 2, and at four additional stations, labeled 7.5, 8.5, 9.5,

and 10.5, centered between the pressure taps. All stations are

downstream of the VGJ holes. Stations 7, 8, 9, 10, and 11 correspond to stations documented in the baseline case in Volino

@4,28#. Profiles at Stations 1–6 are fully documented for the baseline case in Volino @4,28#, and show that the upstream boundary

layer closely follows a laminar solution. Profiles were measured

near but not at the spanwise centerline of the airfoil to insure that

the pressure taps did not interfere with the velocity measurements.

Fig. 1 Schematic of the test section

Table 1 Test section parameters

Axial

Chord

@mm#

True

Chord

@mm#

Pitch

@mm#

Span

@mm#

Suction

side, Ls

@mm#

Inlet

flow

angle

Exit

flow

angle

153.6 170.4 136.0 660.4 228.6 35° 60°

Fig. 2 Drawing of suction side airfoil with cavity and VGJ’s:

„a... full airfoil, „b... cross section of VGJ holes

Table 2 Measurement station locations, local acceleration „inviscid soln...., and measured local free-stream turbulence

Station s/Ls K3106  ̄

u8/U` @%#  ̄

v8/U` @%#

1 0.111 6.32 0.44

2 0.194 4.80 0.39

3 0.278 3.44 0.37

4 0.361 3.00 0.38

5 0.444 2.48 0.39

6 0.528 20.08 0.41

7 0.611 23.24 0.47 0.05

8 0.694 23.80 0.47 0.12

9 0.777 22.32 0.48 0.14

10 0.861 22.12 0.54 0.11

11 0.944 20.72 0.51 0.11

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NASA/CR—2012-217656 39
Profiles were acquired with a hot-wire anemometer ~AA Lab Systems model AN-1003! and a single sensor boundary layer probe

~TSI model 1218-T 1.5!. The sensor diameter is 3.8 mm, and the

active length is 1.27 mm. At each measurement location, data

were acquired for 26 s at a 20-kHz sampling rate (219 samples!.

All raw data were saved. The high sampling rate provides an

essentially continuous signal, and the long sampling time results

in low uncertainty in both statistical and spectral quantities. Data

were acquired at 60 wall normal locations in each profile, extending from the wall to the free stream, with most points concentrated

in the near wall region. The closest point was within 0.1 mm of

the wall, which corresponds to y/Ls50.0004 and about 0.015

boundary layer thicknesses. Flow direction in a separation bubble

cannot be determined with a single-sensor hot wire, but velocity

magnitude can be measured and was found to be essentially zero

within the bubbles of the present cases. Determining the direction

was not therefore considered essential. Uncertainty in the mean

velocity is 3–5% except in the very near wall region, where nearwall corrections ~Wills @30#! were applied to the mean velocity.

Uncertainties in the momentum and displacement thicknesses

computed from the mean profiles are 10%. Uncertainty in the

shape factor H is 8%. Local skin friction coefficients were computed from the near wall mean velocity profiles using the technique of Volino and Simon @31#. This technique accounts for

streamwise pressure gradient effects on the mean profile. The uncertainty in Cf is 8%. The uncertainty in the fluctuating streamwise velocity is below 10%. As explained in Volino @4# based on

the work of Ligrani and Bradshaw @32#, spatial averaging effects

due to the finite length of the hot-wire sensor should not be significant in the present case.

Profiles were also acquired using a cross-sensor boundary layer

probe ~TSI 1243-20!. The sensors are 51-mm-diameter hot films

with 1.02-mm active lengths. The probe is used to document the

instantaneous turbulent shear stress, 2u8v8. Profiles were acquired at the same stations as with the single-sensor probe. Data

were acquired at 25 locations in each profile, extending from 1

mm from the wall to the free stream. Sampling rates and times

were the same as for the single-sensor profiles. The vortices induced by the oscillating jets cause significant secondary velocity,

particularly at the streamwise stations immediately downstream of

the jet holes. The magnitude of these secondary velocity components remains below 20% of the local streamwise velocity, however, so they should not cause significant error in the hot-wire

measurements. The uncertainty in 2u8v8 is 10%.

The VGJ velocities were measured using a hot-film probe ~TSI

model 1210-10A! with a 0.25-mm active sensor length. The sensor was placed directly over the exit of the jet hole. During outflow from the hole, the jet was expected to blow directly across

the sensor, providing an accurate measure of the jet velocity. Uncertainty in the velocity is 5% and results mainly from uncertainty

in the position of the sensor, which could lead to a slightly lower

velocity reading than the velocity at the jet exit plane. During

inflow of the oscillating jet, the flow is expected to behave more

like a sink flow than a jet. The measured velocity does not therefore provide an accurate indicator of the velocity inside the hole

during inflow. The jet velocity was calibrated against the rms input voltage to the speaker with the main flow in the wind tunnel

turned off, and the calibration was used to set the jet velocity in

later experiments. The jet velocity is fixed by the frequency and

amplitude of the displacement of the speaker diaphragm, which

causes a pressurization of the cavity relative to the pressure at the

jet exits. The dynamic pressure of the flow through the test section

is about 2.4 Pa at the jet location. Since the test section exits to

atmosphere, the dynamic pressure results in an average pressure

of 2.4-Pa vacuum in the cavity, and a 2.4-Pa pressure difference

across the speaker diaphragm. This pressure is much smaller than

the pressure experienced by the speaker when driving the jets, and

is not expected to influence its motion. If the amplitude of the

diaphragm motion is unchanged by the presence of flow in the test

section, the jets should drive approximately the same mass flow

through the holes with the wind tunnel on or off. The jet velocity

at the exit plane will presumably be affected by the mean flow,

however.

The measured maximum jet exit velocity was 9.4 m/s for most

of the cases presented below. Given the sensor length and the

diameter of the holes, the measured velocity is an average over the

middle 30% of the jet. The Reynolds number based on this velocity and the jet diameter is 500. At this Reynolds number, approximately 30 diameters would be needed to establish fully developed

laminar flow inside the jet holes. Since the length to diameter ratio

in only 7.5 and the jets are unsteady, the jet velocity is not expected to have a fully developed parabolic laminar profile, but

rather a more flat profile. Given the averaging due to the sensor

size and the expectation of a flat profile, the instantaneous mean

velocity of the jet is assumed to approximately equal the measured

velocity. Ideally this assumption would be checked with a survey

of velocity across the jet exit plane, but the very small jet diameter

precludes an accurate survey. The uncertainty in instantaneous

mean velocity is, therefore, higher than the 5% uncertainty in the

measured velocity. The uncertainty is estimated to be between 10

and 20%.

Data Processing. In addition to conventional time averaging,

the velocity data were ensemble averaged relative to the time

within each jet oscillation cycle. For this purpose, the speaker

input voltage was digitized simultaneously along with the instantaneous velocity data. Data were ensemble averaged at 24 instances within the cycle. At each instance, data were averaged

over 1/180th of the cycle. For each 26-s data trace, this results in

roughly 3000 data points to average for each ensemble. With this

many data points to average over a 26-s time record, the ensemble

averaged results are well resolved and have uncertainties as low as

those given above for the time averaged results. The start of the

cycle was arbitrarily chosen as the instant when the speaker input

voltage crossed from negative to positive. As will be shown below, this roughly corresponds to the beginning of the jets’ outward

pulse.

Results

Jet Velocity. Figure 3 shows a typical time trace of the measured jet velocity and the speaker input voltage. The frequency of

the input signal was set to a nominal value of 10 Hz ~actual value

was 10.5 Hz!. The maximum and average velocities in each outward pulse were 9.4 and 5.9 m/s, respectively. The jet velocity can

be expressed as a blowing ratio B, defined as

Fig. 3 Time trace of jet velocity and speaker input voltage,

F¿Ä0.65

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NASA/CR—2012-217656 40
B5rjetVjet /r `U` ,

where rjet5r ` since the jet fluid comes from the boundary layer,

and U`(52 m/s) is the local free-stream velocity at the jet location. For the present cases Bmax54.7 during each cycle and Bave

53.0 during the outward pulse. The mass flux of the jets can be

compared to the velocity deficit in the boundary layer as

M5@rjetVjet~pDjet

2 /4!S#/@r `U`d*#,

where S50.118 holes/mm is the number of holes per unit span.

The displacement thickness, d*~51.08 mm! is the baseline case

value at the jet location. This gives Mmax50.258 and Mave

50.162. This could be interpreted to mean that 16% of the boundary layer in terms of displacement thickness is sucked off during

the inflow half of each jet cycle, and then re-injected into the

boundary layer during the outward pulse.

The momentum coefficient cm is defined as the ratio of the jet

momentum to the free-stream dynamic pressure. Using the definition of Bons et al. @25#,

cm5@rjetVjet

2 ~pDjet2 /4!S#/@r `U`2 Chord#,

where the axial chord is 153.6 mm. The maximum value of cm in

each cycle is 0.0085. The average value of Vjet

2 for the outpulse is

44 m2

/s2. This gives an average cm50.0042.

Figure 3 shows that the speaker input voltage and the jet velocity are slightly out of phase. The jets lag the input voltage by

about 0.006 s, which is a dimensionless lag Dt/T of 0.063, where

T is the jet oscillation period of 0.095 s. The lag is expected, as the

jets respond dynamically to the pressurizing of the cavity in the

airfoil by the speaker, and there is no reason to expect the speaker

voltage and the jet velocity to be exactly in phase. The finite

distance from the speaker to the jet holes ~of the order 0.5 m! and

the finite speed of sound ~340 m/s! will also lead to a time lag of

the order 1 ms. The time lag increases with distance from the

speaker along the airfoil span. At high jet frequencies, the time lag

causes the jets along the span to be significantly out of phase with

each other. Measurements, however, show that the jet amplitude

along the span is uniform, regardless of the frequency. Variation in

phase along the span could lead to difficulty in practice if attempts

were made to time the jet pulsing to other cyclic events such as

wake passing. It is not an issue in the present study. At the relatively low frequency of 10 Hz, the phase lag was not significant,

and the jets were uniform in both phase and amplitude along the

span.

Pressure Profiles. Pressure profiles were acquired for several

jet amplitudes and jet frequencies. The general finding was that

the jets were effective over a broad frequency range, so long as

the amplitude was sufficiently high. Figure 4 shows Cp profiles

for a range of jet amplitudes, with the jet frequency set to

10.5 Hz (F150.65). In all cases there is good agreement between

the data and an inviscid solution for the Pak-B airfoil on the

pressure side and the upstream portion of the suction side. In the

adverse pressure gradient region on the suction side, differences

are clear. Without the jets the boundary layer separates and does

not reattach, as indicated by the region of constant Cp values.

With Bmax51.9 the boundary layer still does not reattach, but

there is some sign that Cp is starting to drop at the last pressure

tap. For the cases with Bmax54.7 and above, the boundary layer

does not appear to separate. Significant case to case differences in

Cp are present right at the suction peak, but these are likely due to

the injection of the jets at this location and their effect on the flow

over the adjacent pressure tap. The differences diminish rapidly

and are essentially gone by the next downstream measurement

station. Bons et al. @24# demonstrated effective flow control with

Bmax as low as 0.4 in their study. The significantly lower Re in the

present study may explain the need for stronger jets. Figure 5

shows Cp profiles for several different jet frequencies with Bmax

held approximately constant at about 5. There does not appear to

be any clear separated region in any of the cases with jets. The

suction peak is higher in all the cases with jets than in the baseline

case. This is an expected result, suggesting that the lift will be

higher when the boundary layer is attached. The F150.65 case

appears to agree most closely with the inviscid solution. The

broad range of effective frequencies agrees with the results of

Bons et al. @25#.

The objective of the present study is not to establish the optimal

jet conditions for the present case, but to investigate in detail a

case in which the jets provide effective flow control. The dimensionless frequency F150.65 was chosen since it appeared to provide slightly better results than the other cases in Fig. 5, and

Bmax54.7 was chosen since it was the lowest effective blowing

ratio tested. Figure 6 shows the Cp profile for the chosen case

with jets along with the baseline case and the passive bar case of

Volino @5#.

Velocity Profiles. Figure 7 shows mean velocity  ̄u8 and turbulent shear stress profiles for the present case, the baseline case,

and the passive bar case. In the baseline case, the mean profiles

show the boundary layer is on the verge of separating at station 7,

is clearly separated at station 8, and the separation bubble grows

through station 11. In the bar case, the boundary layer separates

from the bar. The separation bubble is visible at station 7, and it

grows through station 9. At station 10, the near wall velocities

Fig. 4 Cp profiles, F¿Ä0.65, various blowing ratios

Fig. 5 Cp profiles, B maxÉ5, various F¿

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NASA/CR—2012-217656 41
begin to rise, indicating the beginning of reattachment, and the

mean profile shows clear reattachment at station 11. With the

VGJ’s, the behavior is much different. There is no clear separation

bubble. The mean profile appears to have an attached, laminarlike

shape. The momentum deficit with the jets appears to be significantly lower at station 11 than in the case with the bars, indicating

lower losses with the jets.

The  ̄u8 values in the baseline and bar cases are very low at

station 7, as expected since the boundary layer is still laminar. A

peak appears downstream in the shear layer over the separation

bubble. In the bar case,  ̄u8 begins to rise in the near wall region at

station 9, signaling imminent reattachment. At stations 10 and 11,

 ̄

u8 in the bar case rises to the high values typical of a transitional

boundary layer. In the jet case,  ̄u8 is high at all stations. At the

upstream locations it is much higher than would be expected for a

turbulent boundary layer. As will be shown below, however, much

of the contribution to  ̄u8 is from 10-Hz unsteadiness associated

with the jets and is not turbulence. The turbulent shear stress

profiles show considerable momentum transport in the jet case at

all stations, which helps to explain how the boundary layer remains attached. Unlike in a turbulent boundary layer, the 2u8v8

peak is well away from the wall. In the baseline case the shear

layer does not transition to turbulent, and 2u8v8 remains near

zero. In the bar case 2u8v8 profiles were not acquired.

Figure 8 shows shape factor and momentum thickness as computed from the mean profiles of Fig. 7. The shape factor H provides a measure of the state of the boundary layer with respect to

separation and transition. The shape factor in the baseline and bar

cases rises rapidly after separation, as d* increases while u remains nearly constant. In the baseline case the boundary layer

never reattaches. In the bar case, transition and reattachment occur, causing d* to fall and u to rise. The shape factor begins to

drop toward a turbulent value of about 1.6, but does not reach this

value, indicating that the recovery from the separation is not comFig. 6 Cp profiles, comparison of baseline, passive bar, and

present jet case with B maxÄ4.7 and F¿Ä0.65

Fig. 7 Station 7–11 dimensionless time averaged profiles; comparison of baseline, passive

bar and present jet cases: „a... mean velocity, „b... u ̄ 8, „c... turbulent shear stress

Fig. 8 Time averaged shape factor and momentum thickness

verses streamwise location; comparison of baseline, passive

bar and present jet cases: „a... H, „b... u

770 Õ Vol. 125, OCTOBER 2003 Transactions of the ASME

NASA/CR—2012-217656 42
plete. In the jet case, the shape factor remains at a laminar value

of about 2. It never rises to a separated flow value, nor does it

drop to a turbulent value. The momentum thickness provides a

measure of the losses in a boundary layer. If the boundary layer

reattaches before the trailing edge, the suction side boundary layer

losses will be the dominant losses in an LPT passage ~Howell

et al. @6#!. In the baseline case the boundary layer does not reattach, so although u remains low, high losses would be expected in

the wake downstream of the airfoils. For the bar and jet cases,

however, the boundary layer is attached at the trailing edge. Momentum thickness is about 20% higher in the bar case, indicating

that the jets are better able to control separation, while causing

lower losses.

Ensemble Average Velocity Profiles. The time averaged profiles of Fig. 7 indicate that the jets are effective in controlling the

boundary layer, but they do not explain the mechanism by which

the jets work. Figure 9 shows ensemble averaged mean velocity

profiles. Profiles are shown for nine streamwise stations at 24 time

increments within the jet oscillation cycle. Figures 10 and 11

show the corresponding  ̃u8 and 2 ̃

u8 ̃

v8 profiles. Examining the

profiles at station 7, the mean profile initially appears to be laminar and attached, and both the  ̃

u8 and 2 ̃

u8 ̃

v8 values are near zero,

indicating again that the flow is laminar. Given the phase lag

shown in Fig. 3 between the speaker input and the jets, and the

finite convection time between the jet hole and station 7, one

would expect that the disturbance created by the jet outpulse

should arrive at station 7 at t/T50.18. In fact, however, the disturbance is not seen in the mean profile until t/T50.333. In agreement, the  ̃

u8 and 2 ̃

u8 ̃

v8 values also rise above zero at t/T

50.333. This may suggest that the rising jet velocity must reach a

sufficiently high amplitude before it can significantly affect the

boundary layer. Comparing the observed phase lag and the jet

velocity of Fig. 3 suggests that the jet velocity must be about 8

m/s, corresponding to an instantaneous B54, for the jets to be

effective. Continuing forward in time at station 7, the jets cause a

large disturbance in the mean profiles that continues until t/T

50.667. The time t/T50.667 corresponds very closely with the

end of the jet outpulse when the phase lag and convection time

from the jet holes to station 7 are considered. The large local

minima and maxima in the mean velocity profiles indicate the jets

are not merely adding turbulence to the boundary layer, but are

inducing some flow structure, most likely streamwise vortices.

Fig. 9 Ensemble averaged dimensionless mean velocity „U ̃ ÕUe... profiles

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NASA/CR—2012-217656 43
These vortices likely cause spanwise variation in the velocity, particularly at the stations nearest the jets. In the present study, data

were only acquired at one spanwise location. Spanwise surveys

should be considered in future work. The  ̃u8 and 2 ̃

u8 ̃

v8 profiles

exhibit large peaks in the regions where the mean velocity gradients are highest in Fig. 9. The 2 ̃

u8 ̃

v8 values have the appropriate

sign, corresponding to the sign of dU ̃/dy in the mean profiles. It

should be noted that the magnitude of  ̃u8 is smaller in the ensemble averaged profiles of Fig. 10 than in the time averaged

profiles of Fig. 7. This indicates that much of the contribution to

 ̄

u8 in Fig. 7 is due to 10-Hz oscillations and not to turbulence.

After t/T50.667, the mean velocity profile resumes a laminar

shape. There is no tendency toward boundary layer separation.

The boundary layer did not separate in the baseline case at this

station ~Fig. 7!, but the mean profile in the baseline case did

appear closer to separation than in the present case.

Moving to the downstream stations, the leading edge of the

disturbance, as observed in the mean profiles and the  ̃u8 and

2 ̃u8 ̃

v8 profiles moves to later values of t/T, as expected since the

disturbance takes some time to convect downstream. The leading

edge of the disturbance appears to move at about 90% of the local

free-stream velocity, which corresponds with the expected leading

edge celerity of a turbulent spot ~e.g., Gostelow et al. @26#,

Schulte and Hodson @27#!. The trailing edge of the disturbance

appears to move at about 45% of the local free-stream velocity,

agreeing with the expected trailing edge celerity of a turbulent

spot. Because the leading and trailing edge celerities are different,

the boundary layer is only disturbed by the jets during approximately 30% of the cycle at station 7, but is disturbed during approximately 70% of the cycle at station 11. The beginning of the

disturbance at each station is seen simultaneously in the mean

profile and the  ̃u8 and 2 ̃

u8 ̃

v8 profiles of Figs. 9–11. At the trailing edge of the disturbance, return of the mean velocity profile to

a smooth shape corresponds closely with the return of 2 ̃u8 ̃

v8 to

near zero ~see, for example, the profiles at station 8 at t/T

50.833 or station 10.5 at t/T50.417). The  ̃u8 profiles, in contrast, take somewhat longer to return to an undisturbed condition.

At station 11, for example, there is at least a small near wall  ̃u8

peak at all times, while the 2 ̃u8 ̃

v8 values are essentially zero

between t/T50.625 and 0.792.

Fig. 10 Ensemble averaged rms value of dimensionless streamwise fluctuating velocity

„u ̃ 8ÕUe... profiles

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NASA/CR—2012-217656 44
The flow structure at the downstream stations appears to be less

distinct than at stations 7 and 7.5. Inflection points are still present

in the mean profiles, but the local minima and maxima in the

mean profiles are less sharp. Still, the mean profile shapes and the

multiple peaks in the  ̃u8 and 2 ̃

u8 ̃

v8 profiles, particularly those far

from the wall ~see for example the station 10 profiles of Figs. 10

and 11 between t/T50.75 and 0.958! are evidence that the flow

structures induced by the jets persist downstream.

After the disturbance caused by the jets passes, the boundary

layer eventually separates at locations between stations 8 and 11.

At station 8.5, for example, the trailing edge of the disturbance

passes at about t/T50.9. By t/T50.1 the near wall profile appears separated. The separation appears to coincide with the return

of  ̃u8 to near zero. The separation appears to persist until about

t/T50.6, after the start of the next disturbance event. The reattachment within the disturbance event corresponds to the motion

of high  ̃u8 and 2 ̃

u8 ̃

v8 into the near wall region at t/T50.667.

The same sequence of events is visible at downstream stations. At

station 10.5, for example, the trailing edge of the disturbance has

passed by t/T50.4, but the boundary layer remains attached until

about t/T50.75. Reattachment is visible at t/T51, and a near

wall peak emerges at the same time in 2 ̃u8 ̃

v8. It should be noted

that the separation bubble remains thin at all stations, and never

begins to approach the thickness observed in the baseline flow of

Fig. 7.

The period between the passage of the disturbance and boundary layer separation is believed to indicate a ‘‘calmed’’ region.

Calmed regions have been observed in previous studies to follow

turbulent spots and wake induced turbulent strips. The duration of

the calmed region increases at the downstream stations, since the

trailing edge celerity of the calmed region, shown in Fig. 9 to be

about 0.3 the local free-stream velocity ~in agreement with previous studies of calmed regions!, is slower than the trailing edge

celerity of the disturbed region. Hence at station 8.5 the calmed

region extends for Dt/T of about 0.2, while at station 10.5 it

extends for Dt/T of about 0.4. The presence of the calmed region

may help to limit the separation bubble thickness.

Figure 12 shows the ensemble  ̃u8 data in a different format.

Contours of near wall  ̃u8 at y50.095 mm (y/Ls50.0004) are

shown in a time-space plot. The horizontal axis shows the dimensionless streamwise location, and the vertical axis indicates the

dimensionless time within the cycle. Two complete cycles are

Fig. 11 Ensemble averaged dimensionless turbulent shear stress „Àu ̃ 8v ̃ 8ÕUe

2... profiles

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NASA/CR—2012-217656 45
shown to better illustrate the periodicity of the event. The data in

the bottom half of the figure is shown again in the top half. In this

format, it is clear that the jet outpulse causes a high  ̃u8 event to

appear at s/Ls50.6 ~station 7! and t/T50.6. This event then proceeds downstream in a widening wedge of turbulence. The lower

and upper slopes of the wedge indicate its leading and trailing

edge celerities, respectively. Within the wedge, a local  ̃u8 peak is

visible at s/Ls50.7 ~station 8!. This peak indicates that the flow

structure produced by the jets has penetrated very near the wall.

Slightly farther downstream at s/Ls50.75 ~station 8.5! a local

minimum is visible in the  ̃u8 contours. This may indicate that the

flow structure induced by the jets has lifted off the wall slightly.

The mean profiles of Fig. 9, support this, showing lower near wall

mean velocity gradients at station 8.5, which indicates that high

speed fluid is not as effectively brought into the near wall region

at this station. Farther downstream in Fig. 12, the  ̃u8 level in the

turbulent wedge rises again as a fully turbulent boundary layer

begins to develop. Outside of the wedge the  ̃u8 level is very low,

indicating laminar flow.

Figure 13 shows ensemble averaged 2 ̃u8 ̃

v8 contours at y/Ls

50.0128 in the format of Fig. 12. As in Fig. 12, the wedge of

turbulence is clear. Within the wedge momentum transport is high,

but outside the wedge the flow appears to be laminar.

Figure 14 shows a time-space plot of the local ensemble averaged free-stream velocity. The free-stream velocity varies both

spatially and temporally. Spatially, the shape of the airfoil passage

causes the flow to decelerate. Temporally, when the jets cause a

turbulent event to moves down the surface, it causes the boundary

layer to thicken and accelerates the free-stream. During the laminar and calmed periods the boundary layer is thinner and the

free-stream velocity is lower. Figure 15 shows the local Thwaites

parameter, lu , as computed from the free-stream velocity data of

Fig. 14. In a laminar boundary layer, separation is expected when

lu is less than 20.082. With the exception of the turbulent strip,

where the boundary layer is locally accelerated, lu is below

20.082 at most times and locations on the surface, and at some

times is below 22. Hence it is not a surprise that the boundary

layer tends to separate when not controlled by turbulence or a

calmed region.

Integral Parameters. Local displacement and momentum

thickness values can be computed from the ensemble mean velocity profiles of Fig. 9. Momentum thickness is shown in Fig. 16.

Both d* and u grow in the streamwise direction and are about

twice as large in the turbulent region than in the laminar flow

region. The slow growth of the momentum thickness in the nonturbulent flow indicates that losses should be low in this region, as

might be expected based on the low  ̃u8 and 2 ̃

u8 ̃

v8 values of Figs.

Fig. 12 Time-space plot of ensemble averaged u ̃ 8ÕUe at yÕLs

Ä0.0004

Fig. 13 Time-space plot of ensemble averaged Àu ̃ 8v ̃ 8ÕUe

2 at

yÕLsÄ0.0128

Fig. 14 Time-space plot of ensemble averaged dimensionless

free-stream velocity U ̃ ` ÕUe

Fig. 15 Time-space plot of ensemble averaged Thwaites acceleration parameter lu

774 Õ Vol. 125, OCTOBER 2003 Transactions of the ASME

NASA/CR—2012-217656 46
10 and 11. At s/Ls50.6 ~station 7!, d* and u in the laminar flow

are both roughly 67% of their values in the baseline case of Volino

@4#. The lower values in the present case may be due to the suppression of the large separation bubble downstream, which

changes the local pressure gradient at station 7. The suction during

the inflow portion of the jet cycle could also be an explanation,

but d* and u are uniformly low in the laminar flow at station 7. If

the suction were causing a thinner boundary layer, one would

expect the effect to be stronger during the time of strongest inflow

into the jet holes. Figure 17 shows the local shape factor H, computed from the d* and u values. Within the turbulent region, the

shape factor remains between 2 and 2.4, which is well above the

expected value of 1.6 for a fully turbulent boundary layer subject

to the strong adverse pressure gradient of the present case. In the

nonturbulent flow, H reaches values as high as 3.4, which is consistent with the observed flow separation.

Skin Friction Coefficient. It is clear from the mean profiles

of Fig. 9 that the local wall shear stress varies greatly during the

jet oscillation cycle. At the upstream stations at t/T50.6, for example, the jets bring high speed fluid very close to the wall, producing a very high mean velocity gradient at the wall. At other

times at all stations, the boundary layer appears very laminarlike,

indicating a relatively low wall shear stress. When the boundary

layer is separated the wall shear goes to zero. Figure 18 shows the

local skin friction coefficients Cf . The white regions with dots

indicate where the flow is separated and the local skin friction is

essentially zero. At s/Ls50.6 ~station 7! there is a laminar boundary layer with Cf50.005 between jet outward pulses. The corresponding momentum thickness Reynolds is about 60. The skin

friction drops in the streamwise direction and the boundary layer

separates at s/Ls50.7 ~station 8!. At t/T50.6 and s/Ls50.6, the

skin friction is much higher, following the jet outpulse. The momentum thickness Reynolds number is as high as 180, and Cf

reaches values as high as 0.014. Moving downstream along this

turbulent region, Cf rapidly drops to 0.001 at t/T50.8 and s/Ls

50.75 ~station 8.5!. The turbulence at this time and position is

also low ~Fig. 10!. As stated above, the vortices produced by the

jets appear to quickly bring high speed fluid near the wall, resulting in high skin friction and turbulence, but by s/Ls50.75 this

effect may weaken, resulting in lower Cf and  ̃

u8. Moving farther

downstream to t/T51.1 and s/Ls50.85, Cf rises again to as high

as 0.007. The corresponding Reu is about 200 at this time and

location. Cf50.007 is about what one would expect for a fully

turbulent boundary layer with Reu5200, based on the standard

correlation Cf50.0256/Reu

0.25 ~Schlichting @33#!.

Figure 18 clearly shows the calmed region described above. It

is the triangular shaped region centered at t/T50.4 and s/Ls

50.9 that in the figure is bounded below by the strip of high Cf

and above by the separated flow region. Comparing Figs. 12, 13,

16, and 18 shows that the calmed region has low skin friction, low

momentum thickness, and low turbulence. Losses should therefore be low for the calmed region, as expected.

A comparison of the momentum thickness and skin friction in

Figs. 16 and 18 is interesting. The wedge of high u overlaps both

the upper portion of the separated flow region and the strip of high

Cf . As shown in the profiles of Figs. 10 and 11, downstream of

s/Ls50.75 ~station 8.5!, the turbulence and mixing induced by the

jet outpulse initially affects the outer part of the boundary layer

and does not immediately eliminate the separation bubble. Figure

16 shows that this outer region mixing causes a rise in the momentum thickness. It is only somewhat later and farther downstream that this mixing moves into the near wall region, causing

reattachment and high Cf . Kaszeta et al. @34# reported a similar

result for flow over the Pak-B airfoil subject to wake passing

events. They observed a time lag between the wake arrival and

near wall transition.

Kaszeta et al. @34# also observed a relation between the temporal acceleration and deceleration of the flow associated with the

wake passing, and its relation to transition and the thickening and

thinning of the boundary layer. Similarly, a comparison of Figs.

Fig. 16 Time-space plot of ensemble averaged momentum

thickness u in mm

Fig. 17 Time-space plot of ensemble averaged shape factor H

Fig. 18 Time-space plot of ensemble averaged skin friction

coefficient Ct ; white areas with dots indicate separated flow

Journal of Turbomachinery OCTOBER 2003, Vol. 125 Õ 775

NASA/CR—2012-217656 47
15, 16, and 18 show the relation between the local acceleration, u

and Cf . The region of overlap between high u and separated flow

corresponds to the region of strong deceleration centered at s/Ls

50.9 and t/T50.85. One could argue that the deceleration inhibits reattachment in spite of the mixing in the outer part of the

boundary layer. At slightly later t/T, the flow is accelerated. Under the combination of acceleration and high 2 ̃

u8 ̃

v8 the boundary

layer reattaches.

Open Questions. The first objective of the present study was

to assess the effectiveness of synthetic VGJ’s on separation control for an LPT airfoil and provide some description of the mechanisms through which the jets work. This has been addressed

above. The second objective was to use these results to generate

questions for further consideration. These are discussed below.

What is the effect of jet frequency on flow control? In the

present case, with F150.65, the airfoil boundary layer at any

given time was approximately half disturbed by the jets and half

laminar or calmed. The effects of one jet outpulse are present in

the boundary layer at all times. Separation occurred, but the separation bubble was small and did not appear to result in any harmful effects. How thick a separation bubble is tolerable? If F1 were

increased, the duration of each event would be reduced, but more

events would be present in the boundary layer at any given time.

It is not clear what effect this would have. In other studies with

synthetic jets, some have found that F1 of the order 1, as in the

present study, is most effective ~e.g., Seifert and Pack @35#!. Others, such as Amitay and Glezer @21# report that under some conditions jets with F1 of the order 10 are more effective. Lowering

the jet frequency might be beneficial. Losses appear to be generated primarily in the flow disturbed by the jets, while the flow

between these events is laminar or calmed with low losses. Reducing F1 could presumably result in a smaller fraction of the

airfoil covered by disturbed flow at any time. At some point, however, if the jets are too widely spaced, the flow will become uncontrolled at times and revert to the large separation bubble of the

baseline case. Based on the trailing edge celerity of the disturbed

flow, if F1 were reduced below 0.45, there would be instances

within the jet cycle when no disturbances would be present in the

boundary layer. If F1 were reduced below 0.3, there would be

times when neither disturbed or calmed flow would be present. In

flow around a single airfoil, Seifert and Pack @35# found that jets

with F1 between 0.5 and 1.5 were most effective at all Re, but

that with F150.25 the jets were ineffective. Bons et al. @25#, in

contrast, found that their VGJ’s were effective at F1 as low as

0.1. They suggested that the more controlled nature of the LPT

flow, where adjacent airfoils provide covered turning, might explain the lower effective F1 in their study. It should be noted that

the Bons et al. @25# experiments were conducted at a Reynolds

number of 60,000, where separation effects are not as severe as in

the present case with Re525,000.

What is the effect of jet amplitude on flow control? The pressure profiles of Fig. 4 indicate that the jet amplitude must be

sufficiently high for the jets to be effective. It is expected that

using a higher amplitude than necessary will result in higher

losses, but the extent to which the losses would increase is not

known. Volino @5# using passive bars found that cases with a small

separation bubble followed by reattachment had lower losses than

cases in which the boundary layer was tripped to turbulent to

prevent separation. Low amplitude jets could potentially produce

a similar effect.

What is the effect of jet wave form? In the present case the

speaker was driven with a sine wave input. A square wave with a

short duty cycle, as in the study of Bons et al. @25#, might be

better. By keeping the duty cycle short, the amount of calmed flow

relative to disturbed flow could be increased. This might allow a

reduction of losses without a sacrifice of separation control.

What is the effect of jet inflow? The outward flow portion of

each jet cycle appears to dominate the flow control. It is possible,

however, that the inflow may help reduce the boundary layer

thickness in the undisturbed flow periods and help reduce separation effects. A direct comparison of cases at the same Re using

synthetic VGJ’s and pulsed VGJ’s could help answer this question.

What is the effect of jet geometry? The VGJ’s are clearly effective. Through the generation of streamwise vortices they appear to provide more mixing than would be produced with normal

jets that simply produced turbulence. The relative magnitudes of

the effects of turbulence and streamwise vortices in enhancing

mixing are uncertain. Might the generation of turbulent spots with

a normal jet be sufficient? Which would result in lower losses?

What are the effect of Reynolds number and free-stream turbulence level? The baseline cases of Volino @4# show that at higher

Re and higher FSTI the boundary layer is more likely to transition

and reattach even without flow control. The generation of calmed

attached flow between jet events could still prove beneficial at

higher Re, however. There would be a tradeoff between the losses

generated by the jets themselves and the reduction in losses the

jets might provide by reducing the separation bubble thickness

and producing low-loss calmed regions. It might be possible to

control the flow at higher Re with significantly lower blowing

ratios than in the present case. Natural transition in the undisturbed flow between jet events and its interaction with calmed

regions would be an added complication at higher Re not seen in

the present study. If the Re were sufficiently high so that separation did not occur in the uncontrolled case, it is unlikely that the

jets would provide any benefit. Schulte and Hodson @27# noted

that the presence of calmed regions produced by unsteady wake

passing could not significantly lower the losses in an already attached boundary layer.

Conclusions

1. Synthetic vortex generator jets proved effective for controlling boundary layer separation on an LPT airfoil at very low Reynolds numbers. The separation bubble was effectively eliminated,

and losses were lower than in a similar case with passive flow

control.

2. The VGJ’s prevent separation by bringing high momentum

fluid into the near wall region and by promoting momentum transport through turbulent mixing.

3. The disturbance produced by the VGJ’s behaves in many

ways like the disturbance associated with a turbulent spot or a

wake induced turbulent strip. The leading edge celerity of the

disturbance is approximately 0.9U` , and the trailing edge celerity

is about 0.45U` . A calmed region with a trailing edge celerity of

0.3U` follows the disturbance. The calmed region is resistant to

separation.

4. The adverse pressure gradient in the present case was strong

enough so that the boundary layer did separate after the passage of

the calmed flow. The separation bubble remained thin, however.

The appearance of a disturbance did not immediately induce reattachment. The disturbance appeared initially in the outer part of

the boundary layer. After some lag time the disturbance spread

into the near wall region and caused reattachment.

5. While much has been learned regarding the effectiveness

and physics of synthetic VGJ’s in LPT flows, many questions

remain regarding their applicability under different flow conditions and their optimal design. These questions have been discussed.

Acknowledgments

This work was sponsored by the NASA Glenn Research Center.

The grant monitor is Dr. David Ashpis. Additional matching support was provided through a U.S. Naval Academy Recognition

Grant. Mr. Dale Boyer of the Technical Support Department at the

776 Õ Vol. 125, OCTOBER 2003 Transactions of the ASME

NASA/CR—2012-217656 48
Naval Academy fabricated the test section and the airfoils. Midshipman Dirk Heron conducted preliminary tests which led into

the present study.

Nomenclature

B 5 blowing ratio

Cf 5 skin friction coefficient

Cp 5 2(PT2P)/rUe

2

, pressure coefficient

cm 5 momentum coefficient

D 5 diameter

F1 5 (0.442Ls)/(UeT), dimensionless frequency

FSTI 5 free-stream turbulence intensity

H 5 d*/u, shape factor

K 5 (n/U`

2 )(dU` /ds), acceleration parameter

Ls 5 suction surface length

M 5 jet to boundary layer mass flux ratio

P 5 pressure

PT 5 upstream stagnation pressure

Re 5 UeLs /n, exit Reynolds number

Reu 5 momentum thickness Reynolds number

S 5 jet holes per unit span

s 5 streamwise coordinate, distance from leading edge

T 5 jet oscillation period

t 5 time

U 5 mean streamwise velocity

U` 5 local free-stream velocity

Ue 5 nominal exit free-stream velocity, based on inviscid

solution

Vjet 5 jet velocity

 ̄

u8 5 time averaged rms streamwise fluctuating velocity

 ̃

u8 5 ensemble averaged rms streamwise fluctuating velocity

2u8v8 5 time averaged turbulent shear stress

2 ̃u8 ̃

v8 5 ensemble averaged turbulent shear stress

y 5 cross-stream coordinate, distance from wall

d* 5 displacement thickness

lu 5 Reu

2 K, Thwaites parameter

n 5 kinematic viscosity

r 5 density

u 5 momentum thickness

Subscripts

ave 5 average over jet outpulse

jet 5 jet condition

max 5 maximum in jet cycle

` 5 free stream

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Journal of Turbomachinery OCTOBER 2003, Vol. 125 Õ 777

NASA/CR—2012-217656 49
Douglas G. Bohl

Ralph J. Volino

e-mail: volino@usna.edu

Department of Mechanical Engineering,

United States Naval Academy,

Annapolis, MD 21402

Experiments With ThreeDimensional Passive Flow

Control Devices on Low-Pressure

Turbine Airfoils

The effectiveness of three-dimensional passive devices for flow control on low pressure

turbine airfoils was investigated experimentally. A row of small cylinders was placed at

the pressure minimum on the suction side of a typical airfoil. Cases with Reynolds

numbers ranging from 25,000 to 300,000 (based on suction surface length and exit

velocity) were considered under low freestream turbulence conditions. Streamwise pressure profiles and velocity profiles near the trailing edge were documented. Without flow

control a separation bubble was present, and at the lower Reynolds numbers the bubble

did not close. Cylinders with two different heights and a wide range of spanwise spacings

were considered. Reattachment moved upstream as the cylinder height was increased or

the spacing was decreased. If the spanwise spacing was sufficiently small, the flow at the

trailing edge was essentially uniform across the span. The cylinder size and spacing

could be optimized to minimize losses at a given Reynolds number, but cylinders optimized for low Reynolds number conditions caused increased losses at high Reynolds

numbers. The effectiveness of two-dimensional bars had been studied previously under

the same flow conditions. The cylinders were not as effective for maintaining low losses

over a range of Reynolds numbers as the bars. DOI: 10.1115/1.2137743

Introduction

Boundary layer separation is a known problem on some modern

low-pressure turbine LPT airfoils, due to the strong adverse

pressure gradients created when designers impose higher loading

in an effort to improve efficiency and lower cost by reducing

airfoil count in engines. Separation bubbles, particularly those

which fail to close, can result in a significant loss of lift and a

subsequent degradation of engine efficiency e.g., Hourmouziadis

1, Mayle 2, and Sharma et al. 3. The problem is particularly

relevant in aircraft engines. Airfoils optimized to produce maximum power under takeoff conditions may still experience boundary layer separation at cruise conditions, due to the lower density

and therefore lower Reynolds numbers at altitude. A component

efficiency drop of 2% may occur between takeoff and cruise conditions in large commercial transport engines, and the difference

could be as large as 7% in smaller engines operating at higher

altitudes 4,5. Prediction and control of suction side separation,

without sacrifice of the benefits of higher loading, is therefore,

crucial for improved engine design.

Separation on airfoils is complicated by boundary layer transition. Separated flow transition in the LPT has been the focus of

several recent studies. Volino 6provides a review and describes

as follows the transition process on the suction side of a typical

LPT airfoil. The strong acceleration on the leading section of the

airfoil keeps the boundary layer thin and laminar, even in the

presence of elevated freestream turbulence. In most cases Volino

6observed that the boundary layer separated just downstream of

the suction peak. If transition then occurred in the shear layer over

the separation bubble, it caused the boundary layer to reattach.

A few recent studies have focused on control of transition and

reattachment in the LPT. Some have used active devices. Huang et

al. 7and Hultgren and Ashpis 8employed high voltage electrodes to produce glow discharge plasma in a boundary layer to

control separation. Bons et al. 4used steady and pulsed vortex

generator jets. Volino 9used oscillating vortex generator jets

with no net mass flow. Sieverding et al. 10used adjustable devices built into the suction surface.

While active flow control provides a means for adjusting to

changing flow conditions and in some cases the benefits of calmed

regions 4,9, passive flow control holds the advantage of simplicity. Van Treuren et al. 11, Lake et al. 12, Murawski and Vafai

13, Byerley et al. 14, Volino 15, Sieverding 10, Vera et al.

16, and Zhang and Hodson 17used various passive devices

under LPT conditions to control separation and in many cases

reduce losses. Most employed a relatively simple modification,

such as a small trip wire or bar essentially roughness, on the

suction surface of an airfoil.

Successful flow control results in a thin, attached boundary

layer at the trailing edge of an airfoil, thereby reducing losses. The

consensus of the studies listed above is that a device on the suction surface should be placed at or slightly downstream of the

pressure minimum. This is a logical result, since the effects of a

device farther upstream would be damped by the favorable pressure gradient, and a device too far downstream would lie under

the separation bubble and be ineffective. Volino 15used rectangular bars and found that the optimal bars were not large enough

to immediately trip the boundary layer to turbulent, but instead

allowed a small separation bubble to form. The bars introduced

small disturbances that grew and caused transition and reattachment to move upstream of their location in the uncontrolled case,

as explained in Volino and Bohl 18. The optimal bar height

depended on the flow conditions. As Re or freestream turbulence

is lowered, the separation bubble becomes larger, so a larger bar is

needed to produce enough of a disturbance to move transition

sufficiently far upstream. A flow control device producing too

small a disturbance will allow a larger separation bubble than

desired, resulting in a thicker boundary layer downstream of reattachment and higher losses. Similarly, too large a disturbance will

Contributed by the International Gas Turbine Institute IGTI of ASME for publication in the JOURNAL OF TURBOMACHINERY. Manuscript received October 1, 2004;

final manuscript received December 1, 2004. IGTI Review Chair: K. C. Hall. Paper

presented at the ASME Turbo Expo 2005: Land, Sea and Air, Reno, NV, June 6–9,

2005, Paper No. GT2005-68969.

Journal of Turbomachinery Copyright © 2006 by ASME APRIL 2006, Vol. 128 / 251

NASA/CR—2012-217656 50
move transition farther upstream than necessary, resulting in a

longer turbulent region and higher losses. Volino 15, Sieverding

et al. 10, and Zhang and Hodson 17all found that under steady

flow conditions, devices optimized for low Re tend to increase

losses at high Re. Devices optimized for high Re can be too small

to be effective at low Re. Unsteady wakes from upstream airfoils

promote transition and reattachment, and Zhang and Hodson 17

found that in unsteady flow optimal control was achieved using

smaller devices than in comparable steady flow cases. This made

it possible in unsteady flow to reduce losses with a single device

over a wider range of Re.

Passive flow control devices of various geometries have been

tested, but it is still uncertain if any particular device is superior.

In preliminary testing, Volino 15considered trip wires, rectangular bars, and delta wing vortex generators. All produced similar

results. Sieverding et al. 10found that straight trip wires were

somewhat better than rows of spherical roughness elements, but

only a limited number of cases were tested. Lake et al. 12found

dimples superior to other devices, presumably because the

dimples produced less blockage than devices that protruded into

the flow. Again, however, the number of cases considered was

limited, and more recent evidence 15,17suggests that optimal

devices should be quite small and produce minimal blockage even

if they do extend into the flow. Zhang and Hodson 17noted

differences in transition location with straight and “wavy” trip

wires and rectangular bars. The sharp backward facing step on a

bar, for example, produced an earlier transition than a round trip

wire of the same height. Still, this does not preclude that a bar and

a slightly larger wire could produce comparable results.

It is still possible that some devices might prove better for

reducing losses than others. Reynolds number can vary by an

order of magnitude during engine operation given the change in

ambient pressure between takeoff and cruise. Since passive devices by definition cannot be adjusted as conditions change, it is

highly desirable to use devices that reduce losses over as large a

Reynolds number range as possible. Given the potential payoff of

a more efficient engine, it is worthwhile to further consider passive flow control devices of different geometries.

A row of small vertical cylinders is considered in the present

study. The cylinders are located at the pressure minimum on the

suction surface and the spacing between cylinders is varied. The

geometry was chosen in the hope that in comparison to twodimensional bars or trips wires of the same height, isolated elements might produce a stronger disturbance due to the threedimensional nature of the flow around them, while presenting less

blockage due to the gaps between elements. The net result would

presumably be successful separation control with lower losses.

Experimental conditions match the low freestream turbulence

cases of Volino 15. Details of the experimental conditions and

the results are presented below.

Experiments

Experiments were conducted in a low speed wind tunnel, described by Volino et al. 19. Briefly, air enters through blowers

and passes through a honeycomb, a series of screens, two settling

chambers, and a three-dimensional contraction before entering the

test section. At the exit of the contraction, the mean velocity is

uniform to within 1%. The freestream turbulence intensity is

0.5% ± 0.05%. Nearly all of this freestream “turbulence” is actually streamwise unsteadiness at frequencies below 20 Hz and is

not associated with turbulent eddies. The root-mean-square intensities of the three components of the unsteadiness are 0.7%, 0.2%,

and 0.2% in the streamwise, pitchwise, and spanwise directions,

respectively.

The test section, shown in Fig. 1, follows the contraction and

consists of the passage between two airfoils. Details are listed in

Table 1 and more information is available in Volino 6. A large

span to chord ratio of 4.3 was chosen to insure two-dimensional

flow at the spanwise centerline of the airfoils, where all measurements were made. Upstream of each airfoil are flaps, which control the amount of bleed air allowed to escape from the passage.

The flaps, along with a tailboard on the pressure side of the passage, are adjusted to produce the correct leading edge flow and

pressure gradient along the airfoils. The flow in the passage

matches that in a multiblade cascade. The geometry of the passage

corresponds to that of the Pak-B airfoil, which is an industry

supplied research airfoil that is representative of a modern LPT

design. It has been used in several studies, as noted in Volino 6.

Experimental conditions match the smooth airfoil baseline

cases of Volino 6, who considered five Reynolds numbers Re

= 25, 000, 50,000, 100,000, 200,000, and 300,000. Reynolds

numbers are based on the exit velocity from the passage and the

suction surface length, Ls. Comparison data from cases with

passive-bar flow control are from Volino 15. The bars were of

uniform rectangular cross section and extended along the airfoil,

as shown in Fig. 2. The trailing edge of the bar was located at

s/Ls= 0.51, near the suction surface velocity peak. All bars were

6 mm wide in the streamwise direction. Bar heights of 0.4, 0.8,

and 1.6 mm were used. The bar heights were all less than 1% of

Ls. They compare to local boundary layer thickness at the bar

location of about 3.8, 2.7, 2.0, 1.4, and 1.2 mm in the base line

Re= 25, 000– 300, 000 cases, respectively.

A row of small vertical cylinders is used in the present experimental cases. The cylinders are D= 6 mm in diameter with one

end affixed to the suction surface and the other extending into the

flow. The cylinders are located in a line at the suction peak, in the

same location as the bars described above, as shown in Fig. 2.

Cylinder heights of 0.4 and 1.6 mm were considered, matching

the smaller and larger bar heights of Volino 15. Center to center

cylinder spacings ranging from a pitch, P, of 1D i.e., the cylinders were touching to 30D were considered.

Measurements. Pressure surveys were made for each case using a pressure transducer 0 – 870 Pa range Validyne transducer

and a Scanivalve. Stagnation pressure was measured with a pitot

tube upstream of the passage inlet, and 11 pressure taps were

located on each airfoil along their spanwise centerlines. The uncertainty in Cp is 7% at Re= 25, 000, and 4% in other cases. Most

of this uncertainty is due to bias error. Stochastic error was minimized by averaging pressure transducer readings over a 10 s period. The flow control cylinders were moved in the spanwise diFig. 1 Schematic of the test section

252 / Vol. 128, APRIL 2006 Transactions of the ASME

NASA/CR—2012-217656 51
rection allowing documentation with the pressure taps directly

downstream of the center of one cylinder and with the taps midway between the centers of two adjacent cylinders.

Profiles of the streamwise velocity component in the suction

side boundary layer were measured for the Re= 50, 000 and

300,000 cases near the trailing edge at s/Ls= 0.944. Profiles were

measured at several spanwise locations relative to the position of

the upstream cylinders. Data were acquired with a hot-wire anemometer AA Lab Systems model AN-1003 and a single sensor

boundary layer probe TSI model 1218-T1.5. The sensor diameter is 3.8 m, and the active length is 1.27 mm. At each measurement location, data were acquired for 26 s at a 20 kHz sampling rate 219 samples. Data were acquired at 60 wall normal

locations in each profile, extending from the wall to the

freestream, with most points concentrated in the near wall region.

The closest point was within 0.1 mm of the wall, which corresponds to y /Ls= 0.0004 and between 0.02 and 0.04 boundary layer

thicknesses. Uncertainties in the mean velocity are 3–5% except

in the very near wall region where near-wall corrections Wills

20 were applied to the mean velocity. Uncertainties in the momentum and displacement thicknesses computed from the mean

profiles are 10%. Uncertainty in the shape factor, H, is 8%. The

uncertainty in the fluctuating streamwise velocity is below 10%,

except in the very near wall region y1 mm, y /Ls0.004 of

the Re= 300, 000 cases, where spatial averaging effects, due to the

finite length of the hot-wire sensor, may become important. This is

discussed in Volino 6.

Results

Pressure Profiles. Suction side pressure profiles for the cases

with larger 1.6 mm high cylinders are shown in Figs. 3–5. Also

shown in each figure are the corresponding baseline results from

Volino 6, the 1.6-mm-thick bar case results from Volino 15,

and an inviscid solution for flow through the passage. Figure 3

shows results at Re= 25, 000 with various cylinder spacings and

the pressure taps directly downstream of one cylinder. The data

agree with the inviscid solution in the favorable pressure gradient

region. The near wall flow slows as it approaches a cylinder or

bar, causing a drop in the measured Cp below the inviscid solution at the fifth pressure tap. If the boundary layer separates and

does not reattach, there is a drop in the suction peak, as indicated

by low Cp values. In addition to this effect, if the blockage caused

by a bar or cylinder is sufficiently high, the streamlines immediately downstream will be displaced as they flow over the sixth

pressure tap, which may cause the measured local Cp to either rise

or fall relative to the inviscid solution, depending on the geometry

of a particular case. For controlling separation, cylinders with

spacing of P= 2D or more are ineffective at Re= 25, 000. There is

a plateau in Cp in the adverse pressure gradient region extending

to the trailing edge, indicating a separation bubble that does not

reattach. The P= 1D touching cylinders and the solid bar result

in a large separation bubble which appears to be starting to reattach by the trailing edge, as indicated by the drop in Cp back

toward the inviscid value at the last pressure tap. Figure 4 show

the results at Re= 50, 000 with the pressure taps directly downstream of one cylinder z/P= 0, Fig. 4a and with the taps downstream of the midpoint between adjacent cylinders z/P= 0.5, Fig.

4b. In the base line case, the boundary layer does not reattach.

Table 1 Test section parameters

Fig. 2 Scale drawing of suction side airfoil showing location

of „a... bar, or „b... cylinders Fig. 3 Pressure profiles, large cylinders, Re=25,000, z/P=0

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NASA/CR—2012-217656 52
With P= 1D or the solid bar, reattachment moves upstream to

s/Ls= 0.78. With P= 2D reattachment occurs at s/Ls= 0.86. With

P2D the effect of the cylinders is not observed until the last

pressure tap, but even with P= 30D there is some effect on Cp at

s/Ls= 0.94 and z/P= 0. As the spacing increases, the Cp values

increase from the inviscid solution toward the baseline case value.

At z/P= 0.5 and s/Ls= 0.94, Cp is affected for spacings up to

20D, although not as strongly as it is at z/P= 0. For P= 24D and

30D, the cylinders are too far apart to affect Cp at z/P= 0.5. The

results at Re= 100, 000 are shown in Fig. 5. In the base line case,

there is a clear separation bubble, and it reattaches near the trailing edge. With P3D, the separation is essentially eliminated.

With P= 5D there is a separation, but reattachment by s/Ls

= 0.78. For P5D, the Cp values are nearly uniform across the

span. For 10DP30D, the boundary layer appears to be at

least starting to reattach by s/Ls= 0.86, with the effect on Cp

greater at z/P= 0. Results for the Re= 200, 000 cases not shown

are similar to those described above. The boundary layer separates

in the base line case and reattaches by s/Ls= 0.86. Cylinders with

P16D effectively suppress the separation across the span. For

P16D, the boundary layer is attached at s/Ls= 0.78 and z/P

= 0, but is still separated at z/P= 0.5. At Re= 300, 000 there is only

a small separation bubble in the base line case, and it reattaches

by s/Ls= 0.78. Cylinders with spacing up to 30D appear to suppress this small bubble across the span.

Comparing the results at z/P= 0 and z/P= 0.5, the spanwise

influence of each cylinder can be estimated. If the region affected

by the cylinder is assumed to spread linearly in the spanwise

direction as the flow convects downstream, the half angle for the

spreading is between roughly 30 and 40 deg. Changes in Cp at

z/P= 0.5 could be due to local turbulence arising from the cylinder induced disturbances. It is also possible, however, that these

changes in Cp are due to changes in the mean flow resulting from

the cylinder induced changes at z/P=0.

The smaller 0.4 mm high cylinders and bar are ineffective at

Re= 25, 000 and 50,000. The boundary layer separates and does

not reattach. At Re= 100, 000, as shown in Fig. 6, the cylinders

cause the first indication of reattachment to move upstream from

near the trailing edge to s/Ls= 0.78 or 0.86. Cylinder spacings up

to 10D were considered, and the Cp values decrease from the base

line case values toward the inviscid solution as P is decreased. No

significant spanwise variation in Cp was observed. The results at

z/P= 0.5 are essentially the same as those shown in Fig. 6 for

z/P= 0. Results for the Re= 200, 000 cases are shown in Fig. 7.

The results are very similar to those of the Re= 100, 000 cases, but

with reattachment shifted about 0.08Ls upstream. When Re

= 300, 000 the base line separation bubble is small, and cylinders

effectively eliminate it for all value of P.

The presumably more complex three-dimensional flow around

isolated P1D cylinders did not produce disturbances better

Fig. 4 Pressure profiles, large cylinders, Re=50,000, „a... z/P

=0, „b... z/P=0.5

Fig. 5 Pressure profiles, large cylinders, Re=100,000, „a...

z/P=0, „b... z/P=0.5

254 / Vol. 128, APRIL 2006 Transactions of the ASME

NASA/CR—2012-217656 53
able to control separation than the touching P= 1D cylinders or

two-dimensional bars. In fact, the opposite was observed. The

pressure profiles show that as the spacing between cylinders increases, their ability to promote reattachment decreases. This is

not simply due to the finite spanwise influence of each cylinder, as

very little spanwise variation was observed for cases with P

10D. Whether increasing the spacing results in reduced flow

blockage and losses will be considered next as the velocity profile

results are considered.

Velocity Profiles. Mean and fluctuating streamwise velocity

profiles for the large cylinder Re= 50, 000 cases at z/P= 0 are

shown in Fig. 8. Data were acquired near the trailing edge s/Ls

= 0.94. The base line case shows a thick separation bubble in the

mean profile and a small peak in u in the shear layer over the

bubble. The two-dimensional bar case exhibits a fully attached

turbulent mean profile shape and a typical turbulent u profile

with a near wall peak. For the cases with cylinder spacing up to

P= 5D, the mean profiles appear attached and turbulent, but the

trend is toward a less full profile as P increases. This trend continues for the cases with P20D, with the mean profiles in these

cases appearing only partially reattached. The peaks in the u profiles increase in magnitude and move farther from the wall as

the cylinder spacing is increased. High peaks away from the wall

are typical of transitioning and reattaching boundary layers. Transition begins in the shear layer over the separation bubble, so u should be high at this location.The high upeaks result from the

switching between intermittently attached-turbulent-like and

separated-laminar-like states. Figure 9 shows the profiles at z/P

= 0.5. For P5D, the profiles appear essentially the same as those

at z/P= 0. For P20D the mean profiles show a separation

bubble, somewhat thinner than that of the base line case, and u profiles with only a small peak in the shear layer over the separation bubble. Figure 10 shows the velocity profiles at several spanwise positions for the P= 30D case. The progression from attached flow at z/P= 0 to separated flow at z/P= 0.5 is clear. The

profiles at z/P= 0.13 and 0.20 are on the edge of the attached flow

region with u peaks between the high peaks at low z/P and the

small shear layer peaks at higher z/P. The influence of the cylinders to a spanwise position between z/P= 0.13 and 0.20 suggests

a half angle for the spread of the cylinder influence of about

20 deg. The velocity profile data of Figs. 8–10 are consistent with

the conclusions drawn from the corresponding pressure coefficient

data of Fig. 4.

The information concerning separation in the mean profiles can

be presented in terms of the shape factor, H. The shape factor rises

to about 4 when a laminar boundary layer separates, and reaches

higher values as a separation bubble thickens. This is due to an

increase in the displacement thickness while the momentum thickness remains nearly constant. If the boundary layer reattaches, the

Fig. 6 Pressure profiles, small cylinders, Re=100,000, z/P=0

Fig. 7 Pressure profiles, small cylinders, Re=200,000, z/P=0

Fig. 8 Trailing edge velocity profiles, large cylinders, Re

=50,000, z/P=0, „a... U/Ue, „b... u/Ue

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displacement thickness drops and the momentum thickness begins

to rise. The shape factor reaches a turbulent value of about 1.4

after recovering from the separation. Figure 11 shows H as a

function of z/P for the Re= 50, 000 cases with the larger cylinders. In the base line case, H= 6.5, indicating a thick separation

bubble. With the two-dimensional bar, H= 1.6, indicating that the

boundary layer has reattached and is nearing fully developed turbulent conditions. With cylinder spacing up to P= 5D, H is spanwise uniform. With P= 5D, H is about 2.3, indicating that the

boundary layer has reattached but is not fully recovered from the

separation. For P20D, H3 at z/P= 0, indicating the boundary

layer has just begun to reattach at this location. For P20D and

higher z/P, H is between 4 and 5.5 indicating the boundary layer

is still separated. These values are still below the base line value

of 6.5, however, indicating that even the widely spaced cylinders

are effective in keeping theseparation bubble thinner.

Velocity profiles for the Re= 300, 000 cases with the larger bar

and cylinders are shown in Figs. 12–14. Figure 12 shows the

profiles at z/P= 0. The mean profiles show that the boundary layer

is attached in all cases. The u peaks are near the wall, as expected

for attached turbulent boundary layers. The base line case has the

highest u peak, indicating that it is the case with the least developed turbulence. The bar and cylinders move transition upstream,

resulting in a more developed turbulent boundary layer by the

trailing edge. Both the mean and u profiles collapse for the cases

with P20D. This suggests that the cylinders are far enough

apart in these cases so that they act as if isolated from each other.

Both the mean and u profiles show that the boundary layer becomes thicker as the cylinder spacing is decreased. The P= 1D

case has a second u peak away from the wall that is not present

Fig. 9 Trailing edge velocity profiles, large cylinders, Re

=50,000, z/P=0.5, „a... U/Ue, „b... u/Ue

Fig. 10 Trailing edge velocity profiles, large cylinders, Re

=50,000, P/D=30, „a... U/Ue, „b... u/Ue

Fig. 11 Shape factor, H, at trailing edge, large cylinders, Re

=50,000

256 / Vol. 128, APRIL 2006 Transactions of the ASME

NASA/CR—2012-217656 55
in the two-dimensional bar case. The touching cylinders appear to

generate more turbulence than the bar. Figure 13 shows the profiles at z/P= 0.5. The profiles for the P= 24D and 30D cases are

very similar to the baseline results, again suggesting that with

large enough spacing the cylinders are effectively isolated from

each other and cannot directly influence the flow across the entire

span. As P is reduced, the boundary layer becomes thicker, and u assumes a more turbulent like shape. Figure 14 shows the profiles

at several spanwise positions for the P= 30D case. The cylinders

appear to influence the boundary layer u nearly uniformly from

z/P= 0 to z/P= 20. For z/P27, the cylinders have little influence and the profiles are very similar to the base line case profiles.

The spanwise extent of the cylinder influence spreads at a half

angle of about 20 deg for a strong effect on the boundary layer

and about 30 deg for some effect.

Figure 15 shows the shape factor at s/Ls= 0.94 for the profiles

of the Re= 300, 000 cases with the larger cylinders. The boundary

layer is attached, so the shape factor is between 1.4 and 1.7 in all

cases. Figure 16 shows the momentum thickness for these cases.

The momentum thickness is related to losses in the boundary

layer, and in cases with equal shape factor and exit flow angle, the

momentum thickness is directly proportional to profile losses

Howell et al. 21. In the Re= 50, 000 cases, the large variation

in H Fig. 11 precludes a comparison of losses based on , but

with the smaller range of H at Re= 300, 000 Fig. 15 the comparison is appropriate. Figure 16 shows that the cylinders cause a

rise in momentum thickness above the base line case value and

that the effect increases as cylinder spacing is decreased. Since the

separation bubble is small even in the base line case, the cylinders

and bars add an unnecessary disturbance and increase losses. The

more widely spaced cylinders cause less blockage and create less

of a disturbance, therefore the losses are lower.

Velocity profiles for the Re= 50, 000 cases with the smaller bar

and cylinders are shown in Fig. 17. Cylinder spacings up to P

= 5D were considered. No variation was observed between the

results from different spanwise locations, so only the results from

z/P= 0 are shown. The boundary layer did not reattach for cases

with P2D, although the separation bubble was slightly thinner

and there was a slight increase in the u peak compared to the

base line case. With P= 1D the boundary layer appears to be on

the verge of reattachment, and with the two-dimensional bar the

boundary layer has just begun to reattach. The smaller cylinders

are inadequate for control of the boundary layer at this Reynolds

number, as previously indicated by the pressure profiles.

The effect of the smaller cylinders on the velocity profiles at

Re= 300, 000 are shown in Fig. 18. No spanwise variation was

observed, so only results from z/P= 0 are shown. The mean profile for the P/D= 1 case is noticeably different than those for the

other cases, including the two-dimensional 2D bar case. As was

noted above, the touching cylinders apparently generate more turbulence that the 2D bar, resulting in a thicker boundary layer. As

was the case with the larger cylinders Fig. 12, the smaller cylinders cause a drop in the u peak from the base line case value,

indicating that the cases with cylinders are closer to fully develFig. 12 Trailing edge velocity profiles, large cylinders, Re

=300,000, z/P=0, „a... U/Ue, „b... u/Ue

Fig. 13 Trailing edge velocity profiles, large cylinders, Re

=300,000, z/P=0.5, „a... U/Ue, „b... u/Ue

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oped turbulent behavior. Shape factors are shown in Fig. 19 as a

function of spanwise position. There is little variation between

cases, as expected since the separation bubble is small and the

boundary layer has fully reattached in all cases. The momentum

thickness is shown in Fig. 20. All of the cases with cylinders have

higher momentum thickness than the base line case. The case with

P= 1D has the highest values, again indicating that the touching

cylinders generate higher losses than the two-dimensional bars or

more widely spaced cylinders.

Fig. 15 Shape factor, H, at trailing edge, large cylinders, Re

=300,000

Fig. 14 Trailing edge velocity profiles, large cylinders, Re

=300,000, P/D=30, „a... U/Ue, „b... u/Ue

Fig. 16 Momentum thickness at trailing edge, large cylinders,

Re=300,000

Fig. 17 Trailing edge velocity profiles, small cylinders, Re

=50,000, z/P=0, „a... U/Ue, „b... u/Ue

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NASA/CR—2012-217656 57
Discussion

The results presented above suggest that the half angle for the

spreading of the disturbances from the cylinders is about 30 deg.

This is a rough estimate due to the finite spacing between pressure

tap locations and the finite number of cylinder spacings investigated, but the half angle roughly agrees with the spreading angle

for turbulent spots in an adverse pressure gradient, as given by

D’Ovidio et al. 22. This angle suggests that a cylinder spacing of

between P/D= 5 and P/D= 10 is necessary to insure flow control

across the span. The velocity profiles showed good spanwise uniformity for P/D10.

The optimal spacing for spanwise disturbances was investigated

in an analytical study by Tumin and Ashpis 23. They considered

disturbances in both favorable and adverse pressure gradient

flows, including the favorable pressure gradient found in the upstream region of the present test section. Although these results do

not apply directly to the adverse pressure gradient region of the

present study, they can be extrapolated to the present conditions.

They suggest an optimal spacing for maximum disturbance

growth in the Re= 50, 000 cases of P/Ls0.03. For the Re

= 300, 000 cases, the optimal spacing would be P/Ls0.012.

With the present cylinder diameter of 6 mm, these spacings correspond to P/D= 1 and P/D= 0.4, respectively, and could not be

achieved unless the cylinders were touching. Consistent with this,

the present results show that the disturbance created by the cylinders decreases as the spacing increases. To better test the Tumin

and Ashpis 23results, however, smaller diameter, separated cylinders with the recommended spacing should be considered.

The results at Re= 50, 000 show that the larger cylinders with

the closest spacing are needed to effectively control separation. At

Re= 300, 000, the separation bubble is small in the base line case,

and cylinders or bars cause transition to move upstream, increasing losses. Hence, the thinnest, most widely spaced which in the

limit means nonexistent devices are optimal. The cylinders which

are best at Re= 50, 000 cause significantly higher losses at Re

= 300, 000. This is the same result found by Volino 15using

two-dimensional bars, and agrees with other findings in the literature, as noted above. So long as the cylinders are close enough to

provide spanwise uniformity, it appears that varying the cylinder

or bar thickness and varying the cylinder spacing are both effective for controlling the transition location and moving it to an

optimal location for minimizing losses.

The present results can address the question of whether separated cylinders provide an advantage over a two-dimensional bar.

A case with a two-dimensional bar that is capable of controlling

separation while keeping losses to a minimum at a low Reynolds

number should be compared to a case with cylinders that are

thicker than this optimal bar but produce the same reattachment

and low losses. Cases with these same geometries should then be

compared at a high Reynolds number to see which results in lower

high-Re losses. The thin bar in the present study results in marFig. 18 Trailing edge velocity profiles, small cylinders, Re

=300,000, z/P=0, „a... U/Ue, „b... u/Ue

Fig. 19 Shape factor, H, at trailing edge, small cylinders, Re

=300,000

Fig. 20 Momentum thickness at trailing edge, small cylinders,

Re=300,000

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ginally reattached flow at Re= 50, 000. The larger cylinders result

in spanwise uniform reattachment at Re= 50, 000 when P/D=5.

The momentum thicknesses at s/Ls= 0.94 for these two cases are

within 8% of each other. At Re= 300, 000 the thin bar results in a

momentum thickness 7% above the base line case value, while the

thick, P/D= 5 cylinders result in a spanwise averaged momentum

thickness that is 46% above the baseline value. Clearly the thin

two-dimensional bar is superior to the thicker cylinders. This

comparison is not completely adequate, however, since the thin

bar only causes marginal reattachment at Re= 50, 000, while the

thick cylinders result in a somewhat more complete reattachment.

Volino 15also considered an intermediate bar with thickness

twice that of the thin bar considered above. At Re= 50, 000 the

intermediate bar induced complete reattachment with H= 1.8 and

/b= 1.23. These results are very close to those obtained with the

large cylinders spaced at P/D= 3. At Re= 300, 000 the intermediate bar resulted in H= 1.59 and /b= 1.29. The thick, P/D= 3

spaced cylinders resulted in an approximately equal shape factor,

but /b= 1.72. With this better comparison, it is still clear that a

thin bar is superior to larger cylinders.

The present results support the conclusions of Sieverding 10

and Zhang and Hodson 17who found that two-dimensional bars

or trips are as good or better than three-dimensional devices for

controlling separation. The present results cannot be considered

absolutely conclusive, however, since only a single geometry was

considered under a limited number of conditions. The number of

possible geometries and spacings for three dimensional devices is

infinite, so it will never be possible to prove conclusively through

experiments that two-dimensional devices are always better. Further study of devices such as the dimples considered by Lake et al.

12would be useful. The effects of high freestream turbulence

and unsteady wakes should also be considered. Perhaps thinner,

smaller diameter cylinders with spacings closer to those extrapolated from Tumin and Ashpis 23could provide better separation

control with lower losses. This is merely speculation. What can be

said is that the present results add to the evidence that simple

two-dimensional bars are preferable.

Conclusions

A row of small cylinders located at the suction peak on a LPT

airfoil were effective for separation control. The reattachment of

the boundary layer moves upstream as the cylinder height is increased or the cylinder spacing is decreased. The half angle of the

spreading of the disturbance created by the cylinders was of the

order 30 deg. This is roughly the same as the expected spreading

angle for a turbulent spot under the same adverse pressure gradient conditions. Based on this angle, the maximum allowable spacing for spanwise uniform separation control can be determined.

By varying the cylinder height and spacing, an optimal reattachment location can be achieved for minimum losses at a given

Reynolds number. Cylinders optimized for low Reynolds numbers

resulted in higher losses at high Reynolds numbers. The present

results add to the evidence that three-dimensional passive flow

control devices are not as effective as two-dimensional bars for

minimizing losses over a range of Reynolds numbers.

Acknowledgment

This work was sponsored by the NASA Glenn Research Center.

The grant monitor was Dr. David Ashpis. The first author received

matching support from the Office of Naval Research.

Nomenclature

Cp  2pT−p/Ue

2

, pressure coefficient

D  cylinder diameter

H  shape factor, 	* /

Ls  suction surface length

P  center to center spacing of cylinders

p  pressure

pT  upstream stagnation pressure

Re  UeLs/
, exit Reynolds number

s  streamwise coordinate, distance from leading edge

Ue  nominal exit freestream velocity, based on inviscid

solution

u  mean streamwise velocity

u  rms streamwise fluctuating velocity

y  cross-stream coordinate, distance from wall

z  spanwise coordinate

	*  displacement thickness


  kinematic viscosity

  density

  momentum thickness

b  momentum thickness in base line case

References

1Hourmouziadis, J., 1989, “Aerodynamic Design of Low Pressure Turbines,”

AGARD Lecture Series 167.

2Mayle, R. E., 1991, “The Role of Laminar-Turbulent Transition in Gas Turbine

Engines,” ASME J. Turbomach., 113, pp. 509–537.

3Sharma, O. P., Ni, R. H., and Tanrikut, S., 1994, “Unsteady Flow in Turbines,”

AGARD Lecture Series 195, Paper No. 5.

4Bons, J. P., Sondergaard, R., and Rivir, R. B., 2001, “Turbine Separation

Control Using Pulsed Vortex Generator Jets,” ASME J. Turbomach., 123, pp.

198–206.

5Volino, R. J., and Hultgren, L. S., 2001, “Measurements in Separated and

Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions,” ASME J. Turbomach., 123, pp. 189–197.

6Volino, R. J., 2002, “Separated Flow Transition Under Simulated LowPressure Turbine Airfoil Conditions: Part 1—Mean Flow and Turbulence Statistics,” ASME J. Turbomach., 124, pp. 645–655.

7Huang, J., Corke, T., and Thomas, F., 2003, “Plasma Actuators for Separation

Control on Low Pressure Turbine Blades,” AIAA Paper No. 2003-1027.

8Hultgren, L. S., and Ashpis, D. E., 2003, “Demonstration of Separation Delay

With Glow Discharge Plasma Actuators,” AIAA Paper No. 2003-1025.

9Volino, R. J., 2003, “Separation Control on Low-Pressure Turbine Airfoils

Using Synthetic Vortex Generator Jets,” ASME J. Turbomach., 125, pp. 765–

777.

10Sieverding, C. H., Bagnera, C., Boege, A. C., Antòn, J. A. C., and Luère, V.,

2004, “Investigation of the Effectiveness of Various Types of Boundary Layer

Transition Elements of Low Reynolds Number Turbine Bladings,” ASME Paper No. GT2004-54103.

11Van Treuren, K. W., Simon, T., von Koller, M., Byerley, A. R., Baughn, J. W.,

and Rivir, R., 2002, “Measurements in a Turbine Cascade Flow Under Ultra

Low Reynolds Number Conditions,” ASME J. Turbomach., 124, pp. 100–106.

12Lake, J. P., King, P. I., and Rivir, R. B., 2000, “Low Reynolds Number Loss

Reduction on Turbine Blades With Dimples and V-Grooves,” AIAA Paper No.

00-738.

13Murawski, C. G., and Vafai, K., 1999, “Effect of Variable Axial Chord on a

Low-Pressure Turbine Blade,” J. Propul. Power, 15, pp. 667–674.

14Byerley, A. R., Störmer, O, Baughn, J. W., Simon, T. W., VanTreuren, K. W.,

and List, J., 2002, “Using Gurney Flaps to Control Laminar Separation on

Linear Cascade Blades,” ASME J. Turbomach., 125, pp. 114–120.

15Volino, R. J., 2003, “Passive Flow Control on Low-Pressure Turbine Airfoils,”

ASME J. Turbomach., 125, pp. 754–764.

16Vera, M., Hodson, H. P., and Vazquez, R., 2004, “The Effects of a Trip Wire

and Unsteadiness on a High Speed Highly Loaded Low-Pressure Turbine

Blade,” ASME Paper No. GT2004-53822.

17Zhang, X. F., and Hodson, H., 2004, “Combined Effects of Surface Trips and

Unsteady Wakes on the Boundary Layer Development of an Ultra-High-Lift

LP Turbine Blade,” ASME J. Turbomach., 127, pp. 479–488.

18Volino, R. J., and Bohl, D. G., 2003, “Separated Flow Transition Mechanisms

and Prediction With High and Low Freestream Turbulence Under Low Pressure Turbine Conditions,” ASME Paper No. GT2004-63360.

19Volino, R. J., Schultz, M. P., and Pratt, C. M., 2001, “Conditional Sampling in

a Transitional Boundary Layer Under High Free-Stream Turbulence Conditions,” ASME J. Fluids Eng., 125, pp. 28–37.

20Wills, J. A. B., 1962, “The Correction of Hot-Wire Readings for Proximity to

a Solid Boundary,” J. Fluid Mech., 12, pp. 65–92.

21Howell, R. J., Ramesh, O. N., Hodson, H. P., Harvey, N. W., and Schulte, V.,

2001, “High Lift and Aft-Loaded Profiles for Low-Pressure Turbines,” ASME

J. Turbomach., 123, pp. 181–188.

22D’Ovidio, A., Harkins, J. A., and Gostelow, J. P., 2001, “Turbulent Spots in

Strong Adverse Pressure Gradients: Part 2—Spot Propagation and Spreading

Rates,” ASME Paper No. 2001-GT-0406.

23Tumin, A., and Ashpis, D. E., 2003, “Optimal Disturbances in Boundary Layers Subject to Streamwise Pressure Gradients,” AIAA J., 41, pp. 2297–2300.

260 / Vol. 128, APRIL 2006 Transactions of the ASME

NASA/CR—2012-217656 59
Y. B. Suzen

Assistant Professor

Department of Mechanical Engineering and

Applied Mechanics,

North Dakota State University,

Fargo, ND 58105

P. G. Huang

Professor and Chair

Mechanical and Materials Engineering

Department,

Wright State University,

Dayton, OH 45435

D. E. Ashpis

Aerospace Engineer

Mem. ASME

NASA Glenn Research Center at Lewis Field,

Cleveland, OH 44135

R. J. Volino

Associate Professor

Mem. ASME

Department of Mechanical Engineering,

United States Naval Academy,

Annapolis, MD 21402-5042

T. C. Corke

Clark Chair Professor

Fellow ASME

F. O. Thomas

Professor

Mem. ASME

J. Huang

Graduate Assistant

Department of Aerospace and Mechanical

Engineering,

Center for Flow Physics and Control,

University of Notre Dame,

Notre Dame, IN 46556

J. P. Lake

Special Projects Flight Commander,

586th FLTS/DON,

Holloman AFB, NM 88330

P. I. King

Professor

Mem. ASME

Department of Aeronautics and Astronautics,

Air Force Institute of Technology,

Wright-Patterson AFB, OH 45433

A Computational Fluid Dynamics

Study of Transitional Flows in

Low-Pressure Turbines Under a

Wide Range of Operating

Conditions

A transport equation for the intermittency factor is employed to predict the transitional

flows in low-pressure turbines. The intermittent behavior of the transitional flows is taken

into account and incorporated into computations by modifying the eddy viscosity, t, with

the intermittency factor, . Turbulent quantities are predicted by using Menter’s twoequation turbulence model (SST). The intermittency factor is obtained from a transport

equation model which can produce both the experimentally observed streamwise variation of intermittency and a realistic profile in the cross stream direction. The model had

been previously validated against low-pressure turbine experiments with success. In this

paper, the model is applied to predictions of three sets of recent low-pressure turbine

experiments on the Pack B blade to further validate its predicting capabilities under

various flow conditions. Comparisons of computational results with experimental data

are provided. Overall, good agreement between the experimental data and computational

results is obtained. The new model has been shown to have the capability of accurately

predicting transitional flows under a wide range of low-pressure turbine conditions.

DOI: 10.1115/1.2218888

1 Introduction

The process of transition from laminar to turbulent flow is a

major unsolved problem in fluid dynamics and aerodynamics. One

area where the transition process plays an important role and is

even more complicated due to the diverse flow conditions encountered is the low-pressure turbine applications. Transitional flows

in these applications are affected by several factors such as varying pressure gradients, wide range of Reynolds number and

freestream turbulence variations, flow separation, and unsteady

wake-boundary layer interactions. Accurate simulation and prediction of transitional flows under these diverse conditions is key

Contributed by the Turbomachinery Division of ASME for publication in the

JOURNAL OF TURBOMACHINERY. Manuscript received February 14, 2004; final manuscript received February 13, 2006. Review conducted by R. L. Davis.

Journal of Turbomachinery Copyright © 2007 by AIAA. Reprinted by permission. JULY 2007, Vol. 129 / 527

NASA/CR—2012-217656 60
to design of more efficient jet engines.

In low-pressure turbine applications, flow over the blades is

mostly turbulent at the high Reynolds number conditions encountered at takeoff and the efficiency is at its design maximum. However, at lower Reynolds number conditions which correspond to

high altitudes and cruise speeds the boundary layers on the airfoil

surface have a tendency to remain laminar; hence, the flow may

separate on the suction surface of the turbine blades before it

becomes turbulent. This laminar separation causes unpredicted

losses, substantial drops in efficiency, and increase in fuel consumption 1–3.

In order to calculate the losses and heat transfer on various

components of gas turbine engines, and to be able to improve

component efficiencies and reduce losses through better designs,

accurate prediction of development of transitional boundary layers

is essential 1.

One approach proven to be successful for modeling transitional

flows is to incorporate the concept of intermittency into computations. This can be done by multiplying the eddy viscosity obtained

from a turbulence model, t, used in the diffusive parts of the

mean flow equations, by the intermittency factor, Simon and

Stephens 4. This method can be easily incorporated into any

Reynolds averaged Navier-Stokes solver. In this approach, the intermittency factor, , can be obtained from an empirical relation

such as the correlation of Dhawan and Narasimha 5, or it can be

obtained from a transport model.

Dhawan and Narasimha 5correlated the experimental data

and proposed a generalized intermittency distribution function

across flow transition. Gostelow et al. 6extended this correlation

to flows with pressure gradients under the effects of a range of

freestream turbulence intensities. Solomon et al. 7, following the

work of Chen and Thyson 8, developed an improved method to

predict transitional flows involving changes in pressure gradients.

These empirical methods led to development of transport equations for intermittency.

Steelant and Dick 9proposed a transport equation for intermittency, in which the source term of the equation is developed

such that the distribution of Dhawan and Narasimha 5across

the transition region can be reproduced. Steelant and Dick used

their model, coupled with two sets of conditioned Navier-Stokes

equations, to predict transitional flows with zero, favorable, and

adverse pressure gradients. However, since their technique involved the solution of two sets of strongly coupled equations, the

method is not compatible with existing computational fluid dynamics CFD codes, in which only one set of Navier-Stokes

equations is involved. Moreover, the model was designed to provide a realistic streamwise behavior but with no consideration

of the variation of in the cross-stream direction.

Cho and Chung 10developed a k--turbulence model for

free shear flows. Their turbulence model explicitly incorporates

the intermittency effect into the conventional k- model equations

by introducing an additional transport equation for . They applied this model to compute a plane jet, a round jet, a plane far

wake, and a plane mixing layer with good agreement. Although

this method was not designed to reproduce flow transition, it provided a realistic profile of in the cross-stream direction.

Suzen and Huang 11developed an intermittency transport

equation combining the best properties of Steelant and Dick’s

model and Cho and Chung’s model. The model reproduces the

streamwise intermittency distribution of Dhawan and Narasimha

5and also produces a realistic variation of intermittency in the

cross-stream direction. This model has been validated against European Research Community On Flow Turbulence And Combustion ERCOFTAC benchmark T3-series experiments reported by

Savill 12,13, low-pressure turbine experiments of Simon et al.

14, and separated and transitional boundary layer experiments of

Hultgren and Volino 15with success 11,16–21.

In this paper we concentrate on prediction of three recent lowpressure turbine experiments on the Pratt and Whitney’s Pack B

blade under low Reynolds number conditions using the transport

model for intermittency. Due to the fact that the Pack B blade is

very sensitive to changes of flow conditions, it is an ideal test

blade for validating the transition/turbulence models. The three

sets of experiments considered are conducted by Lake et al.

3,22, Huang et al. 23, and Volino 24at three independent

facilities. These experiments provide an extensive database for

investigating transitional flows under low-pressure turbine conditions and are employed as benchmark cases for further testing of

the predicting capabilities of the current intermittency model. A

summary of the experiments are given in the next section. In Sec.

3, the intermittency transport model is presented and implementation of the model and the empirical correlations employed for the

onset of transition are described. In Sec. 4, the predictions of the

new intermittency model are compared against the experimental

data. Conclusions are provided in Sec. 5.

2 Low-Pressure Turbine Experiments

In this paper, we concentrate on computation of three sets of

low-pressure turbine experiments using the intermittency transport

model. These experiments are conducted by Lake et al. 3,22,

Huang et al. 23, and Volino 24. In these experiments Pratt and

Whitney’s Pack B blade is used; the details of the blade are shown

in Fig. 1. Overall, these experiments cover a Reynolds number

range from 10,000 to 172,000 and the freestream turbulence intensity ranges from 0.08% to 4%. The cases and data used for

comparison in this paper are summarized in Table 1. In the following sections details of these experimental efforts are given.

2.1 Pack B Blade Cascade Experiments of Lake et al.

[3,22]. Lake et al. 3,22conducted experiments on the Pack B

blade in order to identify methods for reducing separation losses

on low-pressure turbine blades under low Reynolds number conditions. In the experiments, they investigated flows at low Reynolds numbers of 43,000, 86,000, and 172,000 based on inlet

Fig. 1 P&W Pack B blade cascade details

528 / Vol. 129, JULY 2007 Transactions of the ASME

NASA/CR—2012-217656 61
velocity and axial chord and freestream turbulence intensities

FSTI of 1% and 4%. These conditions are similar to those encountered at high-altitude, low-speed flight of reconnaissance unmanned aerial vehicles used by USAF.

In Lake’s experiments, surface pressure coefficients, boundary

layer velocity, and turbulence profiles, total pressure loss data

were obtained at FSTI= 1% and FSTI= 4%. The test setup shown

in Fig. 2 included eight blades with axial chord of 0.1778 m

7 in., and blade spacing of 0.1575 m 6.2 in.. The blades were

numbered 1 through 8 starting from the inside bend. Boundary

layer measurements were taken on blade 5 and surface pressures

were measured around blades 4 and 6. In this paper, the Pack B

blade experiments with Reynolds numbers of 86,000 and 172,000

and freestream turbulence intensities of 1% and 4% are computed

and comparison of pressure distributions between experiments

and computations are performed.

Fig. 2 Experimental setup used by Lake et al. †3,22‡

Table 1 Details of the experiments used for comparison with computations

Source Test Section

Cx

m

Re

UinCx / FSTI %

Data used for

Comparison

Lake et al. 3,22P&W Pack B cascade 0.1778 86,000 1 & 4 Cp distribution

172,000 1 & 4 Cp distribution

Huang et al. 23P&W Pack B cascade 0.1595 10,000 0.08 Cp distribution

25,000 0.08 Cp distribution

50,000 0.08, 1.6, 2.85 Cp distribution,

velocity profilesa

75,000 0.08, 1.6, 2.85 Cp distribution,

velocity profilesa

100,000 0.08, 1.6, 2.85 Cp distribution,

velocity profilesa

Volino 24P&W Pack B single

passage

0.1537 10,291 0.5 Cp distribution,

velocity profiles

20,581 0.5 Cp distribution,

velocity profiles

41,162 0.5 Cp distribution,

velocity profiles

82,324 0.5 Cp distribution,

velocity profiles

a

Velocity profiles are available for FSTI= 0.08% and 2.85% from experiments.

Journal of Turbomachinery JULY 2007, Vol. 129 / 529

NASA/CR—2012-217656 62
2.2 Pack B Blade Cascade Experiments of Huang et al.

[23]. Huang et al. 23conducted experiments on Pack B blade

cascade for a range of Reynolds numbers and turbulence intensities. The Reynolds numbers range from 10,000 to 100,000 based

on inlet velocity and axial chord as listed in Table 1. In their

experiments the blades had an axial chord length of 0.1595 m

6.28 in.. The freestream turbulence intensity in the tunnel was

measured as 0.08%. In order to increase the turbulence intensity,

two grids with different mesh sizes were used. One of the grids

had the mesh size of 0.0254 m denoted as grid 0 and the other

had 0.008 m denoted as grid 3. The decay of turbulence after the

grids was measured using crosswire and they are shown in Figs. 3

and 4 along with the computed results for grid 0 and grid 3,

respectively. The grids were movable in the tunnel so that the

turbulence level of the flow that reaches the blades could be controlled by moving the grid that is, by increasing or decreasing the

distance between the grid and the blade. Experiments were performed for Reynolds numbers 50,000, 75,000, and 100,000, with

grids placed 0.762 m 30 in. away from the blade leading edge,

corresponding to turbulence intensities of 2.85% and 1.6% at the

leading edge for grid 0 and grid 3, respectively. For Re

= 100,000, grid 0 is placed at 0.5588 m 22 in. and 0.3556 m

14 in., corresponding to turbulence intensities of 3.62% and

5.2%, respectively. Pressure coefficient data are available for all

cases and detailed boundary layer measurements are available for

Re= 50,000, 75,000, and 100,000 with FSTI= 0.08% and 2.85%

cases. The cases and data used for comparisons in this paper are

listed in Table 1.

2.3 Pack B Experiments of Volino [24]. Volino 24investigated the boundary layer separation, transition, and reattachment

under low-pressure turbine airfoil conditions. The experiments included five different Reynolds numbers ranging between 10,291

and 123,492 and freestream turbulence intensities of 0.5% and

9%. The test section consisted of a single passage between two

Pack B blades as shown in Fig. 5. The axial chord length of the

blades was 0.1537 m 6.05 in.. There are flaps located upstream

of each blade to control the amount of bleed air allowed to escape

from the passage. These flaps were adjusted by matching measured pressure distribution for a high Reynolds number with the

inviscid pressure distribution on the blade. In addition to the upstream bleed flaps, a tailboard on the pressure side was used to set

the pressure gradient. The compiled data include pressure surveys,

mean and fluctuating velocity profiles, intermittency profiles, and

turbulent shear stress profiles. It was observed that the effect of

high Reynolds number or high freestream turbulence level was to

move transition upstream. Transition started in the shear layer

over the separation bubble and led to rapid boundary layer reattachment. At the lowest Re case, transition did not take place

before the trailing edge and the boundary layer did not reattach.

The beginning of transition corresponded to the beginning of a

significant rise in the turbulent shear stress. These experimental

results provide detailed documentation of the boundary layer and

extend the existing database to lower Reynolds numbers. The

cases used for comparisons with computations in this paper are

listed in Table 1 along with the type of data used for comparisons.

3 Intermittency Transport Model

In this section, the transport model for intermittency is presented. The model combines the transport equation models of

Steelant and Dick 9and Cho and Chung 10. Details of the

development and implementation of the transport model are given

in Suzen and Huang 11,16,17, and in Suzen et al. 18.

The model equation is given by

Fig. 3 Comparison of computed and experimental decay of

turbulence for experiments of Huang et al. †23‡, with grid 0

Fig. 4 Comparison of computed and experimental decay of

turbulence for experiments of Huang et al. †23‡, with grid 3

Fig. 5 Schematic of the test section for experiments of Volino

†24‡

530 / Vol. 129, JULY 2007 Transactions of the ASME

NASA/CR—2012-217656 63


t

+

uj

xj

= 1 − 



1 − F2C0ukukfsfs

+ F

C1

k ij

ui

xj

− C2

k3/2



ui

ukuk

1/2

ui

xj



xj

	

+ C3

k2





xj



xj

+



xj

1 − l

 + 1 − tt



xj

 1

The distributed breakdown function, fs has the form

fs = as

4 + bs3 + cs2 + ds + e

gs

3 + h 2

where s=s−st, and s is the distance along the streamline coordinate, and st is the transition location. The coefficients are

a = 50n

U

b = − 0.4906

c = 0.204

n

U 

−0.5

d = 0.0 e = 0.04444

n

U 

−1.5

h = 10e g = 50 3

The shear stresses are defined as

ij = t


ui

xj

+

uj

xi

− 2

3

uk

xk

ij − 2

3

kij 4

The blending function F is constructed using a nondimensional

parameter k/W, where k is the turbulent kinetic energy and W is

the magnitude of the vorticity. The blending function has the form

F = tanh4


 k/W

2001 − 0.1

0.3  5

The model constants used in Eq. 1 are

l = t = 1.0 C0 = 1.0 C1 = 1.6

C2 = 0.16 C3 = 0.15

The intermittency is incorporated into the computations simply by

multiplying the eddy viscosity obtained from a turbulence model,

t, by the intermittency factor, . Simon and Stephens 4showed

that, by combining the two sets of conditioned Navier-Stokes

equations and making the assumption that the Reynolds stresses in

the nonturbulent part are negligible, the intermittency can be incorporated into the computations by using the eddy viscosity, t

*

which is obtained by multiplying the eddy viscosity from a turbulence model, t, with the intermittency factor, . That is

t

* = t 6

is used in the mean flow equations. It must be noted that does

not appear in the generation term of the turbulent kinetic energy

equations.

Computations of the experiments are performed using a recently developed multiblock Navier-Stokes solver, called GHOST.

The code was developed at the University of Kentucky, by Huang,

and is a pressure-based code based on the SIMPLE algorithm with

second-order accuracy in both time and space. Advection terms

are approximated by a QUICK scheme and central differencing is

used for the viscous terms. The “Rhie and Chow” momentum

interpolation method 25is employed to avoid checkerboard oscillations usually associated with the nonstaggered grid arrangement. This code is capable of handling complex geometries, moving, and overset grids and includes multiprocessor computation

capability using message passing interface MPI. Since multiple

processors are used during the computations, it is more efficient to

divide the computational domain into several smaller pieces with

very fine grids and distribute the zones to processors with the

consideration of load balancing. This code has been used extensively in a recent turbulence model validation effort Hsu et al.

26 and computations of unsteady wake/blade interaction Suzen

and Huang 27 conducted at the University of Kentucky.

The multiblock grid systems used in the computations are obtained by conducting a series of grid refinement studies in order to

ensure that the details of the flow field are captured accurately and

the results are grid independent. All grid systems have first y+ less

than 0.5 near solid walls.

In using this intermittency approach, the turbulence model selected to obtain t must produce fully turbulent features before

transition location in order to allow the intermittency to have full

control of the transitional behavior. Menter’s 28SST model satisfies this requirement. It produces almost fully turbulent flow in

the leading edge of the boundary layer and therefore is used as a

baseline model to compute t and other turbulent quantities in the

computations 18.

The value of n used in evaluating the constants given by Eq.

3 is provided by the following correlation for zero-pressure gradient flows 18

nˆ = nv2

/U3 = 1.8 	 10−11Tu7/4 7

When flows are subject to pressure gradients, the following correlation is used

nˆ

nˆZPG

= 

M1−exp0.75	106Kt

Tu−0.7, Kt 
 0

10−3227Kt

0.5985

, Kt  0 
 8

with M defined as

M = 850Tu−3 − 100Tu−0.5 + 120

where nˆZPG is the value for flow at zero pressure gradient and

can be obtained from Eq. 7, and Kt=/Ut

2

dU/dxt is the flow

acceleration parameter. The favorable pressure gradient part of the

above correlation for Kt0 is from Steelant and Dick 9. The

portion of the correlation for adverse pressure gradient flows for

Kt
0 is formulated using the transition data of Gostelow et al. 6

and Simon et al. 14Suzen et al. 18.

Fig. 6 Multiblock grid used for computations of experiments

of Lake et al. †3,22‡ and FSTI=0.08% experiments of Huang et

al. †23‡

Journal of Turbomachinery JULY 2007, Vol. 129 / 531

NASA/CR—2012-217656 64
The current approach uses the intermittency transport model to

obtain the intermittency distribution for the transitional flows,

while the onset of transition is defined by correlations.

The onset of attached flow transition is determined by the following correlation in terms of turbulence intensity, Tu, and the

acceleration parameter, Kt,

Ret = 120 + 150Tu−2/3coth40.3 − Kt 	 105

9

where Kt was chosen as the maximum absolute value of that parameter in the downstream deceleration region 18. This correlation maintains the good features of Abu-Ghannam and Shaw 29

correlation in the adverse pressure gradient region, and in addition

reflects the fact that the flow becomes less likely to have transition

when subject to favorable pressure gradients by rapidly rising as

Kt becomes positive.

In order to determine the onset of separated flow transition Rest

is expressed in terms of the turbulence intensity Tu and the

momentum thickness Reynolds number at the point of separation

Res in the form 19

Rest = 874Res

0.71 exp− 0.4Tu10

This correlation provides a better representation of the experimental data than Davis et al. 30correlation and is used to predict

onset of separated flow transition in the present computations.

4 Results and Discussion

4.1 Simulations of Experiments of Lake et al. [3,22]. The

intermittency model is applied to predict the Pack B blade experiments of Lake et al. 3,22. In the computations, flows at Reynolds numbers of 86,000 and 172,000 based on inlet velocity and

axial chord with freestream intensities of 1% and 4% were

investigated.

The computations were performed using the grid system shown

in Fig. 6 consisting of five zones obtained as a result of a grid

refinement study. In the grid refinement study computations were

performed on a series of successively finer grids and the variations

in the results were observed. The grid shown in Fig. 6 was chosen

to be adequate for obtaining grid-independent solutions for all

cases. The four zones on which the blade grid is superposed each

have 125	225 grid points and the O-type grid around the blade

has 401	101 points with first y+ less than 0.5.

The comparisons of computed and experimental pressure coefficient distributions are shown in Figs. 7a–7d. In these figures,

the experimental distributions correspond to the measurements

made on test blades 4 and 6.

The computed results compare well with the experiments for

high turbulence intensity, FSTI= 4%, cases shown in Figs. 7a

and 7c. However, for FSTI= 1% cases shown in Figs. 7b and

Fig. 7 Comparison of computed pressure coefficient with experiments of Lake et al. †3,22‡

532 / Vol. 129, JULY 2007 Transactions of the ASME

NASA/CR—2012-217656 65
7d, the extent of the separation bubbles is underpredicted in the

computations. For example, for Re= 86,000, FSTI= 1%, shown in

Fig. 7d, the flow reattaches earlier in computations than it does

in the experiment, as can be observed from the difference in the

pressure coefficient distributions between x/Cx= 80 to 85%.

The comparison of computed total pressure loss coefficients

with experiments is shown in Fig. 8. For the Re= 86,000 case, the

computed loss coefficient is in good agreement with the experiments for both FSTI levels. However, for the Re= 172,000 case

the computations underpredicted the loss coefficient compared to

experiments for both FSTI= 1% and FSTI= 4%. From Fig. 8 it is

evident that the cascade losses decrease as the Reynolds number

increases. This reduction in cascade losses with increasing Reynolds number is due to the decrease in size of the separated flow

region on the suction side of the blades.

The onset of separation locations, reattachment locations, and

onset of transition locations on the suction surface are summarized in Table 2 for these cases, along with the corresponding

values from experiments. In the experiments, the onset of transition locations and the reattachment locations are not reported. The

experimental onset of separation and reattachment points are extracted from the experimental pressure coefficient data. The onset

of separation is taken to be the axial location where the plateau in

the pressure coefficient distribution of the suction side begins, and

the reattachment point is taken to be the axial location after the

sharp change in Cp following the plateau. This procedure may

lead to an error of approximately ±1.5% of axial chord in the

estimated onset locations.

The onset of separation, reattachment, and onset of transition

locations are plotted against Reynolds number in Figs. 9a and

9b for FSTI= 4% and 1%, respectively. The uncertainty in the

estimated experimental values is indicated by error bars in the

figures. For the high turbulence intensity case, computation predicts onset of separation and reattachment slightly upstream of the

experiment. For the low FSTI case shown in Fig. 9b, the separation zone is predicted smaller than the experiments. The onset of

transition is predicted over the separated flow region in the shear

layer. From comparison of these figures it is evident that, with

decreasing freestream turbulence intensity, the separation zone becomes larger, and for a given FSTI condition, the separated flow

region gets smaller with increasing Reynolds number.

4.2 Simulations of Experiments of Huang et al. [23]. In this

set of experiments, first the cases with no grid in tunnel corresponding to FSTI= 0.08% are computed. In these computations,

the same grid system used for the computations of experiments of

Lake et al. 3,22shown in Fig. 6 is used.

The comparisons of the computed and the experimental pressure coefficients are shown in Figs. 10a–10e for Re= 100,000,

75,000, 50,000, 25,000, and 10,000 based on inlet velocity and

axial chord. The agreement between the experiments and computations is very good for all cases.

The computed total pressure loss coefficients are compared to

the available data for Re= 25,000 and 50,000 in Fig. 8. The loss

coefficients predicted in the computations are 2% to 3% higher

compared to the experiments for both Reynolds numbers.

The onset of separation, transition, and reattachment locations

are tabulated in Table 3 for all cases and plotted against Reynolds

number in Figs. 11a–11c for FSTI= 0.08%, 1.6%, and 2.85%,

Fig. 8 Comparison of computed total pressure loss coefficients with experiments of Lake et al. †3,22‡ and Huang et al.

†23‡

Table 2 Separation, reattachment, and transition locations for cases of Lake et al. †3,22‡

Re

UinCx /

FSTI

%

xs/Cx

Computation

xs/Cx

Experiment

xr/Cx

Computation

xr/Cx

Experiment

xtr/Cx

Computation

172,000 4 0.732 0.74 0.83 0.84 0.806

86,000 4 0.725 0.74 0.86 0.88 0.832

172,000 1 0.728 0.72 0.82 0.84 0.808

86,000 1 0.722 0.72 0.87 0.90 0.849

Fig. 9 Comparison of separation, reattachment, and transition

locations for experiments of Lake et al. †3,22‡

Journal of Turbomachinery JULY 2007, Vol. 129 / 533

NASA/CR—2012-217656 66
respectively.

Computed velocity profiles at seven axial stations along the

suction surface of the blade are compared to the experiments for

Re= 100,000, 75,000, and 50,000 in Figs. 12–14, respectively.

For the Re= 100,000 case, the computed velocity profiles compare very well with the experiment as shown in Figs. 12a–12g.

At the first three measurement stations, flow is laminar and attached as shown in Figs. 12a–12c. Flow separation takes place

at x/Cx= 0.725 and the separated flow region is visible in Figs.

12d and 12e, corresponding to axial locations of x/Cx= 0.75

and 0.80. The flow transition and reattachment takes place around

x/Cx= 0.84 in the computation. Reattachment location is earlier

than the experiment which takes place at x/Cx= 0.875. In Fig.

12f corresponding to axial station of x/Cx= 0.85 the computed

flow field has already attached, although the experimental profile

indicates a very small separation zone close to wall. At x/Cx

= 0.9 the flow is completely attached as shown in Fig. 12g.

When the Reynolds number is reduced to 75,000, the size of the

separation bubble increases as can be observed from the comparison of the velocity profiles shown in Figs. 13a–13g. At this

Reynolds number the flow separates around x/Cx0.72 and reattaches around x/Cx0.87. The transition onset location is predicted at x/Cx= 0.854. The size of the separation bubble is larger

than the Re= 100,000 case from comparison of Figs. 13d–13f

and 12d–12f.

Next, the Reynolds number is reduced to 50,000 and the comparison of computed and experimental velocity profiles is shown

in Figs. 14a–14g. For this case the separation bubble is much

larger from the previous cases and extends until x/Cx0.975 in

the experiment and x/Cx0.93 in the computations, as can be

seen in Figs. 14d–14g. Computations predicted the transition

onset location at x/Cx= 0.89. In the computations, the onset of

separation is predicted well in agreement with experiment; however, the reattachment point is earlier, making the size of the separation bubble smaller when compared to experiment. This is evident from the comparison of velocity profiles at the last two

stations shown in Figs. 14f and 14g.

The onset of separation and reattachment points for FSTI

= 0.08% cases is predicted upstream of the experiments as shown

in Fig. 11a.

Next, the high FSTI cases are computed using the six zone

multiblock grid system shown in Fig. 15. The computational domain is extended upstream of the blade in order to specify the

correct turbulence intensity at the inlet and to match the decay of

Fig. 10 Comparison of computed pressure coefficients with experiments of Huang et al. †23‡ for FSTI=0.08% cases

Table 3 Separation, reattachment, and transition locations for cases of Huang et al. †23‡

Re

UinCx /

FSTI

%

xs/Cx

Computation

xs/Cx

Experiment

xr/Cx

Computation

xr/Cx

Experiment

xtr/Cx

Computation

10,000 0.08 0.661 0.725  ̄ ̄ ̄ 25,000 0.08 0.656 0.725 0.980  ̄ 0.936

50,000 0.08 0.714 0.725 0.925 0.975 0.890

75,000 0.08 0.718 0.725 0.860 0.870 0.854

100,000 0.08 0.725 0.725 0.840 0.875 0.840

50,000 1.6 0.722 0.728 0.900 0.900 0.854

75,000 1.6 0.728 0.730 0.867 0.875 0.834

100,000 1.6 0.732 0.730 0.860 0.877 0.821

50,000 2.85 0.728 0.722 0.887 0.900 0.837

75,000 2.85 0.732 0.729 0.840 0.870 0.816

100,000 2.85 0.735 0.734 0.842 0.850 0.806

534 / Vol. 129, JULY 2007 Transactions of the ASME

NASA/CR—2012-217656 67
turbulence that reaches the blade. The matched computed and experimental turbulence decays are shown in Figs. 3 and 4 for grid

0 and grid 3, respectively. The cases considered have the grids

placed 0.762 m 30 in upstream of the blade, corresponding to

turbulence intensities of 2.85% and 1.6% for grid 0 and grid 3,

respectively.

The comparison of the computed and the experimental pressure

coefficient distributions for Re= 50,000, 75,000, and 100,000 for

FSTI= 2.85% cases is shown in Fig. 16. The agreement is very

good between computations and experiments.

Comparisons of computed velocity profiles with the experiments for Re= 100,000 are given in Figs. 17a–17g. In this

case, the flow separates around x/Cx0.74 and reattaches at

x/Cx0.85. The onset of transition is predicted at x/Cx= 0.806.

The computed size and extent of the separation bubble is in good

agreement with the experiment as tabulated in Table 3 and as can

be seen in Figs. 11c and 16d–16f.

For the lower Reynolds number of 75,000, computed velocity

profiles are compared with the experiments in Figs. 18a–18g.

The agreement between experiment and computation is good prior

to the reattachment as shown in Figs. 18a–18e. There is a

discrepancy in the reattachment region. The flow separation takes

place around x/Cx0.73 and reattaches at x/Cx0.87 according

to the experiment, whereas computation predicts reattachment earlier at around x/Cx0.84 with the onset of transition predicted at

x/Cx= 0.816. The difference in reattachment points is evident in

the comparison of the computed and experimental velocity profiles shown in Fig. 18f. At this station the experimental profile

indicates separated flow and the computed profile shows an already attached flow.

The next case considered has the same FSTI= 2.85% but with

Reynolds number being reduced to 50,000. The comparison of

velocity profiles is shown in Figs. 19a–19g. The computations

agree well with the experiment, and the size and extent of the

separation bubble are well predicted as can be seen from Fig.

11c. The onset of separation is around x/Cx0.72 and the flow

reattaches around x/Cx0.9, with transition onset at x/Cx

= 0.837.

In Fig. 20, computed and experimental pressure coefficient distributions for grid 3 case which correspond to FSTI= 1.6% are

compared for Re= 50,000, 75,000, and 100,000. Again, very good

agreement between computations and experiments is obtained.

The onset of separation and reattachment locations shown in Fig.

11b compares well with the experiments.

Overall, Figs. 11a–11c indicate that, as FSTI increases, the

separated flow region decreases, and at a given FSTI, increasing

Reynolds number has the same effect on the separated flow region.

4.3 Simulations of Pack B Experiments of Volino [24]. In

computation of experiments of Volino 24the flow field is modeled with the 31-zone multiblock grid shown in Fig. 21 obtained

as a result of a series of grid refinement studies. The bleed flaps

below the lower blade and above the upper blade are defined by

fitting third-order polynomials through the available points obtained from experimental setup; these curves are used as the flap

shapes in generating the computational grid. Initial computations

indicated that the shape of the bleed flaps and the orientation of

the tailboard behind the upper blade greatly affect the computed

results, especially the onset of separation and reattachment points

on the lower blade’s suction surface. In order to select the most

accurate orientation for the tailboard and the shape of the bleed

flaps, several test computations were performed for the case with

Re= 41,162 and FSTI= 0.5% using different tailboard orientations

and bleed flap shapes. In these computations the main goal was to

match the experimental velocity profiles in the laminar flow part

and to capture the correct onset point of separation. Once an acceptable geometry is obtained, the final bleed flap shapes and

tailboard orientation are used for computation of all other Reynolds number cases.

Computed pressure coefficient distributions are compared to

experiments in Figs. 22a–22d for Re= 82,324, 41,162, 20,581,

and 10,291, and the separation onset, reattachment, and transition

onset information is summarized in Table 4. The Cp comparison

for Re= 82, 324 shown in Fig. 22a indicates that the computation

predicts early reattachment of the flow; in the recovery region

following reattachment the pressure coefficient distribution is

overpredicted.

The computed pressure coefficient distributions for the lower

Reynolds number cases shown in Figs. 22c and 22d compare

well with experiments. For the Re= 41,162 case shown in Fig.

22b, the onset of separation and reattachment locations matches

the experiment as given in Table 4; however, in the recovery reFig. 11 Comparison of separation, reattachment, and transition locations for experiments of Huang et al. †23‡

Journal of Turbomachinery JULY 2007, Vol. 129 / 535

NASA/CR—2012-217656 68
gion the pressure coefficient distribution is overpredicted.

Computed velocity profiles are compared to experiment at 11

stations along the suction surface of the blade in Figs. 23a–23k

for Re= 82, 324 and FSTI= 0.5%. The results compare well with

the experiment up to x/Cx= 0.732 shown in Figs. 23a–23g.

After this station flow separation takes place. Separation onset and

reattachment are slightly earlier in the computations compared to

experiment as given in Table 4. This also can be observed from

the velocity profiles at stations x/Cx= 0.798 to 0.912 shown in

Figs. 23h–23j. Overall computations compare well with the

experimental measurements.

Next the Reynolds number is reduced to 41,162 and the computed and experimental velocity profiles are compared in Figs.

24a–24k. The computed profiles agree well with experiments

except at x/Cx= 0.912 shown in Fig. 24j. At this station the

computation indicates a smaller separated flow region close to

reattachment in contrast to the experiment. However, the flow

reattaches around x/Cx= 0.95 both in computation and experiment, and in the next measurement station the agreement is well.

The next case considered has a Reynolds number of 20,581.

Computed velocity profiles are shown along with the experimental

data at 11 axial stations in Figs. 25a–25k. In this case flow

separates around x/Cx0.76 and does not reattach in experiment;

however, computations indicated reattachment at x/Cx0.98.

This discrepancy is evident from the comparison of velocity proFig. 12 Comparison of computed velocity profiles with experiments of Huang et al. †23‡, Re=100,000, FSTI=0.08% case

Fig. 13 Comparison of computed velocity profiles with experiments of Huang et al. †23‡, Re=75,000, FSTI=0.08% case

Fig. 14 Comparison of computed velocity profiles with experiments of Huang et al. †23‡, Re=50,000, FSTI=0.08% case

Fig. 15 Grid used for computation of experiments of Huang et

al. †23‡ with FSTI=1.6% and 2.85%

536 / Vol. 129, JULY 2007 Transactions of the ASME

NASA/CR—2012-217656 69
Fig. 16 Comparison of computed pressure coefficients with experiments of Huang et al. †23‡ for FSTI=2.85% cases

Fig. 17 Comparison of computed velocity profiles with experiments of Huang et al. †23‡, Re=100,000, FSTI=2.85% case

Fig. 18 Comparison of computed velocity profiles with experiments of Huang et al. †23‡, Re=75,000, FSTI=2.85% case

Fig. 19 Comparison of computed velocity profiles with experiments of Huang et al. †23‡, Re=50,000, FSTI=2.85% case

Fig. 20 Comparison of computed pressure coefficients with experiments of Huang et al. †23‡ for FSTI=1.6% cases

Journal of Turbomachinery JULY 2007, Vol. 129 / 537

NASA/CR—2012-217656 70
files at the last two measurement stations shown in Figs. 25j and

25k. The computation indicates a smaller separated region in

these stations and finally reattaches very close to the trailing edge.

Onset of transition was predicted at x/Cx= 0.978.

The final case in this set of experiments is the one with Re

= 10,291. The computed velocity profiles compare very well with

the experimental data as shown in Figs. 26a–26k. In this case

the flow separates around x/Cx0.76 and does not reattach. The

flow is completely laminar; transition was not predicted on the

blade.

5 Concluding Remarks

A transport equation for the intermittency factor is employed to

predict three sets of recent low-pressure turbine experiments on

the Pack B blade. The intermittent behavior of the transitional

flows is taken into account by modifying the eddy viscosity with

the intermittency factor. Comparisons of the computed and experimental data are made and overall good agreement with the experimental data is obtained. The predicting capabilities of the current

intermittency approach and the intermittency transport model in

prediction of transitional flows under a wide range of lowpressure turbine conditions is demonstrated.

Fig. 21 Thirty-one zone multiblock grid used for computation

of experiments of Volino †24‡

Fig. 22 Comparison of computed pressure coefficient distributions with experiments of Volino †24‡, FSTI=0.5%

538 / Vol. 129, JULY 2007 Transactions of the ASME

NASA/CR—2012-217656 71
Table 4 Separation, reattachment, and transition locations for cases of Volino †24‡

Re

UinCx /

FSTI

%

xs/Cx

Computation

xs/Cx

Experiment

xr/Cx

Computation

xr/Cx

Experiment

xtr/Cx

Computation

10,291 0.5 0.760 0.750  ̄ ̄ ̄

20,581 0.5 0.765 0.760 0.980  ̄ 0.978

41,162 0.5 0.760 0.770 0.950 0.950 0.840

82,324 0.5 0.757 0.767 0.890 0.900 0.857

Fig. 23 Comparison of computed velocity profiles with experiments of Volino †24‡, Re=82,324, FSTI=0.5%

Fig. 24 Comparison of computed velocity profiles with experiments of Volino †24‡, Re=41,162, FSTI=0.5%

Journal of Turbomachinery JULY 2007, Vol. 129 / 539

NASA/CR—2012-217656 72
Acknowledgment

This work is supported by NASA Glenn Research Center under

Cooperative Agreement NCC3-590 and followed by NCC3-1040.

The project is part of the Low Pressure Turbine Flow Physics

Program of NASA-Glenn.

The experimental efforts at University of Notre Dame are supported by Cooperative Agreement NCC3-935 and Cooperative

Agreement NCC3-775 and the research at U.S. Naval Academy is

supported by Contract C-31011-K. This paper was originally published as AIAA Paper 2003-3591.

Nomenclature

Cp 
 pressure coefficient, 2P−P/Uin

2 

Cx 
 axial chord

FSTI 
 freestream turbulence intensity %

Kt 
 flow acceleration parameter /Ue

2

dUe / ds

k 
 turbulent kinetic energy

Lx 
 axial chord

N 
 nondimensional spot breakdown rate parameter, nt

3 /

Fig. 25 Comparison of computed velocity profiles with experiments of Volino †24‡, Re=20,581, FSTI=0.5%

Fig. 26 Comparison of computed velocity profiles with experiments of Volino †24‡, Re=10,291, FSTI=0.5%

540 / Vol. 129, JULY 2007 Transactions of the ASME

NASA/CR—2012-217656 73
n 
 spot generation rate

P 
 static pressure

Ptotal 
 total pressure

Re 
 Reynolds number

Rest 
 st−ssUe /

Ret 
 tUe /

s 
 streamwise distance along suction surface

Tu 
 turbulence intensity %, u/U

U 
 boundary layer streamwise velocity

Ue 
 local freestream velocity

Uin 
 inlet freestream velocity

u 
 friction velocity

W 
 magnitude of vorticity

yn 
 distance normal to the wall

y+ 
 ynu/


 intermittency factor

 
 momentum thickness

 
 pressure gradient parameter 2 /dU/ ds

 
 molecular viscosity

t 
 eddy viscosity

 
 /

t 
 t /

 
 total pressure loss coefficient,

2Ptotalinlet−Ptotalexit/Uin

2 

 
 density

 
 spot propagation parameter

Subscripts

e 
 freestream

s 
 onset of separation

t 
 onset of transition

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Journal of Turbomachinery JULY 2007, Vol. 129 / 541

NASA/CR—2012-217656 74
1 Copyright © 2002 by ASME 

Proceedings of 

ASME TURBO EXPO 2002 

June 3-6, 2002, Amsterdam, The Netherlands

GT-2002-30236 

SEPARATED FLOW TRANSITION UNDER SIMULATED LOW-PRESSURE TURBINE 

AIRFOIL CONDITIONS: PART 1 – MEAN FLOW AND TURBULENCE STATISTICS

Ralph J. Volino

Department of Mechanical Engineering 

United States Naval Academy 

Annapolis, Maryland 21402 

Email: volino@usna.edu

ABSTRACT 

 Boundary layer separation, transition and reattachment have been 

studied experimentally under low-pressure turbine airfoil conditions. 

Cases with Reynolds numbers (Re) ranging from 25,000 to 300,000 

(based on suction surface length and exit velocity) have been 

considered at low (0.5%) and high (9% inlet) free-stream turbulence 

levels. Mean and fluctuating velocity and intermittency profiles are 

presented for streamwise locations all along the airfoil, and turbulent 

shear stress profiles are provided for the downstream region where 

separation and transition occur. Higher Re or free-stream turbulence 

level moves transition upstream. Transition is initiated in the shear 

layer over the separation bubble and leads to rapid boundary layer 

reattachment. At the lowest Re, transition did not occur before the 

trailing edge, and the boundary layer did not reattach. Turbulent shear 

stress levels can remain low in spite of high free-stream turbulence and 

high fluctuating streamwise velocity in the shear layer. The beginning 

of a significant rise in the turbulent shear stress signals the beginning 

of transition. A slight rise in the turbulent shear stress near the trailing 

edge was noted even in those cases which did not undergo transition or 

reattachment. The present results provide detailed documentation of 

the boundary layer and extend the existing database to lower Re. The 

present results also serve as a baseline for an investigation of 

turbulence spectra in Part 2 of the present paper, and for ongoing work 

involving transition and separation control. 

NOMENCLATURE 

Cf skin friction coefficient 

Cp ( ) 2

T − ρU/PP2 e , pressure coefficient 

FSTI free-stream turbulence intensity 

H δ

∗

/θ, shape factor 

K ) /( )( / 2 ν U dU∞∞ dx , acceleration parameter 

Ls suction surface length 

 hot-wire sensor length 

P pressure 

PT upstream stagnation pressure

Re LU se /ν , exit Reynolds number 

Res U∞s/ν, local Reynolds number

Reθ momentum thickness Reynolds number 

s streamwise coordinate, distance from leading edge 

U mean streamwise velocity 

U∞ local free-stream velocity 

Ue nominal exit free-stream velocity, based on inviscid solution 

u′ rms streamwise fluctuating velocity 

uτ τ w / ρ , friction velocity 

− vu ′′ time averaged turbulent shear stress 

v′ rms wall normal fluctuating velocity 

y cross-stream coordinate, distance from wall 

y

+ yuτ/ν, distance from wall in wall coordinates 

δ

∗ displacement thickness 

γ intermittency, fraction of time flow is turbulent 

ν kinematic viscosity 

ρ density 

τw wall shear stress 

θ momentum thickness 

INTRODUCTION 

 Modern low-pressure (LP) turbine airfoils are subject to 

increasingly stronger pressure gradients as designers impose higher 

loading in an effort to improve efficiency and lower cost by reducing 

the number of airfoils in an engine. If the adverse pressure gradient on 

the suction side of these airfoils becomes strong enough, the boundary 

layer will separate. Separation bubbles, particularly those which fail 

to reattach, can result in a significant loss of lift and a subsequent 

degradation of engine efficiency (e.g. Hourmouziadis [1], Mayle [2], 

and Sharma et al. [3]). The problem is particularly relevant in aircraft 

engines. Airfoils optimized to produce maximum power under takeoff 

conditions may still experience boundary layer separation at cruise 

conditions, due to the thinner air and lower Reynolds numbers at 

altitude. A component efficiency drop of 2 percent may occur 

between takeoff and cruise conditions in large commercial transport 

engines, and the difference could be as large as 7 percent in smaller 

engines operating at higher altitudes. Component life may also be 

NASA/CR—2012-217656 75
2 Copyright © 2002 by ASME 

affected by more than an order of magnitude (Hodson [4]). Because 

the LP turbine produces the bulk of the net power in many engines, 

changes in its component efficiency can result in nearly equal changes 

in overall engine efficiency (Wisler [5]). There are several sources for 

losses in an engine, including secondary flows, but the suction side 

boundary layer has been identified as the primary source of losses in 

the LP turbine (Curtis et al. [6]). Prediction and control of suction side 

separation, without sacrifice of the benefits of higher loading, is 

therefore, necessary for improved engine design. 

 Separation on LP turbine airfoils is complicated by boundary layer 

transition. Turbulent boundary layers are much more resistant to 

separation than laminar boundary layers. A substantial fraction of the 

boundary layer on both sides of a turbine airfoil may be transitional 

(Mayle [2]), so accurately predicting transition location is crucial for 

accurate prediction of separation. Transition prediction for turbine 

airfoils is complex and can depend on a number of factors, including 

the free-stream turbulence intensity (FSTI), streamwise pressure 

gradient, airfoil curvature, surface roughness, and the unsteadiness 

associated with passing wakes from upstream stages. Several 

transition mechanisms are possible under engine conditions. Mayle 

[2] classified the modes of transition as “natural transition” involving 

Tollmien-Schlichting waves; “bypass” transition caused by high freestream turbulence or other large disturbances; “separated flow” 

transition of the shear layer over a separation bubble; “periodicunsteady” transition, which might also be called wake-induced 

transition; and reverse transition. If transition occurs far enough 

upstream, it can prevent separation. If transition occurs in the shear 

layer over a separation bubble, it will tend to induce boundary layer 

reattachment. The lower the Reynolds number, the farther 

downstream transition will tend to occur, hence the problems 

associated with performance at altitude. 

 Boundary layer transition has been studied extensively, and in 

recent years several studies have focused on transition in the LP 

turbine. Halstead et al. [7] present a study from a rotating cascade 

with multiple stages and FSTI characteristic of engine conditions. The 

adverse pressure gradients in this study were not strong enough to 

induce separation, however. Solomon [8] subsequently modified the 

facility and provides documentation of separation from a more 

aggressive airfoil. Gier and Ardey [9] provide another example from a 

rotating facility. Boundary layers and separation bubbles on flat plates 

subject to adverse pressure gradients have been considered in several 

studies. Recent work has included the studies of Hatman and Wang 

[10], Sohn et al. [11], Lou and Hourmouziadis [12], Volino and 

Hultgren [13] and Yaras [14]. On airfoils, either in cascade or singlepassage cascade-simulators, studies have included Murawski et al. 

[15], Qiu and Simon [16] and Simon et al. [17]. Cascades with 

moving wakes, simulating the effect of upstream blade rows, were 

utilized by Brunner et al. [18], Stadtmüller et al. [19], Howell et al. 

[20] and Kaszeta et al. [21]. Numerical studies have included the 

work of Dorney et al. [22], Chernobrovkin and Lakshminarayana [23], 

Huang and Xiong [24], and Thermann et al. [25]. A few studies have 

included attempts to control transition and separation. Howell et al. 

[20] studied modified airfoil shapes, Van Treuren et al. [26] utilized 

vortex generators, and Lake et al. [27] considered various passive 

devices including dimples. Bons et al. [28] showed considerable 

success using both steady and pulsed vortex generator jets. The 

preceding list of studies, while long, is by no means inclusive. It is 

merely a sample of recent work, biased toward the most recent studies. 

 Much has been learned from the work to date, but the nature of 

separated flow transition is still not completely clear, and existing 

models are still not as robust as needed for accurate prediction. The 

present study expands the existing database. The flow through a 

single-passage cascade-simulator is documented under both high and 

low FSTI conditions at several different Reynolds numbers. The 

geometry of the passage corresponds to that of the “Pak-B” airfoil, 

which is an industry supplied research airfoil that is representative of a 

modern, aggressive LP turbine design. This geometry was used in 

several of the studies mentioned above ([15, 16, 17, 21, 22, 23, 27, 

28]), and the pressure profile from the suction side of this airfoil was 

matched in the flat plate study of Volino and Hultgren [13]. Previous 

work has included documentation at Re (based on suction surface 

length and exit free-stream velocity) as low as 50,000 (e.g. [13, 16, 

21]). The present work includes the first complete documentation (to 

the author’s knowledge) of cases with Re as low as 25,000. Also new 

is documentation of the turbulent shear stress in the boundary layer 

under both high and low FSTI. 

 Details of the experimental facility and results of the study follow. 

The present paper focuses on mean and statistical quantities. Part 2 of 

this work [29] includes turbulence spectral results, providing evidence 

of the important transition mechanisms. 

EXPERIMENTS

 Experiments were conducted in a low speed wind tunnel, described 

by Volino et al. [30]. Briefly, air enters through blowers and passes 

through a series of screens, a honeycomb, two settling chambers, and a 

three-dimensional contraction before entering the test section. At the 

exit of the contraction, the mean velocity is uniform to within 1%. 

The FSTI is 0.5%0.05%. Nearly all of this free-stream “turbulence”

is actually streamwise unsteadiness at frequencies below 20 Hz and is 

not associated with turbulent eddies. The rms intensities of the three 

components of the unsteadiness are 0.7%, 0.2% and 0.2% in the 

streamwise, pitchwise and spanwise directions, respectively. For low 

FSTI cases, the test section immediately follows the contraction. For 

high FSTI, a passive grid is installed at the contraction exit followed 

by a 1 m long rectangular settling chamber. Details of the grid are 

available in Volino et al. [30]. At the inlet to the test section the high 

FSTI mean flow and turbulence are spatially uniform to within 3% and 

6% respectively. The free-stream turbulence is nearly isotropic with 

rms intensities of 8.8%, 8.9% and 8.3% in the streamise, pitchwise and 

spanwise directions. The integral length scales of these components 

are 3 cm, 1.6 cm and 1.4 cm. The integral scales were computed from 

the power spectra of each component. 

 The test section, shown in Fig. 1, consists of the passage between 

two airfoils. Details are listed in Table 1. Cascade simulators of this 

type have been used in studies such as Chung and Simon [31], more 

recently in the present facility by Aunapu et al. [32], and with the PakB geometry by Qiu and Simon [16] and Kaszeta et al. [21]. A large 

span to chord ratio of 4.3 was chosen to insure two-dimensional flow 

at the spanwise centerline of the airfoils, where all measurements were 

made. Upstream of each airfoil are flaps, which control the amount of 

bleed air allowed to escape from the passage. These are adjusted to 

produce the correct leading edge flow and pressure gradient along the 

airfoils. A tailboard on the pressure side of the passage also aids in 

setting the pressure gradient. 

 Single passage test sections have several advantages. For a given 

wind tunnel with fixed maximum flow rate, the single passage can be 

considerably larger than a passage in a multi-blade facility. The larger 

size and simpler geometry can also result in better probe access. 

Previous studies (e.g. [31, 32]) demonstrated that the full flow field, 

including the three-dimensional secondary flows near the endwalls, in 

a single passage can be set to match that in a corresponding multiblade cascade. The present test section also has some advantages over 

flat plate test sections. First, the airfoil curvature is matched. Second, 

with an adverse pressure gradient, suction is often needed to prevent 

NASA/CR—2012-217656 76
3 Copyright © 2002 by ASME 

Table 1: Test section parameters 

Axial 

Chord 

[mm] 

True 

Chord 

[mm] 

Pitch 

[mm] 

Span 

[mm] 

Suction 

side, Ls

[mm] 

Inlet 

flow 

angle 

Exit 

flow 

angle 

153.6 170.4 136.0 660.4 228.6 35° 60°

separation on the wall opposite a flat test plate (e.g. Volino and 

Hultgren [13]). A cascade simulator does not require suction due to 

the favorable pressure gradient on the pressure side of the passage. 

 Single passages also have disadvantages. It is, of course, impossible 

to establish periodicity. Stage losses cannot be directly determined 

since there is flow only on one side of each airfoil and the downstream 

wake is, therefore, unlike that in a multi-blade facility. This limitation, 

however, is not prohibitive for the present study. Primary concern is 

with boundary layer separation and transition, which occur in the 

passage. Although the downstream wake may be different, the flow in 

the passage does match that of a multi-blade facility. 

 Ten different cases have been documented including high and low 

FSTI cases at five Reynolds numbers (Re=25,000, 50,000, 100,000, 

200,000, and 300,000). The Reynolds number range is representative 

of conditions from cruise to takeoff. The FSTI levels in an engine may 

vary considerably, but the values in the present work are believed to 

span the range of most interest. Solomon [8] surveyed several studies 

that included wake effects and found FSTI values ranging from 1 to 

5% between wakes and from 3 to 23% within wakes. Wakes can 

affect transition and separation in the boundary layer in three ways. 

First, the FSTI rises during a wake passage compared to the betweenwake value. Second, a calmed region follows wake induced transition. 

The calmed flow is non-turbulent, but unlike a steady non-turbulent 

flow, it can be very resistant to separation. Finally, independent of the 

FSTI effect, each wake includes a mean velocity deficit, resulting in 

temporal deceleration and acceleration as the wake passes. Lou and 

Hourmouziadis [12] separated this temporal effect from the wake 

turbulence effect, using downstream control to create an oscillating 

velocity in their test section. In the present study, only steady flow is 

considered. While the significant effects of temporal acceleration and 

calming are not present, the high and low FSTI cases of the present 

study do allow a means for evaluating the effect of wake turbulence 

level. 

Measurements 

 Pressure surveys were made for each case using a pressure 

transducer (0-870 Pa range Validyne transducer) and a Scanivalve. 

Stagnation pressure was measured with a pitot tube upstream of the 

passage inlet, and eleven pressure taps were located on each airfoil 

along their spanwise centerlines. Locations of the taps on the suction 

side are listed in Table 2 along with measured local FSTI components, 

and the ReK product at these stations based on a non-separating, 

inviscid solution. The pressure distribution on the upstream portion of 

the suction side always closely matched the inviscid solution for flow 

over the airfoil. This allowed the use of the measured static pressure 

at the third pressure tap on the suction side, along with the inviscid 

flow solution for the passage and the upstream stagnation pressure, to 

determine the nominal passage exit velocity, which was used to 

normalize the measured pressure distributions. More conventionally 

the measured inlet velocity and the inlet and exit flow angles are used 

to compute the exit velocity. Because the velocity at the third tap is 

1.9 times that at the passage inlet, and therefore easier to measure, 

using the third tap velocity reduced the bias uncertainty in the pressure 

coefficients, particularly at the lower Reynolds numbers. The 

uncertainty in the suction side pressure coefficients was 7% at the 

lowest Re, and below 4% in other cases. Most of this uncertainty was 

tailboard

flap

flap bleed air

bleed air

35o

60o

Fig. 1: Schematic of the test section 

due to bias error. Stochastic error was minimized by averaging 

pressure transducer readings over a 10 second period. 

 Velocity profiles were measured at eleven streamwise stations along 

the suction side at the locations given in Table 2. Profiles were 

measured near but not at the spanwise centerline of the airfoil to insure 

that the pressure taps did not interfere with the velocity measurements. 

Profiles were acquired with a hot-wire anemometer (TSI model 

IFA100) and a single sensor boundary layer probe (TSI model 1218-

T1.5). The sensor diameter is 3.8 μm, and the active length is 1.27 

mm. At each measurement location, data were acquired for 26 

seconds at a 20 kHz sampling rate (219 samples). All raw data were 

saved. The high sampling rate provides an essentially continuous 

signal, which is needed for intermittency and spectral post-processing. 

The long sampling time results in low uncertainty in both statistical 

and spectral quantities. Data were acquired at 60 wall normal 

locations in each profile, extending from the wall to the free-stream, 

with most points concentrated in the near wall region. The closest 

point was 0.1 mm from the wall, which corresponds to y/Ls=0.0004 

and between 0.01 and 0.2 boundary layer thicknesses. Flow direction 

in a separation bubble cannot be determined with a single-sensor hotwire, but velocity magnitude can be measured and was found to be 

essentially zero within the bubbles of the present cases. Determining 

the direction was not, therefore, considered essential. At locations 

where the boundary layer was attached, local wall shear stress was 

computed from the near wall profile using the technique of Volino and 

Simon [33]. Uncertainties in the mean velocity are 3-5% except in the 

very near wall region (y

+

<5) where near-wall corrections (Wills [34]) 

were applied to the mean velocity. Uncertainties in the momentum 

and displacement thicknesses computed from the mean profiles are 

10%. Uncertainty in the shape factor, H, and the wall shear stress are 

both 8%. 

 The uncertainty in the fluctuating streamwise velocity is below 

10%, except in the very near wall region, where spatial averaging 

effects become important in some cases. Ligrani and Bradshaw [35, 

36] showed that spatial averaging over the length of a hot-wire sensor 

can result in low apparent u′ . Their experiments were done in a fullyNASA/CR—2012-217656 77
4 Copyright © 2002 by ASME 

Table 2: Measurement stations locations, local acceleration 

(inviscid soln.), and measured local free-stream turbulence 

Station 

s/Ls ReK 

×103

Low 

FSTI

′/ Uu ∞

[%] 

Low 

FSTI

′/ Uv ∞

[%] 

High 

FSTI

′/ Uu ∞

[%] 

High 

FSTI

′/ Uv ∞

[%] 

1 0.111 209.3 0.44 5.2 

2 0.194 121.4 0.39 4.6 

3 0.278 77.6 0.37 4.0 

4 0.361 58.2 0.38 3.5 

5 0.444 37.2 0.39 3.2 

6 0.528 -0.2 0.41 2.8 

7 0.611 -60.6 0.47 0.05 2.9 5.9 

8 0.694 -71.7 0.47 0.12 3.0 6.2 

9 0.777 -57.3 0.48 0.14 3.4 6.6 

10 0.861 -48.7 0.54 0.11 3.8 6.8 

11 0.944 -17.5 0.51 0.11 4.0 6.8 

turbulent boundary layer with Reθ=2600. The spatial averaging effects 

become important when the sensor length is longer than the width of 

the smaller near wall streaks in a turbulent boundary layer. Ligrani 

and Bradshaw [35, 36] found that the spatial averaging effects become 

small when the dimensionless sensor length,  τ /u ν , is less than 

about 25. The error also becomes smaller as the sensor is moved away 

from the wall. This is expected since the average size of the turbulent 

eddies should increase with distance from the wall. The Ligrani and 

Bradshaw [35, 36] results suggest that when the distance from the 

wall, y, is larger than the sensor length,  , that spatial averaging errors 

are under 10% even for large  τ /u ν . Closer to the wall they showed 

errors in u′ as large as 30% when  τ /u ν =60. 

 In the present study,  τ /u ν remains below 25 in all cases with 

Re<200,000. Spatial averaging is not, therefore, expected to be a 

problem, even near the wall. For the Re=200,000 cases,  τ /u ν is 

above 25 at Station 11 of the low FSTI case and at Stations 9-11 of the 

high FSTI case, reaching values as high as 60. Spatial averaging 

should not be significant for y>1 mm (y/Ls>0.004), but may cause 

errors as high as 30% closer to the wall. It is not certain that the errors 

are this large, however. The momentum thickness Reynolds numbers 

in the present cases are all below 700, which is significantly below the 

Reθ=2600 value of the Ligrani and Bradshaw [35, 36] study. This may 

indicate less developed turbulence in the present study, which could 

imply fewer small scale eddies and lower averaging errors. For the 

Re=300,000 cases,  τ /u ν reaches values as high as 90 at Stations 10 

and 11 of the low FSTI case and Stations 9-11 of the high FSTI case. 

As in the Re=200,000 cases, errors should be small when y>1 mm, but 

may be larger closer to the wall. 

 A boundary layer cross-wire probe (TSI model 1243-T1.5) was used 

to measure profiles of the wall normal velocity and turbulent shear 

stress at Stations 7-11 for each case. The upstream boundary layer 

was too thin for cross-wire measurements. Data were acquired at 25 

locations in each profile, beginning 1 mm from the wall and extending 

to the free-stream. Sampling rates and times were the same as for the 

single sensor probe. Uncertainty in the turbulent shear stress is 10%. 

 Measurements with the cross-wire probe are subject to spatial 

averaging errors due to the length of the sensors (1.27 mm active 

length) and the spacing between the two sensors (1 mm). Applying 

the results of Ligrani and Bradshaw [35, 36] and Ligrani et al. [37] to 

the present cases, spatial averaging may be significant at locations 

very near the wall, particularly for the high Reynolds number cases. 

At locations farther from the wall than 1 mm, however, the errors 

should become small and within the 10% uncertainty estimate given 

above. Hence, no measurements were made at y locations below 1 

mm. Because all measurements were at y>1 mm and a boundary layer 

type probe was used, probe blockage effects were not expected to be 

significant. 

 The intermittency, γ, is the fraction of time the flow is turbulent 

within the transition region, and was determined at each measurement 

location using the technique described in Volino et al. [30] with an 

uncertainty of 10%. Turbulent flow in the boundary layer is defined 

here as flow which includes a range of large and small scales, 

turbulence production, and dissipation. Using this definition, a 

boundary layer may be characterized by significant fluctuations but 

still be non-turbulent if these fluctuations are induced by an external 

source which does not cause near wall turbulence production. Such is 

often the case under high FSTI conditions. Free-stream eddies 

“buffet” the boundary layer, inducing non-turbulent boundary layer 

fluctuations. Buffeting may occur through pressure fluctuations. 

Boundary layer fluid is pushed in the wall normal direction across the 

mean gradient in the streamwise velocity, resulting in significant u′

fluctuations. This type of motion was termed “inactive” by Bradshaw 

[38] since it does not result in momentum transport, in spite of 

potentially high u′ levels. In addition to buffeting, some free-stream 

eddies may penetrate into the boundary layer and cause some mixing. 

As described in Volino [39], the eddies which have the greatest effect, 

whether through buffeting or penetrating the boundary layer, will be 

the larger, energy containing eddies. The non-turbulent boundary 

layer subject to these external effects will be characterized by large 

amplitude, low frequency fluctuations. Transition to turbulence is 

characterized not so much by large increases in u′ levels, which may 

remain essentially constant, but by the appearance of higher 

frequencies superimposed on the low frequencies. The higher 

frequencies signal the generation of turbulence in the near wall region. 

Volino et al. [30] provide examples of an intermittent flow switching 

in time between disturbed non-turbulent and turbulent states. 

 The presence or absence of high frequencies in a signal is used to 

distinguish between turbulent and non-turbulent flow, using the 

algorithm presented by Volino et al. [30]. The algorithm is similar to 

others found in the literature. Briefly, the time derivative of a signal is 

computed and compared to a threshold. Rapid (high frequency) 

fluctuations result in high derivatives. When the derivative is larger 

than the threshold, the flow is declared instantaneously turbulent and 

the intermittency function is assigned a value of 1. When the 

derivative is below the threshold, the intermittency is assigned a value 

of 0. The time average of the function is the intermittency, γ. Volino 

et al. [30] showed that intermittency can be computed based on u′ or 

vu ′′ signals with essentially the same result. Results based on u′ are 

presented in this paper. 

RESULTS 

Pressure Profiles

 Pressure coefficients for all ten cases are shown in Fig. 2. At the 

upstream stations on the suction side, there is good agreement between 

the data at all Reynolds numbers and the inviscid flow solution. 

Separation appears to occur at s/Ls of about 0.6 in all cases. These 

results agree with those of Volino and Hultgren [13], who also 

observed that the separation location did not depend strongly on the 

Reynolds number or FSTI. They are in contrast to other studies, such 

as Qiu and Simon [16], which showed that the separation location 

depended more strongly on Re. Reattachment depends strongly on 

NASA/CR—2012-217656 78
5 Copyright © 2002 by ASME 

0 0.2 0.4 0.6 0.8 1 1.2 1.4 0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

Re=25000

s/Ls

Cp

Re=50000

Re=100000

Re=200000

Re=300000

Re=400000

(a) LOW FSTI

inviscid solution

0 0.2 0.4 0.6 0.8 1 1.2 1.4 0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

Re=25000

s/Ls

Cp

Re=50000

Re=100000

Re=200000

Re=300000

Re=400000

(b) HIGH FSTI

inviscid solution

Fig. 2: Cp profiles: (a) low FSTI, (b) high FSTI 

both Re and FSTI. With low FSTI (Fig. 2a), the boundary layer 

appears to be separated in all cases at Stations 8 and 9. It reattaches by 

Station 10 for the Re=200,000 and 300,000 cases, reattaches by Station 

11 for the Re=100,000 case, and does not reattach at all for the 

Re=25,000 and 50,000 cases. Reattachment for the high FSTI cases 

(Fig. 2b) occurs upstream of the low FSTI locations. The boundary 

layer appears to be separated in all cases at Station 8, but has already 

reattached by Station 9 in the Re=200,000 and 300,000 cases. 

Reattachment has occurred by Station 10 for the Re=100,000 case, and 

appears to be beginning at Station 11 for the Re=50,000 case. The 

boundary layer does not appear to reattach when Re=25,000, in spite 

of the high FSTI. Results for the low FSTI cases are very similar to 

the flat plate results of Volino and Hultgren [13]. The present high 

FSTI results appear to show about a 10% larger separation region than 

the flat plate cases of [13]. The differences are small and of the order 

of the resolution of the measurement stations. Any differences 

between the studies are presumably due to differences in the freestream turbulence. Although the high FSTI in both studies was about 

8%, the inlet free-stream turbulence was more anisotropic in Volino 

and Hultgren [13], and the integral length scales of the free-stream 

turbulence in [13] were about double those in the present study. The 

larger length scale presumably caused earlier transition in the Volino 

and Hultgren [13] study, resulting in a slightly shorter separation 

bubble. Comparison to the high FSTI cases of Simon et al. [17] shows 

reattachment about 14% farther upstream in [17] than in the present 

study. With low FSTI, Simon et al. [17] did not observe reattachment 

at all when Re=100,000, while it was observed in the present study. 

For the low FSTI Re=200,000 case, they indicate reattachment about 

6% farther upstream than the present study. Although the streamwise 

pressure gradients were nominally the same in the present study and 

Simon et al. [17], small differences in the pressure gradients along 

with differences in the intensity and length scales of the free-stream 

turbulence were apparently responsible for the differences in 

reattachment location. 

Upstream Boundary Layer 

 The local free-stream velocities at Station 1-5 for all 10 cases 

closely followed the equation 

.0 214

e Ls

s .1 48

U

U















 = ∞ (1) 

which corresponds to the free-stream velocity distribution for a 

Falkner-Skan wedge flow. Figure 3 shows that the 50 mean velocity 

profiles from all 10 cases at these stations collapse onto the same 

Falkner-Skan profile. Skin friction coefficients, shown in Fig. 4, 

which were computed using the near wall profiles, also follow the 

Falkner-Skan solution. There is no significant difference between the 

low and high FSTI cases. Agreement with the flat plate data of Volino 

and Hultgren [13] is good. For the low FSTI cases, it is not surprising 

that the laminar boundary layer closely follows the expected laminar 

solution. Under the same high FSTI inlet conditions, however, Volino 

et al. [30] showed that a non-turbulent boundary layer may be strongly 

influenced by the free-stream turbulence and exhibit large deviation 

from laminar behavior. In the present study, the acceleration 

parameter, K, is in some of the cases over 10 times larger than in the 

Volino et al. [30] study, and the boundary layer thickness is as little as 

1/5 that in [30]. These differences apparently limit the free-stream 

effect on the upstream boundary layer, resulting in the observed 

laminar-like behavior. 

 Fig. 5 shows profiles of the rms fluctuating streamwise velocity, u′ , 

for the 50 upstream profiles. The low FSTI u′ is mainly streamwise 

unsteadiness that scales with the local U∞ . Values are low 

everywhere, increasing slightly from the free-stream value to a peak at 

y/θ=3 and then dropping to zero at the wall. For the high FSTI cases 

the free-stream u′ level does not change significantly within the test 

section and scales with Ue . Since the boundary layer fluctuations are 

caused by the free-stream fluctuations, the u′ profiles collapse when 

normalized on Ue . The collapse is not perfect since the free-stream 

u′ does drop somewhat as the eddies are strained in the accelerating 

flow. Qualitatively the behavior is the same as in the low FSTI cases, 

with high values in the free-stream rising to a peak at y/θ=3 and 

dropping to zero near the wall. The peak in u′ has lower magnitude 

and is farther from the wall than would be expected in a turbulent 

boundary layer. In all cases the boundary layer is clearly laminar-like 

in spite of the high u′ level. This is shown in both the mean velocity 

profiles (Fig. 3) and the local skin friction coefficients (Fig. 4) which 

follow laminar flow solutions. 

NASA/CR—2012-217656 79
6 Copyright © 2002 by ASME 

0 1 2 3 4 5 6 7 8 9 10 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

High FSTI

y/θ

U/U∞

Low FSTI

Fig. 3: Mean velocity profiles from Station 1-5, all cases 

0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2

x 105

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

cf=1.271/Res

1/2

Res

cf

Re=25K,Low FSTI

Re=50K,Low FSTI

Re=100K,Low FSTI

Re=200K,Low FSTI

Re=300K,Low FSTI

Re=25K,High FSTI

Re=50K,High FSTI

Re=100K,High FSTI

Re=200K,High FSTI

Re=300K,High FSTI

Laminar Solution

Fig. 4: Skin friction coefficients from Station 1-6, all cases 

Low FSTI Transition 

 Downstream of Station 6, the pressure gradient becomes adverse 

and separation and transition occur. Figures 6 through 10 show 

profiles at Stations 6-11 of U, u′ , the turbulent shear stress, − vu ′′ , and 

the intermittency, γ, for the five low FSTI cases. The v′ profiles are 

qualitatively very similar to the − vu ′′ profiles in all cases and are not 

shown. In the Re=25,000 case (Fig. 6), the mean velocity profile has 

just separated at Station 7, and the separation bubble grows 

continuously larger at the downstream stations. There is no 

reattachment, but rather a massive separation with a burst bubble at the 

trailing edge. The intermittency is nearly zero everywhere, indicating 

that the shear layer remains essentially laminar. Although the flow is 

laminar, u′ is non-zero. The u′ peak grows as the flow moves 

downstream, and its location is concurrent with the inflection point in 

the mean profile. The − vu ′′ values remain near zero through Station 

10, but then rise at Station 11 with a peak in the shear layer at the same 

0 2 4 6 8 10 12 14 16 18 20 0

0.002

0.004

0.006

0.008

0.01

0.012

0.014

0.016

0.018

0.02

(a) Low FSTI

y/θ

u′/U∞

Station 1

2

3

4

5

0 2 4 6 8 10 12 14 16 18 20 0

0.05

0.1

(b) High FSTI

y/θ

u′/Ue

Station 1

2

3

4

5

Fig. 5: Boundary layer u′ profiles from Stations 1-5; (a) low 

FSTI cases, (b) high FSTI cases 

location as the u′ peak. The magnitude of this peak is extremely low; 

the eddy viscosity at the peak is only about 1/30th of the molecular 

kinematic viscosity. Although not significant in terms of eddy 

transport, this peak may signify the beginning of transition. 

 Figure 7 shows the profiles for the Re=50,000 case. Results are 

very similar to the Re=25,000 case of Fig. 6. The boundary layer does 

not reattach. The − vu ′′ profile at Station 11 again exhibits a low level 

peak. The dimensionless value of this peak is about three times that of 

the peak in Fig. 6, indicating that the Re=50,000 case may be closer to 

transitioning. 

 Results for the Re=100,000 case are shown in Fig. 8. The mean 

velocity profiles show that the boundary layer is on the verge of 

separating at Station 7, but is still attached. It has separated by Station 

8, and the separation bubble grows through Station 10, although it 

does not become as thick as in the lower Re cases. At Station 10, the 

mean velocity near the wall rises slightly above zero, indicating the 

beginning of reattachment. At Station 11 the boundary layer is clearly 

reattached. The intermittency is near zero through Station 10, and then 

NASA/CR—2012-217656 80
7 Copyright © 2002 by ASME 

0 0 0 0 0 0 1 0

0.02

0.04

0.06

0.08

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

0.06

0.08

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.1

0 0 0 0 0 0 0

0.02

0.04

0.06

0.08

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.0001

0 0 0 0 0 0 1 0

0.02

0.04

0.06

0.08

γ

y/Ls

d) 6 7 8 9 10 11

Fig. 6: Station 6-11 profiles for low FSTI, Re=25,000 case: 

(a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency 

0 0 0 0 0 0 1 0

0.02

0.04

0.06

0.08

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

0.06

0.08

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.1

0 0 0 0 0 0 0

0.02

0.04

0.06

0.08

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.0005

0 0 0 0 0 0 1 0

0.02

0.04

0.06

0.08

γ

y/Ls

d) 6 7 8 9 10 11

Fig. 7: Station 6-11 profiles for low FSTI, Re=50,000 case: 

(a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency 

suddenly increases to 1 at Station 11, indicating fully-turbulent flow as 

the boundary layer reattaches. The peak in the intermittency is well 

away from the wall, indicating that transition begins in the shear layer 

over the separation bubble. The u′ profiles exhibit a peak in the shear 

layer at Stations 8 and 9, similar to the behavior at the lower Re. At 

Station 10 there is an increase in u′ near the wall as reattachment 

begins. The turbulent shear stress profile rises above zero at Station 

10 with a dimensionless value that is an order of magnitude larger than 

the peak shown in Fig. 7 for the Re=50,000 case. At Station 11 the 

magnitude of the peak has increased by another order of magnitude 

and the boundary layer is clearly turbulent. The peak in − vu ′′ is well 

away from the wall, indicating that while the boundary layer is 

turbulent and reattached, it has not yet recovered to fully-developed 

turbulent conditions. 

0 0 0 0 0 0 1 0

0.02

0.04

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.2

0 0 0 0 0 0 0

0.02

0.04

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.01

0 0 0 0 0 0 1 0

0.02

0.04

γ

y/Ls

d) 6 7 8 9 10 11

Fig. 8: Station 6-11 profiles for low FSTI, Re=100,000 case: 

(a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency 

0 0 0 0 0 0 1 0

0.02

0.04

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.2

0 0 0 0 0 0 0

0.02

0.04

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.01

0 0 0 0 0 0 1 0

0.02

0.04

γ

y/Ls

d) 6 7 8 9 10 11

Fig. 9: Station 6-11 profiles for low FSTI, Re=200,000 case: 

(a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency 

 Figure 9 shows the profiles for the Re=200,000 case. The mean 

velocity profiles indicate that separation does not occur until near 

Station 8. There is a clear separation bubble at Station 9, and the 

boundary layer is reattached by Station 10. By Station 11 the mean 

profile appears to have recovered to a fully-developed turbulent shape. 

The intermittency jumps from near 0 at Station 9 to 1 at Station 10, 

indicating a rapid transition and reattachment. The magnitude of u′

increases similarly, from a small peak near the inflection point of the 

mean profile at Station 9 to high values throughout the boundary layer 

at Station 10. The turbulent shear stress profiles show the same 

sudden increase between Stations 9 and 10, and the peak is still away 

from the wall at Station 11, indicating that recovery from separation 

may not be fully complete. 

 The Re=300,000 case profiles are shown in Fig. 10. Separation 

again occurs near Station 8. By Station 9 the separation bubble has 

NASA/CR—2012-217656 81
8 Copyright © 2002 by ASME 

0 0 0 0 0 0 1 0

0.02

0.04

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.2

0 0 0 0 0 0 0

0.02

0.04

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.01

0 0 0 0 0 0 1 0

0.02

0.04

γ

y/Ls

d) 6 7 8 9 10 11

Fig. 10: Station 6-11 profiles for low FSTI, Re=300,000 case: 

(a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency 

clearly grown and the non-zero mean velocities near the wall indicate 

that the boundary layer is on the verge of reattachment. The u′ values 

increase greatly between Stations 8 and 9. The u′ profile has a peak 

in the shear layer over the separation bubble and a second peak near 

the wall, which is indicative of reattachment. The intermittency 

profile at Station 9 shows this same double peak. The boundary layer 

is reattached and fully turbulent by Station 10. The turbulent shear 

stress profile rises to a small but discernable non-zero level at Station 

9, corresponding to the beginning of reattachment and the rise of the 

intermittency. By Station 11, − vu ′′ =uτ

2

 near the wall, which would be 

expected for a fully-developed, attached turbulent boundary layer. 

 The velocity data of Figs. 6-10 agree with the pressure profiles of 

Fig. 2a. The separation locations agree, although the velocity profiles 

provide better resolution and indicate that separation does move 

downstream somewhat as Reynolds number increases. The 

reattachment locations indicated by the pressure profiles correspond to 

locations where the velocity profiles have clearly reattached. Incipient 

reattachment is visible in the mean velocity one station upstream of 

full reattachment in some cases. In all cases the magnitude of u′

begins to increase in the shear layer after separation. At first this 

increase in u′ occurs without a corresponding increase in − vu ′′ , 

which remains near zero, and the shear layer remains laminar. As the 

flow continues downstream, low but nonzero − vu ′′ values eventually 

appear in the shear layer, and at the higher Reynolds numbers this is 

quickly followed by a rapid rise of − vu ′′ , sudden transition to 

turbulence, and almost immediate reattachment of the boundary layer. 

At the two lowest Reynolds numbers, the initial rise in − vu ′′ was 

detected, but it occurred so far downstream that transition and 

reattachment never occurred. The present results are consistent with 

those of Hatman and Wang [10], Lou and Hourmouziadis [12] and 

Volino and Hultgren [13], who also considered low FSTI separated 

flow transition. They also reported rapid transition and attributed it to 

the breakdown of a Kelvin-Helmholtz type instability of the shear 

layer. The transition mechanism in the present study will be discussed 

in more detail in Part 2 [29]. 

High FSTI Transition 

 Profiles for the high FSTI cases are shown in Figs. 11-15. For the 

Re=25,000 case (Fig. 11), the mean velocity profiles show that the 

boundary layer has separated by Station 6. The boundary layer 

appears on the verge of reattaching at Stations 10 and 11, but is not 

clearly reattached. The u′ level rises rapidly after separation, with a 

peak in the shear layer at each station. The free-stream buffets the 

shear layer, forcing fluid across a large mean velocity gradient, dU/dy, 

which causes high u′ levels. The same effect is present in the 

attached boundary layer upstream (Fig. 5b), but is damped somewhat 

by the wall. Free-stream buffeting and high u′ do not necessarily 

imply turbulent transport, and the turbulent shear stress remains low 

through Station 9. The − vu ′′ level rises to high levels at Stations 10 

and 11, but the peak is in the shear layer and drops to zero at the wall. 

Perhaps at this very low Reynolds number, even significant transport 

in the shear layer is insufficient to promote full reattachment of the 

boundary layer. Turbulent reattachment may be an intermittent 

phenomenon, related to and much like transition. At the lower Re it 

may occur over an extended distance. The intermittency indicates that 

the flow remains non-turbulent, in spite of the high levels of − vu ′′ at 

the downstream stations. The intermittency function, as defined 

above, only declares the flow turbulent when the velocity fluctuations 

include a full range of both large and small scales. The apparent 

mismatch between the − vu ′′ and γ profiles of Fig. 11 may indicate 

that the fluctuations which cause the turbulent shear stress initially do 

not include this range of scales. Turbulent shear stress spectra are 

presented in Part 2 [29]. Profiles of the wall normal fluctuating 

velocity, v′ , are also shown in Fig. 11. At Stations 7-9, high freestream values drop to zero at the wall, with no peak corresponding to 

the peak in u′ . The free-stream buffeting effect on v′ is damped by 

the wall, a phenomenon also observed in attached, non-turbulent 

boundary layers under high FSTI conditions (Volino et al. [30]). A 

peak emerges in v′ in the shear layer at Station 10 and 11, 

corresponding completely with the rise in − vu ′′ at these stations. The 

link between v′ and − vu ′′ was clear at all Re, making it unnecessary 

to present both v′ and − vu ′′ for the remaining cases. 

 The Re=50,000 case of Fig. 12 is very similar to the Re=25,000 

case. Reattachment is clearer, however, at Station 11, and the 

intermittency is non-zero at this station. Fig. 13 shows the 

Re=100,000 case. In this case the intermittency indicates that 

transition has begun by Station 9, which corresponds to an initial rise 

in − vu ′′ . The mean profile shows that the boundary layer is 

reattached at the last two stations. The high − vu ′′ peaks away from 

the wall indicate that the boundary layer has not fully recovered from 

the separation at Station 11. 

 It is not clear that separation occurs in the Re=200,000 case (Fig. 

14), but the velocity mean profile at Station 8 has an inflection point 

and appears to be close to separating. By Station 9 the boundary layer 

is clearly attached, the intermittency indicates transition is underway, 

u′ is high even near the wall, and − vu ′′ has risen to a turbulent level. 

At Station 11, − vu ′′ reaches a maximum equal to uτ

2

 near the wall, 

indicating a fully-developed attached turbulent flow. 

 The Re=300,000 results of Fig. 15 are very similar to those at 

Re=200,000. Transition begins slightly earlier at Re=300,000, with 

the intermittency greater than zero at Station 8. The thin boundary 

layer at Re=300,000 results in peaks in − vu ′′ at Stations 9 and 10 that 

are too close to the wall to resolve with the cross-wire probe. 

 In general, transition in the high FSTI cases began upstream of the 

locations in the corresponding low FSTI cases, and the transition 

NASA/CR—2012-217656 82
9 Copyright © 2002 by ASME 

0 0 0 0 0 0 1 0

0.02

0.04

0.06

0.08

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

0.06

0.08

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.2

0 0 0 0 0 0 0

0.02

0.04

0.06

0.08

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.01

0 0 0 0 0 0 1 0

0.02

0.04

0.06

0.08

γ

y/Ls

d) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

0.06

0.08

v′/Ue

y/Ls

e) 6 7 8 9 10 11

0.1

Fig. 11: Station 6-11 profiles for high FSTI, Re=25,000 case: 

(a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency, (e) v′

0 0 0 0 0 0 1 0

0.02

0.04

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.2

0 0 0 0 0 0 0

0.02

0.04

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.01

0 0 0 0 0 0 1 0

0.02

0.04

γ

y/Ls

d) 6 7 8 9 10 11

Fig. 12: Station 6-11 profiles for high FSTI, Re=50,000 case: 

(a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency 

region length was longer with high FSTI. This agrees with the 

observations of Volino and Hultgren [13], who also observed that 

transition was less abrupt with high FSTI. Boundary layer u′ levels 

are much higher with high FSTI, but the turbulent shear stress 

magnitude remains low until transition begins and does not appear to 

depend strongly on FSTI. 

Shape Factor and Momentum Thickness 

 As discussed above, stage losses cannot be determined 

quantitatively using a single passage test section, but is possible to 

0 0 0 0 0 0 1 0

0.02

0.04

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.2

0 0 0 0 0 0 0

0.02

0.04

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.01

0 0 0 0 0 0 1 0

0.02

0.04

γ

y/Ls

d) 6 7 8 9 10 11

Fig. 13: Station 6-11 profiles for high FSTI, Re=100,000 

case: (a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency 

0 0 0 0 0 0 1 0

0.02

0.04

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.2

0 0 0 0 0 0 0

0.02

0.04

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.01

0 0 0 0 0 0 1 0

0.02

0.04

γ

y/Ls

d) 6 7 8 9 10 11

Fig. 14: Station 6-11 profiles for high FSTI, Re=200,000 

case: (a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency 

compute the momentum thickness of the suction side boundary layer 

at the trailing edge. As explained by Howell et al. [20], this 

momentum thickness is proportional to the suction side profile loss 

when the boundary layer shape factor and passage exit angle remain 

constant. For those cases in which the boundary layer reattaches and 

recovers to a fully developed turbulent shape, H is approximately 

equal to 1.4. In these cases the suction side profile loss is likely the 

dominant loss mechanism (Howell et al. [20]). In those cases in which 

the boundary layer does not fully reattach, or reattaches near the 

trailing edge, the shape factor will be very large due to a large 

displacement thickness. The momentum thickness may be relatively 

small compared to the displacement thickness in these cases, since the 

wall shear is essentially zero and θ will not grow significantly in the 

free-shear layer. Large losses would then be expected in the wake, 

downstream of the airfoil. 

NASA/CR—2012-217656 83
10 Copyright © 2002 by ASME 

0 0 0 0 0 0 1 0

0.02

0.04

U/Ue

y/Ls

a) 6 7 8 9 10 11

0 0 0 0 0 0 0

0.02

0.04

u′/Ue

y/Ls

b) 6 7 8 9 10 11

0.2

0 0 0 0 0 0 0

0.02

0.04

−u′v′/Ue

2

y/Ls

c) 6 7 8 9 10 11

0.01

0 0 0 0 0 0 1 0

0.02

0.04

γ

y/Ls

d) 6 7 8 9 10 11

Fig. 15: Station 6-11 profiles for high FSTI, Re=300,000 

case: (a) mean velocity, (b) u′, (c) -u′v′, (d) intermittency 

 Fig. 16 shows the shape factor at Station 11 (s/Ls=0.94) as a 

function of Re. In the Re=200,000 and 300,000 cases, the shape factor 

is approximately 1.4, indicating that the boundary layer is reattached 

and that the momentum thickness is a good indicator of overall losses. 

For the lower Re cases, H is significantly higher, particularly in the 

low FSTI cases. The lower H in the high FSTI cases indicates that 

high FSTI helps to keep the separation bubble thinner by promoting 

more mixing in the shear layer over the bubble and by inducing earlier 

transition and reattachment. Fig. 17 shows the Station 11 momentum 

thickness as a function of Re. High FSTI helps keep the separation 

bubble thinner, as shown above in the mean velocity profiles of Figs. 

6-15, which tends to result in lower θ and lower losses when the 

boundary layer reattaches. High FSTI also promotes increased mixing, 

however, which tends to increase θ. For the Re=300,000 cases, Fig. 

17 indicates that the second effect is more significant and θ is higher 

for the high FSTI case. Transition and reattachment occur sufficiently 

far upstream in the low FSTI case that the effect of the high FSTI in 

promoting even earlier reattachment is not enough to counter the 

enhanced mixing effects. This suggests that small, controllable 

separation bubbles may be acceptable or even desirable in some cases, 

as proposed by Hourmouziadis [1] for controlled diffusion blading. 

For the Re=200,000 and Re=100,000 cases, the earlier reattachment 

caused by high FSTI is more significant and θ is lower for the high 

FSTI cases. This result is consistent with the observation in several 

studies that unsteady wakes from upstream airfoils result in lower 

losses. Enhanced mixing in the shear layer explains the higher θ for 

the high FSTI cases at Re=25,000. These momentum thicknesses do 

not relate directly to losses since the shear layer does not reattach in 

either of the Re=25,000 cases. 

Comparison to Correlations 

 Hatman and Wang [10] discuss three modes of separated flow 

transition. Based on their criteria, the present cases all fall into their 

laminar-separation long-bubble category. The data support this; 

separation occurred before transition. Hatman and Wang [10], Mayle 

[2] and others suggest that the Thwaites [40] criteria, Reθ

2

K=-0.082, is 

a good predictor for laminar separation. A laminar, attached flow 

solution would put Reθ

2

K=-0.082 between Stations 6 and 7 in all cases 

0 0.5 1 1.5 2 2.5 3

x 105

1

2

3

4

5

6

7

Re

H

Low FSTI

HIGH FSTI

Fig. 16: Shape factor at Station11 

0 0.5 1 1.5 2 2.5 3

x 105

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

Re

θ [mm]

Low FSTI

HIGH FSTI

Fig. 17: Momentum thickness at Station 11 

of the present study. The presence of the separation bubble changes 

the local acceleration, however, which tends to move the location 

where Reθ

2

K=-0.082. Volino and Hultgren [13] found the Thwaites 

criteria to be a good predictor of separation and it appears to work well 

for the present study as well. Exact prediction of the separation point 

is not straightforward, however, due to the interdependence of local K

values and the separation bubble location. 

 Prediction of transition and reattachment is more difficult. Hatman 

and Wang [10] present a transition correlation based on low FSTI data 

which predicts that transition should not occur in any of the present 

cases. Clearly, however, transition and reattachment do occur. The 

− vu ′′ profiles indicate that transition is imminent even in the low 

FSTI, low Re cases. Volino and Hultgren [13] drew comparisons to 

correlations from Mayle [2] and Davis et al. [41] with mixed results. 

NASA/CR—2012-217656 84
11 Copyright © 2002 by ASME 

Table 3: Distance from separation location to start of 

transition as a fraction of Ls; measured values and 

correlation predictions for each case 

FSTI Re×

10-3

Measured Mayle [2] 

short 

bubble 

Mayle [2] 

long 

bubble 

Davis et 

al. [41] 

25 > 0.41 0.25 0.83 1.54 

50 > 0.33 0.16-0.19 0.53-0.64 0.77 

100 0.17-0.33 0.12-0.14 0.39-0.48 0.38 

200 0-0.17 0.09-0.11 0.31-0.36 0.19 

Low 

300 0-0.08 0.07-0.08 0.23-0.28 0.13 

25 > 0.33 0.27-0.29 0.89-0.98 0.17 

50 0.17-0.33 0.16-0.18 0.53-0.61 0.08 

100 0.08-0.17 0.12-0.14 0.39-0.48 0.04 

200 0-0.08 0.07-0.12 0.25-0.39 0.02 

High 

300 0-0.08 0.06-0.09 0.19-0.31 0.02 

Comparisons to the present data are similarly mixed. The Mayle 

correlations predict the distance from the separation point to the onset 

of transition based on Reθ at the separation location. He presents a 

correlation for short separation bubble length and a correlation for 

long bubble length, which is 3.3 times the short bubble length. The 

Davis et al [41] correlation also predicts the distance from separation 

to transition onset, but as a function of the FSTI. Table 3 presents the 

distance from separation to the start of transition, normalized on the 

suction surface length, for all cases of the present experiments and as 

predicted by the Mayle [2] and Davis et al. [41] correlations. The 

finite spacing between the measurement stations results in uncertainty 

in Reθ at separation and in the exact locations of separation and 

transition, so a range of values is given for each quantity in the table. 

The results of the present cases lie between the Mayle long and short 

bubble correlations to within the resolution of the measurement 

locations. The Davis et al. correlation tends to predict too long a 

distance for the low FSTI cases and too short a distance for the high 

FSTI cases. 

 Existing correlations appear to give reasonable rough estimates of 

separated flow transition in some cases, but they are not particularly 

accurate or robust predictors. The general agreement between the 

similar cases of the present study, Volino and Hultgren [13], and 

Simon et al. [17] suggests that prediction of separated flow transition 

should be possible to some extent. The differences between the results 

of these studies, noted above, suggest that very accurate prediction of 

the flow may prove difficult and strongly dependent on small 

differences in boundary conditions. It is doubtful that a simple, robust 

correlation can be developed to incorporate all relevant boundary 

condition effects and provide very accurate predictions. It is also 

questionable whether the boundary conditions could be specified 

accurately enough for actual engine conditions. Perhaps they can, 

with addition research, however, and it may be possible to improve 

predictions with advanced computational schemes. The difficulty of 

predicting transition suggests that it may be advantageous to develop 

flow control schemes to force transition to occur at desired locations 

rather than try to predict it under existing conditions. 

CONCLUSIONS 

 Separated flow transition has been documented for cases with 

Reynolds numbers ranging from 25,000 to 300,000 at both high and 

low FSTI. The following conclusions can be drawn from the results. 

1. The start of boundary layer reattachment occurs near the start 

of transition, and both depend strongly on Re and FSTI. 

2. High FSTI results in competing effects with regard to losses. It 

causes enhanced mixing, which tends to promote boundary 

layer growth and increase losses. At the same time, it 

promotes earlier transition and reattachment, which reduces 

boundary layer thickness and losses. At the highest Re, 

reattachment occurred shortly after separation regardless of 

FSTI level, and high FSTI resulted in higher losses. At the 

intermediate Re, high FSTI reduced losses. At the lowest Re, 

the boundary layer did not reattach even with high FSTI, so 

losses would be high regardless of FSTI level. 

3. The turbulent shear stress level can remain near zero in spite of 

high FSTI and high u′ in the boundary layer. The beginning 

of a rise in − vu ′′ signals the beginning of transition. In the 

lowest Re cases, transition did not occur and the boundary 

layer did not reattach, but the beginning of a rise in − vu ′′ was 

observed near the trailing edge. This hints that it may be 

possible to induce transition even at very low Re. 

4. The present results agree roughly with similar studies from the 

literature and existing correlations but there are significant 

differences. Attempts to control transition and force its 

location may prove more fruitful than prediction of unmodified 

flow. The present study provides an extensive, detailed 

baseline data set for ongoing flow control experiments. 

ACKNOWLEDGEMENTS 

 This work was sponsored by the NASA Glenn Research Center. 

The grant monitor is Dr. David Ashpis. Additional matching support 

was provided through a U.S. Naval Academy Recognition Grant. Mr. 

Dale Boyer of the Technical Support Department at the Naval 

Academy fabricated the airfoils for the test section. 

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[1] Hourmouziadis, J., 1989, “Aerodynamic Design of Low Pressure 

Turbines,” AGARD Lecture Series 167. 

[2] Mayle, R.E., 1991, “The Role of Laminar-Turbulent Transition in 

Gas Turbine Engines,” ASME Journal of Turbomachinery, 116, 

pp. 509-537. 

[3] Sharma, O.P., Ni, R.H., and Tanrikut, S., 1994, “Unsteady Flow 

in Turbines,” AGARD Lecture Series 195, Paper No. 5. 

[4] Hodson, H.P., 1991, “Aspects of Unsteady Blade-Surface 

Boundary Layers and Transition in Axial Turbomachines,”

Boundary Layers in Turbomachines, VKI Lecture Series 1991-

06. 

[5] Wisler, D.C., 1998, “The Technical and Economic Relevance of 

Understanding Boundary Layer Transition in Gas Turbine 

Engines,” Minnowbrook II, 1997 Workshop on Boundary Layer 

Transition in Turbomachines, LaGraff, J.E., and Ashpis, D.E., 

eds., NASA/CP-1998-206958, pp. 53-64. 

[6] Curtis, E.M., Hodson, H.P., Banieghbal, M.R., Denton, J.D., 

Howell, R.J., and Harvey, N.W., 1996, “Development of Blade 

Profiles for Low-Pressure Turbine Applications,” ASME Paper 

96-GT-358. 

[7] Halstead, D.E., Wisler, D.C., Okiishi, T.H., Walker, G.J., 

Hodson, H.P., and Shin, H.-W., 1997, “Boundary Layer 

Development in Axial Compressors and Turbines: Part 3 of 4 –

LP Turbines,” ASME Journal of Turbomachinery, 119, pp. 225-

237. 

[8] Solomon, W.J., 2000, “Effects of Turbulence and Solidity on the 

Boundary Layer Development in a Low Pressure Turbine,”

ASME Paper 2000-GT-0273. 

NASA/CR—2012-217656 85
12 Copyright © 2002 by ASME 

[9] Gier, J., and Ardey, S., 2001, “On the Impact of Blade Count 

Reduction on Aerodynamic Performance and Loss Generation in 

a Three-Stage LP Turbine,” ASME Paper 2001-GT-0197. 

[10] Hatman, A., and Wang, T., 1999, “A Prediction Model for 

Separated Flow Transition,” ASME Journal of Turbomachinery, 

121, pp. 594-602. 

[11] Sohn, K.H., DeWitt, K.J., and Shyne, R.J., 2000, Experimental 

Investigation of Boundary Layer Behavior in a Simulated Low 

Pressure Turbine,” ASME Journal of Fluids Engineering, 122, 

pp. 84-89. 

[12] Lou, W., and Hourmouziadis, J., 2000, “Separation Bubbles 

Under Steady and Periodic-Unsteady Main Flow Conditions,”

ASME Journal of Turbomachinery, 122, pp. 634-643. 

[13] Volino, R.J., and Hultgren, L.S., 2001, “Measurements in 

Separated and Transitional Boundary Layers Under Low-Pressure 

Turbine Airfoil Conditions,” ASME Journal of Turbomachinery, 

123, pp. 189-197. 

[14] Yaras, M.I., 2001, “Measurements of the Effects of PressureGradient History on Separation-Bubble Transition,” ASME Paper 

2001-GT-0193. 

[15] Murawski, C.G., Sondergaard, R., Rivir, R.B., Simon, T.W., 

Vafai, K., and Volino, R.J., 1997, “Experimental Study of the 

Unsteady Aerodynamics in a Linear Cascade with Low Reynolds 

Number Low Pressure Turbine Blades,” ASME Paper 97-GT-95. 

[16] Qiu, S., and Simon, T.W., 1997, “An Experimental Investigation 

of Transition as Applied to Low Pressure Turbine Suction 

Surface Flows,” ASME Paper 97-GT-455. 

[17] Simon, T.W., Qiu, S., and Yuan, K., 2000, “Measurements in a 

Transitional Boundary Layer Under Low-Pressure Turbine 

Conditions,” NASA/CR-2000-209957. 

[18] Brunner, S., Fottner, L., and Schiffer, H.-P., 2000, “Comparison 

of Two Highly Loaded Low Pressure Turbine Cascades Under 

the Influence of Wake-Induced Transition,” ASME Paper 2000-

GT-268. 

[19] Stadtmüller, P., Fottner, L., and Fiala, A., 2000, “Experimental 

and Numerical Investigation of Wake-Induced Transition on a 

Highly Loaded LP Turbine at Low Reynolds Numbers,” ASME 

Paper 2000-GT-0269. 

[20] Howell, R.J., Ramesh, O.N., Hodson, H.P., Harvey, N.W., and 

Schulte, V., 2001, “High Lift and Aft-Loaded Profiles for LowPressure Turbines,” ASME Journal of Turbomachinery, 123, pp. 

181-188. 

[21] Kaszeta, R.W., Simon, T.W., and Ashpis, D.E., 2001, 

“Experimental Investigation of Transition to Turbulence as 

Affected by Passing Wakes,” ASME Paper 2001-GT-195. 

[22] Dorney, D.J., Ashpis, D.E., Halstead, D.E., and Wisler, D.C., 

1999, “Study of Boundary Layer Development in a Two-Stage 

Low Pressure Turbine,” AIAA Paper 99-0742; also NASA TM1999-208913. 

[23] Chernobrovkin, A., and Lakshminarayana, B., 1999, “Turbulence 

Modeling and Computation of Viscous Transitional Flow for 

Low Pressure Turbines,” Proc. 4th International Symposium on 

Engineering Turbulence Modeling and Measurements, Corsica, 

France. 

[24] Huang, P.G., and Xiong, G., 1998, “Transition and Turbulence 

Modeling of Low Pressure Turbine Flows,” AIAA Paper 98-

0039. 

[25] Thermann, H., Müller, M., and Niehuis, R., 2001, “Numerical 

Simulation of the Boundary Layer Transition in Turbomachinery 

Flows,” ASME Paper 2001-GT-0475. 

[26] VanTreuren, K.W., Simon, T., von Koller, M., Byerley, A.R., 

Baughn, J.W., and Rivir, R., 2001, “Measurements in a Turbine 

Cascade Flow Under Ultra Low Reynolds Number Conditions,”

ASME Paper 2001-GT-0164. 

[27] Lake, J.P., King, P.I., and Rivir, R.B., 2000, “Low Reynolds 

Number Loss Reduction on Turbine Blades With Dimples and VGrooves,” AIAA Paper 00-738. 

[28] Bons, J.P., Sondergaard, R., and Rivir, R.B., 2001, “Turbine 

Separation Control Using Pulsed Vortex Generator Jets,” ASME 

Journal of Turbomachinery, 123, pp. 198-206. 

[29] Volino, R.J., 2002, “Separated Flow Transition under Simulated 

Low-Pressure Turbine Airfoil Conditions: Part 2 - Turbulence 

Spectra,” ASME Paper GT-2002-30237. 

[30] Volino, R.J., Schultz, M.P., and Pratt, C.M., 2001, “Conditional 

Sampling in a Transitional Boundary Layer Under High FreeStream Turbulence Conditions,” ASME Paper 2001-GT-0192. 

[31] Chung, J.T., and Simon, T.W., 1990, “Three-Dimensional Flow 

Near the Blade/Endwall Junction of a Gas Turbine: Visualization 

in a Large-Scale Cascade Simulator, ASME Paper 90-WA/HT-4. 

[32] Aunapu, N.V., Volino, R.J., Flack, K.A., and Stoddard, R.M., 

2000, “Secondary Flow Measurements in a Turbine Passage With 

Endwall Flow Modification,” ASME Journal of Turbomachinery, 

122, pp. 651-658. 

[33] Volino, R.J., and Simon, T.W., 1997, “Velocity and Temperature 

Profiles in Turbulent Boundary Layers Experiencing Streamwise 

Pressure Gradients,” ASME Journal of Heat Transfer, 119, pp. 

433-439. 

[34] Wills, J.A.B., 1962, “The Correction of Hot-Wire Readings for 

Proximity to a Solid Boundary,” Journal of Fluid Mechanics, 12, 

pp. 65-92. 

[35] Ligrani, P.M., and Bradshaw, P., 1987, “Spatial Resolution and 

Measurement of Turbulence in the Viscous Sublayer Using 

Subminiature Hot-Wire Probes,” Experiments in Fluids, 5, pp. 

407-417. 

[36] Ligrani, P.M., and Bradshaw, P., 1987, “Subminiature Hot-Wire 

Sensors: Development and Use,” Journal of Physics E: Scientific 

Instruments, 20, pp. 323-332. 

[37] Ligrani, P.M., Westphal, R.V., and Lemos, F.R., 1989, 

“Fabrication and Testing of Subminiature Multi-Sensor Hot-Wire 

Probes,” Journal of Physics E: Scientific Instruments, 22, pp. 

262-268. 

[38] Bradshaw, P., 1994, “Turbulence: the Chief Outstanding 

Difficulty of Our Subject,” Experiments in Fluids, 16, pp. 203-

216. 

[39] Volino, R.J., 2002, “An Investigation of the Scales in Transitional 

Boundary Layers Under High Free-Stream Turbulence 

Conditions,” ASME Paper GT-2002-30233. 

[40] Thwaites, B., 1949, “Approximate Calculations of the Laminar 

Boundary Layer,” Aeronautical Quarterly, 7, pp. 245-280. 

[41] Davis, R.L., Carter, J.E., and Reshotko, E., 1985, “Analysis of 

Transitional Separation Bubbles on Infinite Swept Wings,” AIAA 

Paper 85-1685. 

NASA/CR—2012-217656 86
1 Copyright © 2002 by ASME 

Proceedings of 

ASME TURBO EXPO 2002 

June 3-6, 2002, Amsterdam, The Netherlands

GT-2002-30237 

SEPARATED FLOW TRANSITION UNDER SIMULATED LOW-PRESSURE TURBINE 

AIRFOIL CONDITIONS: PART 2 – TURBULENCE SPECTRA

Ralph J. Volino

Department of Mechanical Engineering 

United States Naval Academy 

Annapolis, Maryland 21402 

Email: volino@usna.edu

ABSTRACT 

 Spectral analysis was used to investigate boundary layer separation, 

transition and reattachment under low-pressure turbine airfoil 

conditions. Cases with Reynolds numbers ranging from 25,000 to 

300,000 (based on suction surface length and exit velocity) have been 

considered at low (0.5%) and high (9% inlet) free-stream turbulence 

levels. Spectra of the fluctuating streamwise velocity and the turbulent 

shear stress are presented. The spectra for the low free-stream 

turbulence cases are characterized by sharp peaks. The high freestream turbulence case spectra exhibit more broadband peaks, but 

these peaks are centered at the same frequencies observed in the 

corresponding low turbulence cases. The frequencies of the peaks 

suggest that a Tollmien-Schlichting instability mechanism drives 

transition, even in the high turbulence cases. The turbulent shear 

stress spectra proved particularly valuable for detection of the early 

growth of the instability. The predictable nature of the instability may 

prove useful for future flow control work. 

NOMENCLATURE 

FSTI free-stream turbulence intensity 

f frequency in Hz 

Ls suction surface length 

PSD power spectral density of 2 u′ or − vu ′′

Re LU se /ν , exit Reynolds number 

Reδ∗ U∞δ

∗

/ν displacement thickness Reynolds number 

s streamwise coordinate, distance from leading edge 

U∞ local free-stream velocity 

Ue nominal exit free-stream velocity 

u′ streamwise fluctuating velocity 

− vu ′′ turbulent shear stress 

v′ wall normal fluctuating velocity 

y cross-stream coordinate, distance from wall 

δ

*

 displacement thickness 

δs shear layer thickness 

δ99.5 99.5% boundary layer thickness 

ν kinematic viscosity 

INTRODUCTION 

 In Part 1 of the present study [1], the significance of boundary layer 

separation, transition, and reattachment to the flow over modern lowpressure turbine airfoils was discussed. Measured mean velocity and 

statistical turbulence quantities were presented for cases with high and 

low free-stream turbulence intensity (FSTI) and Reynolds numbers 

(based on suction surface length and exit velocity) ranging from 

25,000 to 300,000. The separation point tended to move downstream 

somewhat as Reynolds number increased. Transition and reattachment 

locations moved upstream significantly as Re or FSTI were increased. 

 While the statistical quantities presented in Part 1 [1] provide a 

quantitative description of what happens under different Re and FSTI

conditions, they do not explain the transition mechanism. To better 

explain the transition process and accurately predict or control it, an 

understanding of the physics which cause the results observed in Part 1 

[1] is needed. The present paper uses spectral analysis to investigate 

separated flow transition. 

 Mayle [2] classified the modes of transition as “natural transition,” 

“bypass” transition; “separated flow” transition of the shear layer over 

a separation bubble; “periodic-unsteady” transition, which might also 

be called wake-induced bypass transition; and reverse transition. 

Under low FSTI, zero streamwise pressure gradient conditions, natural 

transition is expected. This type of transition has been extensively 

documented and can be predicted with linear stability analysis. As 

described by Schlichting [3], when the displacement thickness 

Reynolds number exceeds a critical value, the boundary layer becomes 

unstable to small disturbances, which begin to grow as TollmienSchlichting (TS) waves. These waves eventually become three 

dimensional and result in turbulent spots. Under high FSTI, zero 

pressure gradient conditions, large disturbances can cause a bypass of 

the linear growth stages of transition, resulting in the sudden 

appearance of turbulent spots. At intermediate FSTI, elements of both 

bypass and natural transition may be observed. Sohn and Reshotko 

[4], for example, presented data for a 1% FSTI case, showing both 

spectral peaks at possible TS frequencies and broadband unsteadiness 

more typical of bypass transition. 

 In some cases, transition is observed even though linear stability 

theory predicts that the boundary layer should not develop TS waves. 

NASA/CR—2012-217656 87
2 Copyright © 2002 by ASME 

Volino [5], for example, considered a favorable pressure gradient case 

with high FSTI that clearly underwent transition. The boundary layer 

thickness remained low in this case due to the acceleration, resulting in 

Reδ∗ below the critical limit for linear instability. 

 Separated flow transition could potentially include elements of 

either natural or bypass transition. In separated flow cases the 

pressure gradient is adverse, resulting in a boundary layer or shear 

layer that typically is unstable to TS waves. High FSTI, however, 

might be the dominant factor in a separated shear layer, overwhelming 

the effect of any TS waves and producing bypass transition. Hughes 

and Walker [6] list several studies with FSTI below 0.9% in which TS 

waves were detected in adverse pressure gradient cases. They also 

note that Halstead et al. [7] did not detect TS waves in the flow 

through a rotating cascade with more representative, higher FSTI. 

Solomon and Walker [8], however, provide evidence of TS waves 

under conditions similar to those of Halstead et al. [7]. Hughes and 

Walker [6] considered a flow with wakes, in which the FSTI between 

wakes ranged from less than 1% to about 3%, and the FSTI in the 

wakes was about 8%. They provide clear evidence of TS waves. 

 Hatman and Wang [9], Volino and Hultgren [10], and Lou and 

Hourmouziadis [11] all considered low FSTI, adverse pressure 

gradient flows and observed transition in the shear layer over 

separation bubbles. Spectral data in all three studies showed clear 

evidence of an instability along with harmonics. It was expected that 

this instability was very similar to the Kelvin-Helmholtz instabilities 

observed in free shear layers, although the unstable frequencies were 

somewhat different than expected for free shear layers since the 

separation bubbles were bounded by the wall on one side. Volino and 

Hultgren [10] also considered high FSTI cases and observed 

broadband unsteadiness in the spectra of the streamwise fluctuating 

velocity, u′ . Spikes at discreet frequencies, which were observed in 

the low FSTI cases, were not present. They stated that transition in the 

high FSTI cases appeared to be through a bypass mode. 

 Clearly there is some disagreement regarding the transition 

mechanism in separated boundary layers, particularly under high FSTI

conditions. Some of these differences may stem from physical 

differences in the boundary conditions between the various studies. 

The present study addresses the issue through spectral analysis of 

flows over a range of Reynolds numbers at both high and low FSTI. 

Included in the analysis are spectra of the turbulent shear stress, which 

were not considered in previous studies. 

EXPERIMENTS 

 The experimental facility and the cases considered are described in 

detail in Part 1 [1]. A low speed wind tunnel supplies air to a singlepassage cascade-simulator with geometry and flow angles matching 

those for the industry supplied Pak-B airfoil. For the low FSTI cases, 

the background turbulence level for the wind tunnel is nominally 

0.5%, and consists primarily of low frequency unsteadiness. A passive 

grid is used to generate a high inlet FSTI of 8.7%. 

 Velocity data were acquired at 11 streamwise measurement stations 

along the spanwise centerline of the suction side of the passage. 

Station locations are given in Table 1. At each station, instantaneous 

streamwise velocity was measured at 60 locations as a single sensor 

hot-wire probe was traversed from the airfoil surface to the freestream. The voltage from the hot-wire was offset and amplified by a 

factor of 10 and low pass filtered at 10 kHz using signal conditioners 

(TSI model 157). At each location, data were acquired for 26 seconds 

at a 20 kHz sampling rate (219 samples). The high sampling rate 

provides an essentially continuous signal, which is needed for spectral 

processing. The long sampling time results in low uncertainty in both 

statistical and spectral quantities. Two component velocity 

Table 1: Measurement Stations 

Station 1 2 3 4 5 6 

s/Ls 0.111 0.194 0.278 0.361 0.444 0.527 

Station 7 8 9 10 11 

s/Ls 0.611 0.694 0.778 0.861 0.944 

Ls=228.6 

mm 

measurements were made at Stations 7-11 with a cross-wire probe. 

The upstream boundary layer was too thin for cross-wire 

measurements. Data were acquired at 25 locations in each profile, 

beginning 1 mm from the wall and extending to the free-stream. 

Sampling rates and times were the same as with the single sensor 

probe. Power spectra of u′ , the wall normal fluctuating velocity, v′ , 

and the turbulent shear stress, − vu ′′ , were computed for the data from 

all measurement locations. Uncertainties in the u′ , v′ and − vu ′′

spectra are all 10%. Frequencies are resolved from 4.88 Hz to 10 kHz 

in 4.88 Hz increments using a 4096 point Fast Fourier Transform to 

compute the spectra. As a check, the spectra were integrated with 

respect to frequency and found to equal the corresponding time 

averaged Reynolds stresses. 

 Data sets were acquired for cases at high and low FSTI with exit 

Reynolds numbers of 25,000, 50,000, 100,000, 200,000 and 300,000. 

The upstream boundary layer through Station 6 remained laminar in 

all of these cases, with the mean velocity following a Falkner-Skan 

wedge flow solution. Downstream of Station 6 the pressure gradient 

becomes adverse. Separation occurred between Stations 6 and 8, with 

the separation point moving downstream as Re increased. 

Reattachment did not occur in the high or low FSTI, Re=25,000 cases 

or the low FSTI, Re=50,000 case. In all other cases the boundary layer 

did reattach, and the reattachment point moved upstream as Re or FSTI

was increased. Details are available in Part 1 [1]. The u′ level rose in 

the shear layer in each case after the boundary layer separated, but the 

turbulent shear stress remained near zero until transition began. 

Transition began in the shear layer and quickly led to boundary layer 

reattachment. In the Re=25,000 cases, low but non-zero turbulent 

shear stress was observed at the most downstream stations, indicating 

that although the shear layer was still non-turbulent and separated, it 

was showing signs of the start of transition. 

 As discussed in Part 1 [1], the finite length of the hot-wire sensors 

(1.27 mm) will result in some spatial averaging and could result in 

attenuation of the measured fluctuating velocity components. Based 

on the results of Ligrani and Bradshaw [12, 13], it was explained in 

Part 1 that the errors in the rms fluctuating quantities are within the 

10% uncertainty estimates in the majority of cases in the present study. 

Exceptions occur for the Re=200,000 cases at Station 11 of the low 

FSTI case and Stations 9-11 of the high FSTI case. For the 

Re=300,000 cases, larger errors are expected for Stations 10-11 of the 

low FSTI case and Stations 9-11 of the high FSTI case. For these 

cases, errors may be as large as 30% near the wall, but should be under 

10% at y locations greater than 1 mm. As explained by Ligrani et al. 

[14], measurements with the cross-wire probe are subject to potentially 

larger errors due to the finite spacing (1 mm) between the two sensors. 

These errors will be largest near the wall, but become smaller than the 

10% uncertainty for y locations above 1 mm. For this reason, cross 

wire measurements were only made for y>1 mm. 

 The spectra in the present paper are presented to show the energy 

content of the fluctuating quantities as a function of frequency. The 

average errors in these spectra, therefore, should be the same as those 

given above for the corresponding rms quantities. The errors will not 

be uniform with respect to frequency across the spectra, however. As 

explained by Ligrani and Bradshaw [13], spatial averaging effects will 

be most severe for the smallest scales (highest frequencies) in the 

NASA/CR—2012-217656 88
3 Copyright © 2002 by ASME 

flow. Applying the spectral results of Ligrani and Bradshaw [13] to 

the present study, errors due to spatial averaging at y locations above 1 

mm will rise above 10% at frequencies above 150, 300, 600, 1200 and 

1800 Hz for the Re=25,000, 50,000, 100,000, 200,000, and 300,000 

cases respectively. Below these frequencies the errors should be under 

10%. These frequencies are all above the frequencies of the spectral 

peaks in the results presented below, so the peaks should not be 

significantly attenuated. For the Re=25,000 and 50,0000 cases, all 

significant energy in the spectra is below the frequencies given above, 

so there is no significant attenuation of the results at any frequency. 

For the Re=100,000 case, only the high frequency “tail” of the spectra 

will be subject to significant error. Magnitudes at these high 

frequencies may be as much as 30% low. Similarly for the 

Re=200,000 and 300,000 cases, errors will only rise above 10% at 

frequencies about twice those of the spectral peaks. Magnitudes at the 

highest frequencies may be as much as 40% low in these cases. 

FREQUENCIES OF INTEREST 

 Transition in the shear layer over the separation bubble could be 

initiated by Tollmien-Schlichting waves originating in the boundary 

layer upstream of separation. Upstream of Station 6, the pressure 

gradient is favorable and the boundary layer should not develop TS 

waves at any frequencies. At Station 6 the flow begins to decelerate 

and immediately becomes unstable. Walker [15] provides the 

following equation for the frequency of maximum amplification rate 

for TS waves. 

2/3* 2 2.3/2 Re− ∞ = δ πν Uf (1) 

The TS frequencies predicted by Eq. (1) are listed in Table 2. Values 

are given for Stations 6 and 7 for the cases in which the boundary 

layer is still attached at these stations. 

 Another possible path to transition is breakdown of the shear layer 

though a Kelvin-Helmholtz type instability. The frequency of the 

instability should scale with the velocity change across the shear layer 

and inversely with the shear layer thickness. Since the velocity in the 

separation bubble is nearly zero, the velocity difference across the 

shear layer equals the local free-stream velocity. The shear layer 

thickness, δs, is determined from the mean velocity profiles presented 

in Part 1 [1]. Table 2 lists the quantity U s ∞ /δ for each station where 

the boundary layer is separated. 

RESULTS 

Free-Stream Spectra

 The u′ and v′ free-stream spectra are shown in Fig. 1 for each 

station of the low FSTI, Re=300,000 case. Frequency is plotted on a 

log scale versus frequency times power spectral density on a linear 

scale. In these coordinates the area under the curve in any frequency 

band is proportional to the contribution to the quantity of interest in 

that band. The u′ spectra are dominated by low amplitude 

unsteadiness at frequencies below 20 Hz. The u′ spectra for the lower 

Re cases (not shown) have proportionately lower amplitudes, but the 

frequency range remains the same. The frequencies associated with 

turbulent eddies would be expected to scale with the free-stream 

velocity. Since the frequencies in the present cases remain constant as 

Re is changed, the unsteadiness in the wind tunnel is most likely not 

associated with turbulent eddies. The v′ spectra show the same low 

frequencies as u′ but at 1/10th the magnitude. A second lower peak is 

centered at about 100 Hz. This peak is also present with the same 

magnitude in u′ , and is visible in an expanded version of Fig. 1a. The 

Table 2: Most unstable Tollmien-Schlichting frequencies 

prior to separation and s ∞ /U δ values in shear layer 

TS freq. [Hz] U δ s ∞ / [s-1 ×10-3] 

Station Station 

FSTI Re 

×10-3

6 7 6 7 8 9 10 11 

25 80 1.0 0.8 0.6 0.5 0.4 0.4 

50 176 2.1 1.5 1.3 1.2 1.0 

100 443 266 3.9 3.2 3.8 

200 1175 645 11 8.7 

Low 

300 1554 1134 22 20 

25 60 1.1 0.8 0.6 0.4 0.4 0.3 

50 174 3.1 2.4 1.6 1.1 0.8 

100 408 272 4.1 3.1 

200 1138 732 15 

High 

300 1747 1195 25 

100 101 102 103 0

0.001

0.002

0.003

0.004

0.005

0.006

0.007

0.008

0.009

0.01

f [Hz]

f⋅PSD(u′

2) [m2/s2

]

(a)

Station 1

2

3

4

5

6

7

8

9

10

11

100 101 102 103 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1 x 10−3

f [Hz]

f⋅PSD(v′

2) [m2/s2

]

(b)

Station 7

8

9

10

11

Fig. 1: Free-stream spectra for low FSTI, Re=300,000 case; 

(a) u′, (b) v′

NASA/CR—2012-217656 89
4 Copyright © 2002 by ASME 

100 101 102 103 104 0

0.05

0.1

0.15

0.2

0.25

f [Hz]

f⋅PSD(u′

2) [m2/s2

]

(a)

Station 1

2

3

4

5

6

7

8

9

10

11

100 101 102 103 104 0

0.1

0.2

0.3

0.4

0.5

0.6

f [Hz]

f⋅PSD(v′

2) [m2/s2

]

(b)

Station 7

8

9

10

11

Fig. 2: Free-stream spectra for high FSTI, Re=300,000 case; 

(a) u′, (b) v′

frequency of this second peak scales with the free-stream velocity and 

is proportionately lower for the lower Re cases. Its magnitude is very 

low, and by itself would correspond to an rms turbulence level of 

about 0.01%. It is likely the residual turbulence remaining after the 

wind tunnel screens. 

 The u′ and v′ free-stream spectra for the high FSTI Re=300,000 

case are shown in Fig. 2. The magnitude of u′ is about 25 times 

larger than in the low FSTI case of Fig. 1a. At the upstream stations 

there is a peak at about 30 Hz. This peak decays due to streamwise 

straining of the flow as it is accelerated through Station 6. In the 

adverse pressure gradient region downstream of Station 6, a broadband 

peak emerges centered at about 700 Hz. In the lower Re cases, the u′

spectra is qualitatively the same, but the magnitudes scale with 2 U∞

and the frequencies scale with U∞ . The v′ spectra of Fig. 2b all 

appear similar with a broadband peak centered at 60 Hz. As with u′ , 

the magnitudes and frequencies of the v′ spectra in the lower Re cases 

scale with the free-stream velocity, and appear qualitatively similar to 

those of Fig. 2b. 

101 102 103 0

0.002

0.004

0.006

0.008

0.01

0.012

0.014

0.016

0.018

0.02

1.46e−05

f [Hz]

y/Ls

Fig. 3: Contours of f.

PSD(u′

2

), Low FSTI, Re=100,000, 

Station 4; outer contour magnitude and contour spacing 

indicated by numerical value in field of figure 

101 102 103 0

0.001

0.002

0.003

0.004

0.005

0.006

0.007

0.008

0.009

0.01

0.00018

f [Hz]

y/Ls

Fig. 4: Contours of f.

PSD(u′

2

), High FSTI, Re=300,000, 

Station 4; outer contour magnitude and contour spacing 

indicated by numerical value in field of figure 

Upstream Boundary Layer 

 Contours of the boundary layer u′ spectra from Station 4 of the low 

FSTI, Re=100,000 case are shown in Fig. 3. Frequency in Hz is shown 

on the horizontal axis on a log scale, and distance from the wall 

normalized on the suction surface length is on the vertical axis on a 

linear scale. The frequency is left dimensional since there is no single 

appropriate normalization for all regions of the flow. The contours in 

Fig. 3 show the dimensionless magnitude of the spectra as 

22 f ⋅ PSD ′ /)( Uu ∞ . A slice through the data of Fig. 3 at a fixed 

distance from the wall would produce a spectrum in the coordinates of 

Fig. 2, except with dimensionless magnitude. The number (1.46e-05), 

NASA/CR—2012-217656 90
5 Copyright © 2002 by ASME 

0.000104

St. 10

0.00024

St. 11

0

0.02

0.04

1.4e−05

St. 7

(a)

y/Ls

2.2e−05

St. 8

5.2e−05

St. 9

101 102 103 104

0

0.02

0.04

1.74e−05

(e)

y/Ls

f [Hz]

0

0.02

0.04

1.76e−05

(d)

y/Ls

0

0.02

0.04

1.6e−05

(c)

y/Ls

0

0.02

0.04

1.4e−05

(b)

y/Ls

4.8e−05

4.6e−05

1.24e−05

101 102 103 104

1.02e−05

f [Hz]

0.000134

0.000158

4.2e−05

101 102 103 104

0.0028

f [Hz]

0.00026

0.0082

0.0088

101 102 103 104

0.0066

f [Hz]

0.00056

0.022

0.0074

101 102 103 104

0.002

f [Hz]

Fig. 5: Contours of 22 f ⋅PSD ′ U/)u( ∞ , Low FSTI cases; Station number indicated above each column; by row (a) Re=25,000, (b) 

Re=50,000, (c) Re=100,000, (d) Re=200,000, (e) Re=300,000; outer contour magnitude and contour spacing indicated by 

numerical value in field of each subplot, 5 contours shown in each subplot

which appears in the center of the figure, indicates the magnitude of 

the outermost contour and the contour spacing. Hence, the outermost 

contour value is 1.46×10-5, the next contour value is 2.92×10-5, the 

next is 4.38×10-6, etc. The same format is used in all the figures which 

follow. Figure 3 shows a peak near the wall centered at 50 Hz and 

unsteadiness below 20 Hz extending from the free-stream to near the 

wall. The results shown in Fig. 3 are typical of the behavior at 

Stations 1-6 in all the low FSTI cases. The 50 Hz peak was also 

visible in the Re=200,000 and 300,000 cases, but not at the two lower 

Re. This peak did not appear to have any significant effect on the 

downstream boundary layer. 

 Figure 4 shows the spectra from Station 4 of the high FSTI, 

Re=300,000 case in the coordinates of Fig. 3. The Fig. 4 spectra are 

typical of the spectra at Stations 1-6 of all the high FSTI cases. There 

is a near wall peak at 100 Hz, which is 1.7 times the frequency of the 

free-stream v′ frequency peak shown in Fig. 2b. The distance of the 

peak from the wall is larger in the lower Re cases, scaling with the 

boundary layer thickness. The frequency of the peak scales with Re, 

and is consistently 1.7 times the dominant v′ frequency in the freestream. This indicates that the boundary layer unsteadiness is induced 

by free-stream buffeting. Volino [5] discusses free-stream buffeting of 

boundary layers in more detail. 

Low FSTI Transition 

 Upsteam of Station 6, the boundary layer unsteadiness is 

attributable to low amplitude streamwise unsteadiness in the low FSTI

cases, and free-stream buffeting in the high FSTI cases. Buffeting, as 

explained in Part 1 [1], refers to the effect of free-stream pressure 

fluctuations on the boundary layer as fluid is pushed in the wall 

normal direction across the gradient in the mean streamwise velocity. 

Downstream of Station 6, the pressure gradient becomes adverse, and 

the spectra become more interesting. Figure 5 shows the u′ spectra 

contours for Stations 7-11 of all the low FSTI cases. The coordinates 

of each subplot are the same as those of Fig. 3. In the Re=25,000 case 

NASA/CR—2012-217656 91
6 Copyright © 2002 by ASME 

Fig. 6: Spectra of 22 f ⋅PSD ′ U/)u( ∞ , Low FSTI, Re=100,000, 

Station 10

 (top row of Fig. 5), the contours at Station 7 show the low frequency 

unsteadiness observed at the upstream stations. Between Stations 7 

and 10, the magnitude of the peak increases by an order of magnitude. 

The peak at each station moves away from the wall. Comparison to 

the mean velocity profiles presented in Part 1 [1] shows that the 

location of the peak corresponds, not surprisingly, to the shear layer 

over the separation bubble. The Re=50,000 case (row 2) shows similar 

behavior through Station 10, but a second peak emerges at about 78 

Hz at Station 11. This second peak is indicative of transition. 

 In the Re=100,000 case (row 3 of Fig. 5), a sharp peak appears at 

Station 10 at 273 Hz. The contour spacing increases by two orders of 

magnitude between Stations 9 and 10, indicating a similar increase in 

the magnitude of the peak. The peak is so sharp and so large that it 

appears as a line in the contour plot. Figure 6 shows this Re=100,000, 

Station 10 data in the same coordinates as Fig. 5, but three 

dimensionally. The low frequency unsteadiness visible at Station 9 is 

still present, but since its magnitude is only 1/50th that of the 273 Hz 

peak, it is barely visible in Fig. 6. The sharp peak in the shear layer is 

typical of all the low FSTI cases. 

 In the Re=200,000 case (row 4 of Fig. 5), a small peak appears in 

the shear layer at 698 Hz at Station 9. By Station 10 this peak has 

increased in size by two orders of magnitude and is becoming more 

broadband as the boundary layer reattaches and becomes turbulent. 

Although not clear in Fig. 5, the sharp spectral peak remains in the 

center of this broadband turbulence and is clear in the format of Fig. 6. 

Similar behavior is apparent in the Re=300,000 case, but the spectral 

peak and subsequent turbulence are centered at 922 Hz. 

 Contours of the normalized turbulent shear stress spectra, 

( ) 2 f PSD −⋅ ′′ /Uvu ∞ , for the low FSTI cases are shown in Fig. 7. The 

format is the same as in Fig. 5. The shear layer is laminar at Stations 7 

and 8 in all cases, and the turbulent shear stress is near zero. The 

Re=25,000 case shows a low magnitude, 15 Hz peak at Station 10. 

The peak was obscured in the u′ spectra of Fig. 5 due to the presence 

of other streamwise unsteadiness at similar frequencies. The 

magnitude of the 15 Hz shear stress peak increases by a factor of 15 

between Station 10 and 11, but its amplitude is still quite low. Close 

inspection of the u′ spectra (Fig. 5) at Station 11 shows a double peak, 

with one peak at 15 Hz. The mean profiles of Part 1 [1] indicate that 

Table 3: Measured peak frequencies in spectra 

FSTI Re×10-3 25 50 100 200 300 

Low Station 10 10 9 9 9 

f [Hz] 15 78 273 698 922 

High Station 9 9 9 9 9 

f[ Hz] 20 85 230 600 1100 

transition may be imminent, but the shear layer is still laminar at 

Station 11. 

 The Re=50,000 case shows similar behavior to that at Re=25,000. 

A turbulent shear stress peak appears at Station 10 and is 37 times 

larger by Station 11. This peak did not become visible until Station 11 

in the u′ spectra. Similarly, in the Re=100,000 case a sharp turbulent 

shear stress peak emerges at Station 9, one station upstream of its 

appearance in u′ . The shear stress peak appears at Station 9 in the 

Re=200,000 and 300,000 cases. In all cases, the peak appears at the 

same frequency in u′ and − vu ′′ , but tends to become visible earlier in 

the − vu ′′ spectra due to the lower magnitude of the low-frequency 

“noise” in − vu ′′ . 

 The frequencies of the spectral peaks are listed for all cases in Table 

3. Also indicated in the table are the stations at which these 

frequencies were determined. The frequencies of the peaks did not 

change significantly in the streamwise direction. Because the peaks do 

not appear until after the boundary layer has separated, there is some 

reason to believe that a Kelvin-Helmholtz type instability may be 

involved. For the low FSTI cases, comparison of the spectral peak 

frequencies to the U s ∞ /δ values in Table 2 shows that the spectral 

frequencies and U s ∞ /δ values both increase with Re, as expected. 

The ratio of the measured frequency to U s ∞ /δ is not a constant, 

however, ranging from about 0.04 to about 0.08. Comparison of the 

measured frequencies to the TS frequencies of Table 2 shows better 

correlation. At the three highest Re, the measured frequencies agree 

with the TS frequencies at Station 7 (the last station before separation) 

to within 20%. The boundary layer is already separated at Station 7 in 

the Re=25,000 and 50,000 cases, so the frequency comparison must be 

made at Station 6. Agreement with the TS frequencies is not as good 

in these cases as at the high Re. Hughes and Walker [6] note that since 

the most unstable TS frequency changes with streamwise position, the 

frequency observed in the shear layer need not equal the TS value at 

any single upstream position. 

High FSTI Transition 

 Contours of the u′ spectra for the high FSTI cases are shown in Fig. 

8. In the Re=25,000 case (top row), there is low frequency 

unsteadiness at the same frequencies observed upstream (Fig. 4). As 

in the low FSTI cases, the peak u′ location is in the shear layer over 

the separation bubble. In the Re=50,000 case, the outermost contour at 

Station 10 extends to higher frequencies than at the upstream stations, 

and by Station 11 a new peak has emerged at 60 Hz, extending from 

the shear layer down to the wall. The magnitude of this new peak is 

about equal to the magnitude of the original, low frequency peak, 

which is also still visible. Similar behavior is clear for the Re=100,000 

case. Higher frequencies begin to emerge at Station 9, and a large, 

clear peak centered at 260 Hz is visible at Stations 10 and 11. In the 

Re=200,000 and 300,000 cases, some signs of higher frequencies are 

already visible at Station 8, and a double peak is clear at Station 9. At 

Stations 10 and 11, the higher frequency peak overwhelms the lower 

frequencies. 

NASA/CR—2012-217656 92
7 Copyright © 2002 by ASME 

1.72e−06

St. 10

2.6e−05

St. 11

0

0.02

0.04

1.04e−06

St. 7

(a)

y/Ls

1.4e−06

St. 8

4.6e−07

St. 9

101 102 103 104

0

0.02

0.04

2.8e−08

(e)

y/Ls

f [Hz]

0

0.02

0.04

9.8e−08

(d)

y/Ls

0

0.02

0.04

3.4e−07

(c)

y/Ls

0

0.02

0.04

1.58e−07

(b)

y/Ls

6.4e−07

1.66e−06

9.8e−08

101 102 103 104

1.5e−08

f [Hz]

1.02e−06

3.8e−06

1.54e−06

101 102 103 104

0.000168

f [Hz]

4.4e−06

0.005

0.0022

101 102 103 104

0.00088

f [Hz]

0.000162

0.0082

0.00146

101 102 103 104

0.00038

f [Hz]

Fig. 7: Contours of 2 f PSD −⋅ ′′ U/)vu( ∞ , Low FSTI cases; see Fig. 5 caption for further explanation

 Comparing Fig. 8 to Fig. 5, the contour levels in the high FSTI case 

are about 2 orders of magnitude higher than in the low FSTI case at 

Stations 7 and 8. This is expected and due to the higher u′ caused by 

free-stream buffeting in the high FSTI case. Farther downstream, 

however, after the higher frequencies emerge, the contour levels are 

higher for the low FSTI cases. This is somewhat misleading, as the 

rms u′ levels are actually very similar in the high and low FSTI cases 

at these stations. The lower contour levels in the high FSTI case result 

because the u′ fluctuations are distributed over a wider frequency 

band. The spectral peaks of Fig. 8 are much broader than the 

corresponding peaks of the low FSTI cases shown in Figs. 5-7. Figure 

9 shows the u′ spectra at Station 10 of the high FSTI, Re=100,000 

case. Comparing to Fig. 6, the high FSTI case exhibits much more 

low frequency activity due to free-stream buffeting, and the peak is 

clearly broader in frequency than the 273 Hz spike of the low FSTI

case. 

 Figure 10 shows contours of the turbulent shear stress spectra for 

the high FSTI cases. Values at Station 7 and 8 are low, in spite of the 

high FSTI, indicating that much of the u′ unsteadiness observed in 

Fig. 8 does not involve turbulent transport. Distinct, broadband peaks 

emerge by Station 9 in all cases. As in the low FSTI cases, these peaks 

are visible in − vu ′′ farther upstream than they are in u′ . The peak 

becomes visible in − vu ′′ soon after it forms, but the u′ peak must 

grow to become larger than the free-stream induced fluctuations before 

it is discernable. 

 The broad peaks in the high FSTI cases might suggest that transition 

occurs through a bypass mode. This was the conclusion of Volino and 

Hultgren [10]. Closer inspection, however, reveals strong similarity to 

the low FSTI cases. The frequencies of the peaks in the high FSTI

cases are listed in Table 3. The broad nature of the high FSTI case 

peaks, as shown in Fig. 9, result in an uncertainty of about 20% in the 

frequency values in Table 3. With this uncertainty, the frequency at 

each Re is essentially the same as in the corresponding low FSTI case. 

Comparison to the TS frequencies in Table 2 show the same good 

agreement observed in the low FSTI cases. Close inspection of Fig. 10 

reveals that at the four highest Re, the peak in − vu ′′ begins to appear, 

with very low magnitude, at Stations 7 and 8. The boundary layer is 

still thin at these stations, so part of the peak lies closer to the wall 

NASA/CR—2012-217656 93
8 Copyright © 2002 by ASME 

0.0084

St. 10

0.0066

St. 11

0

0.02

0.04

0.00168

St. 7

(a)

y/Ls

0.004

St. 8

0.0072

St. 9

101 102 103 104

0

0.02

0.04

0.00098

(e)

y/Ls

f [Hz]

0

0.02

0.04

0.00084

(d)

y/Ls

0

0.02

0.04

0.00102

(c)

y/Ls

0

0.02

0.04

0.00134

(b)

y/Ls

0.0044

0.0034

0.0024

101 102 103 104

0.00196

f [Hz]

0.0066

0.0042

0.0026

101 102 103 104

0.0028

f [Hz]

0.0054

0.0028

0.00148

101 102 103 104

0.00096

f [Hz]

0.0036

0.0022

0.00114

101 102 103 104

0.0008

f [Hz]

Fig. 8: Contours of 22 f ⋅PSD ′ U/)u( ∞ , High FSTI cases; see Fig. 5 caption for further explanation

than can be measured with the cross-wire probe. The closest 

measurement to the wall with the cross-wire was at y=1 mm, which 

corresponds to y/δ99.5 between 0.2 and 0.65 at the stations in question. 

What is visible, however, is at the same frequencies as present 

downstream. Since the boundary layer is still attached at Station 7, it 

is doubtful these fluctuations could be induced by a Kelvin-Helmholtz 

type instability. Instead, it appears that a TS instability plays a role, 

even in the high FSTI cases. This agrees with the findings of Hughes 

and Walker [6], who used instantaneous wall shear measurements to 

identify wave packets both within and between wakes in an unsteady 

flow. Although it appears that TS waves play a role in the high FSTI 

cases, it should be noted that it is also possible that transition occurs 

though a bypass mode, and that the spectral peaks are due not to TS 

waves but to the turbulence within turbulent spots. Perhaps both TS 

and bypass modes play a role, as suggested by Mayle [2]. 

CONCLUSIONS 

 Turbulence spectra provide valuable insights into the transition 

mechanism. Clear sharp peaks were observed in the spectra of the low 

FSTI cases, at frequencies indicating a probable TS instability 

Fig. 9: Spectra of 22 f ⋅PSD ′ U/)u( ∞ , High FSTI, Re=100,000, 

Station 10

NASA/CR—2012-217656 94
9 Copyright © 2002 by ASME 

0.00042

St. 10

0.00082

St. 11

0

0.02

0.04

9.4e−06

St. 7

(a)

y/Ls

1.62e−05

St. 8

6e−05

St. 9

101 102 103 104

0

0.02

0.04

4e−06

(e)

y/Ls

f [Hz]

0

0.02

0.04

8.4e−06

(d)

y/Ls

0

0.02

0.04

1.04e−05

(c)

y/Ls

0

0.02

0.04

5.4e−06

(b)

y/Ls

2.8e−05

3e−05

4.4e−05

101 102 103 104

3.8e−05

f [Hz]

0.000104

0.00032

0.00044

101 102 103 104

0.00034

f [Hz]

0.00072

0.00086

0.0004

101 102 103 104

0.00032

f [Hz]

0.00092

0.00058

0.0003

101 102 103 104

0.00022

f [Hz]

Fig. 10: Contours of 2 f PSD −⋅ ′′ U/)vu( ∞ , High FSTI cases; see Fig. 5 caption for further explanation

mechanism for the breakdown of the shear layer over the separation 

bubble to turbulence. The spectra were more broadband for the high 

FSTI cases, but the peaks of these spectra occurred at the same 

frequencies as in the corresponding low FSTI cases, suggesting a 

possible similar transition mechanism at high and low FSTI. The 

turbulent shear stress spectra were valuable for detection of 

instabilities upstream of the location where peaks became discernable 

in the u′ spectra. In the high FSTI cases, low magnitude peaks were 

detected in − vu ′′ upstream of separation, further supporting the 

argument for a TS transition. Larger initial disturbances in the high 

FSTI cases resulted in detectable − vu ′′ farther upstream than in the 

low FSTI cases. Since the adverse pressure gradient boundary layer is 

unstable to disturbances over a broad range of frequencies and the 

free-stream turbulence contains a range of scales, broad peaks emerge 

in the boundary layer spectra of the high FSTI cases. These peaks 

contrast with the sharp peaks of the low FSTI cases, which result from 

the growth of small disturbances at the most unstable frequencies. The 

presence and predictability of the shear layer instability may prove 

useful in future attempts to induce transition for separation control. 

ACKNOWLEDGEMENTS 

 This work was sponsored by the NASA Glenn Research Center. 

The grant monitor is Dr. David Ashpis. Additional matching support 

was provided through a U.S. Naval Academy Recognition Grant. 

REFERENCES 

[1] Volino, R.J., 2002, “Separated Flow Transition Under LowPressure Turbine Airfoil Conditions: Part 1 – Mean Flow and 

Turbulence Statistics,” ASME Paper GT-2002-30236. 

[2] Mayle, R.E., 1991, “The Role of Laminar-Turbulent Transition in 

Gas Turbine Engines,” ASME Journal of Turbomachinery, 116, 

pp. 509-537. 

[3] Schlichting, H., 1979, Boundary Layer Theory, 7th ed., McGrawHill, New York. 

[4] Sohn, K.H., and Reshotko, E., 1991, “Experimental Study of 

Boundary Layer Transition with Elevated Freestream Turbulence 

on a Heated Flat Plate,” NASA CR 187068. 

NASA/CR—2012-217656 95
10 Copyright © 2002 by ASME 

[5] Volino, R.J., 2002, “An Investigation of the Scales in Transitional 

Boundary Layers Under High Free-Stream Turbulence 

Conditions,” ASME Paper GT-2002-30233. 

[6] Hughes, J.D. and Walker, G.J., 2001, “Natural Transition 

Phenomena on an Axial Compressor Blade,” ASME Journal of 

Turbomachinery, 123, pp. 392-401. 

[7] Halstead, D.E., Wisler, D.C., Okiishi, T.H., Walker, G.J., 

Hodson, H.P., and Shin, H.-W., 1997, “Boundary Layer 

Development in Axial Compressors and Turbines: Part 3 of 4 –

LP Turbines,” ASME Journal of Turbomachinery, 119, pp. 225-

237. 

[8] Solomon, W.J., and Walker, G.J., 1995, “Incidence Effects on 

Wake-Induced Transition on an Axial Compressor Blade,” Proc. 

12th International Symposium on Air Breathing Engines, 

Melbourne, Australia, pp. 954-964. 

[9] Hatman, A., and Wang, T., 1999, “A Prediction Model for 

Separated Flow Transition,” ASME Journal of Turbomachinery, 

121, pp. 594-602. 

[10] Volino, R.J., and Hultgren, L.S., 2001, “Measurements in 

Separated and Transitional Boundary Layers Under Low-Pressure 

Turbine Airfoil Conditions,” ASME Journal of Turbomachinery, 

123, pp. 189-197. 

[11] Lou, W., and Hourmouziadis, J., 2000, “Separation Bubbles 

Under Steady and Periodic-Unsteady Main Flow Conditions,”

ASME Journal of Turbomachinery, 122, pp. 634-643. 

[12] Ligrani, P.M., and Bradshaw, P., 1987, “Spatial Resolution and 

Measurement of Turbulence in the Viscous Sublayer Using 

Subminiature Hot-Wire Probes,” Experiments in Fluids, 5, pp. 

407-417. 

[13] Ligrani, P.M., and Bradshaw, P., 1987, “Subminiature Hot-Wire 

Sensors: Development and Use,” Journal of Physics E: Scientific 

Instruments, 20, pp. 323-332. 

[14] Ligrani, P.M., Westphal, R.V., and Lemos, F.R., 1989, 

“Fabrication and Testing of Subminiature Multi-Sensor Hot-Wire 

Probes,” Journal of Physics E: Scientific Instruments, 22, pp. 

262-268. 

[15] Walker, G.J., 1989, “Transitional Flow on Axial Turbomachine 

Blading,” AIAA Journal, 27, pp. 595-602. 

NASA/CR—2012-217656 96
 1 Copyright © 2003 by ASME 

Proceedings of ASME Turbo Expo 2003 

Power for Land, Sea, and Air 

June 16-19, 2003, Atlanta, Georgia, USA

GT2003-38728 

PASSIVE FLOW CONTROL ON LOW-PRESSURE TURBINE AIRFOILS

Ralph J. Volino

Department of Mechanical Engineering 

United States Naval Academy 

Annapolis, Maryland 21402 

Email: volino@usna.edu

ABSTRACT 

 Two-dimensional rectangular bars have been used in an 

experimental study to control boundary layer transition and 

reattachment under low-pressure turbine conditions. Cases with 

Reynolds numbers (Re) ranging from 25,000 to 300,000 (based on 

suction surface length and exit velocity) have been considered at low 

(0.5%) and high (8.5% inlet) free-stream turbulence levels. Three 

different bars were considered, with heights ranging from 0.2% to 

0.7% of suction surface length. Mean and fluctuating velocity and 

intermittency profiles are presented and compared to results of 

baseline cases from a previous study. Bar performance depends on the 

bar height and the location of the bar trailing edge. Bars located near 

the suction surface velocity maximum are most effective. Large bars 

trip the boundary layer to turbulent and prevent separation, but create 

unnecessarily high losses. Somewhat smaller bars had no immediate 

detectable effect on the boundary layer, but introduced small 

disturbances which caused transition and reattachment to move 

upstream from their locations in the corresponding baseline case. The 

smaller bars were effective under both high and low free-stream 

turbulence conditions, indicating that the high free-stream turbulence 

transition is not simply a bypass transition induced by the free-stream. 

Losses appear to be minimized when a small separation bubble is 

present, so long as reattachment begins far enough upstream for the 

boundary layer to recover from the separation. Correlations for 

determining optimal bar height are presented. The bars appear to 

provide a simple and effective means of passive flow control. Bars 

which are large enough to induce reattachment at low Re, however, 

cause higher losses at the highest Re. Some compromise would, 

therefore, be needed when choosing a bar height for best overall 

performance. 

NOMENCLATURE 

Cp ( ) 2

T − ρU/PP2 e , pressure coefficient 

d bar height 

FSTI free-stream turbulence intensity 

H δ

∗

/θ, shape factor 

K ) /)(/( dsdUU 2 ν ∞∞ , acceleration parameter 

Ls suction surface length 

P pressure 

PT upstream stagnation pressure

Re LU se /ν , exit Reynolds number 

Red Udd/ν, Reynolds number based on bar height 

Rest U∞(st-ss)/ν, separation to transition distance Reynolds number

Reθ momentum thickness Reynolds number 

s streamwise coordinate, distance from leading edge 

U mean streamwise velocity 

U∞ local free-stream velocity 

Ud mean velocity at bar height in baseline boundary layer 

Ue nominal exit free-stream velocity, based on inviscid solution 

u′ rms streamwise fluctuating velocity 

y cross-stream coordinate, distance from wall 

δ99.5 99.5% boundary layer thickness 

δ

∗ displacement thickness 

γ intermittency, fraction of time flow is turbulent 

ν kinematic viscosity 

ρ density 

θ momentum thickness 

Subscripts 

b baseline flow 

m modified flow 

p suction surface pressure minimum, velocity maximum 

s separation location 

t transition start location 

INTRODUCTION 

 Modern low-pressure turbine (LPT) airfoils are subject to 

increasingly stronger pressure gradients as designers impose higher 

loading in an effort to improve efficiency and lower cost by reducing 

the number of airfoils in an engine. If the adverse pressure gradient on 

the suction side of these airfoils becomes strong enough, the boundary 

layer will separate. Separation bubbles, particularly those which fail 

to reattach, can result in a significant loss of lift and a subsequent 

degradation of engine efficiency (e.g. Hourmouziadis [1], Mayle [2], 

NASA/CR—2012-217656 97
 2 Copyright © 2003 by ASME 

and Sharma et al. [3]). The problem is particularly relevant in aircraft 

engines. Airfoils optimized to produce maximum power under takeoff 

conditions may still experience boundary layer separation at cruise 

conditions, due to the thinner air and lower Reynolds numbers at 

altitude. A component efficiency drop of 2% may occur between 

takeoff and cruise conditions in large commercial transport engines, 

and the difference could be as large as 7% in smaller engines operating 

at higher altitudes. Component life may also be affected by more than 

an order of magnitude (Hodson [4]). Because the LPT produces the 

bulk of the net power in many engines, changes in its component 

efficiency can result in nearly equal changes in overall engine 

efficiency (Wisler [5]). There are several sources for losses in an 

engine, including secondary flows, but the suction side boundary layer 

has been identified as the primary source of losses in the LPT (Curtis 

et al. [6]). Prediction and control of suction side separation, without 

sacrifice of the benefits of higher loading, is therefore, necessary for 

improved engine design. 

 Separation on LPT airfoils is complicated by boundary layer 

transition. Turbulent boundary layers are much more resistant to 

separation than laminar boundary layers. If transition occurs far 

enough upstream, it can prevent separation. If transition occurs in the 

shear layer over a separation bubble, it will tend to induce boundary 

layer reattachment. The lower the Reynolds number, the farther 

downstream transition will tend to occur, hence the problems 

associated with performance at altitude. 

 Separated flow transition has been studied extensively, and in recent 

years several studies have focused on transition in the LPT. Volino [7] 

provides a review of much of that work. Separation can be affected 

through naturally occurring phenomena in an engine and through 

deliberate attempts at flow control. Several studies have shown that 

high free-stream turbulence intensity (FSTI) tends to cause transition 

to move upstream, resulting in a smaller separation bubble. Reducing 

the separation bubble size tends to result in thinner boundary layers 

after reattachment, thereby reducing losses. Moving transition 

upstream, however, results in a longer turbulent region on the airfoil, 

which tends to increase losses. Volino [7] showed that the net result 

of these competing effects depends on the Reynolds number. High 

FSTI tends to reduce losses at low Re. At high Re, where separation 

bubbles are relatively small even with low FSTI, high FSTI results in 

higher losses. At very low Re, boundary layers may fail to reattach 

even with high FSTI (e.g. Volino [7] and Van Treuren et al. [8]). 

Unsteadiness caused by wakes generated upstream of an airfoil has 

been shown in several studies (e.g. Howell et al. [9]) to reduce the 

extent of separation bubbles and reduce losses. As with elevated FSTI, 

wake unsteadiness is most effective at reducing losses at lower Re, 

where the steady flow separation bubbles are largest. Stadtmüller et 

al. [10] found that at high Re, losses were higher with wakes than in 

steady flow. 

 Existing results suggest that separation bubbles should be kept 

small, but without producing an unnecessarily long turbulent region. 

Hourmouziadis [1] discussed “controlled diffusion blading,” in which 

an airfoil is designed so that a small separation bubble is present. The 

bubble itself is not thick enough to produce high losses, and its 

presence allows a shorter turbulent region near the trailing edge. This 

idea is discussed below in conjunction with the present results. 

 Capitalizing on the beneficial effects of unsteady wakes, Howell et 

al. [9] and Brunner et al. [11] studied airfoils modified for higher lift. 

Losses increased with airfoil loading, as adverse pressure gradients 

became stronger and separation bubbles became larger. With wake 

passing, however, the magnitude of the loss increase was in some 

cases relatively small compared to the increase in lift. Aft loaded 

airfoils tended to have lower losses, since separation and transition 

occurred closer to the trailing edge, resulting in a shorter turbulent 

region. 

 While high FSTI and wakes help to mitigate separated flow 

problems, they clearly do not solve all problems, as evidenced by the 

known efficiency drop in modern engines at altitude. Howell et al. [9] 

indicated that their highly loaded airfoils might be close to a limit, and 

that higher loading could cause unacceptable separation problems even 

in the presence of wakes. Looking beyond FSTI and wakes, other 

types of flow control could prove useful. The literature contains 

numerous examples of separation control. Most have been applied to 

external flows over aircraft, but a few studies have considered passive 

devices added to LPT airfoils. Van Treuren et al. [8] utilized vortex 

generators on the suction surface of an LPT airfoil. The vortex 

generators caused reattachment at Re=50,000 (all Re in the present 

paper are based on exit velocity and suction surface length). Losses 

appeared to be slightly lower with the vortex generators. The vortex 

generators were not effective at Re=25,000, and the boundary layer did 

not reattach even with 8% FSTI. Van Treuren et al [8] did not 

consider higher Re. In another study, Lake et al. [12] used various 

passive devices including dimples and boundary layer trips in an LPT 

cascade. They considered cases with Re above 100,000. Murawski 

and Vafai [13] added extensions to the trailing edges of the airfoils in 

their cascade. These extensions tended to move the separation 

location downstream. At low Re, they reduced the length of the 

separation bubble and reduced losses. At high Re, losses increased. 

Byerley et al. [14] used “Gurney flaps” to control separation. These 

devices were trips, near the trailing edge on the pressure side of the 

airfoils. They helped to keep the boundary layer attached on the 

suction side, but also increased losses in the cascade. Active 

separation control has also been employed. Bons et al. [15, 16] used 

steady and pulsed vortex generator jets to successfully control 

separation under LPT conditions. 

 The studies listed above indicate that separation control should be 

possible under LPT conditions. Existing results are, however, limited 

both in the range of Reynolds numbers considered in each study and in 

the types of data acquired. More experiments are needed with various 

types of devices to expand the experimental data base. Detailed 

measurements will also help in the explanation of the physical 

mechanisms by which various devices affect the flow. 

 Passive flow control is considered in the present work. Thin bars of 

rectangular cross section are placed on the suction surface of an LPT 

airfoil near the suction surface velocity peak. Experiments were 

conducted in a single-passage cascade-simulator, described in Volino 

[7]. The geometry of the passage corresponds to that of the “Pak-B” 

airfoil, which is an industry supplied research airfoil that is 

representative of a modern, aggressive LPT design. Volino [7] 

documented cases in the present facility without flow control. These 

serve as baseline cases for the present study. 

EXPERIMENTS

 Experiments were conducted in a low speed wind tunnel, described 

by Volino et al. [17]. Briefly, air enters through blowers and passes 

through a honeycomb, a series of screens, two settling chambers, and a 

three-dimensional contraction before entering the test section. At the 

exit of the contraction, the mean velocity is uniform to within 1%. 

The FSTI is 0.5%±0.05%. Nearly all of this free-stream “turbulence” 

is actually streamwise unsteadiness at frequencies below 20 Hz and is 

not associated with turbulent eddies. The rms intensities of the three 

components of the unsteadiness are 0.7%, 0.2% and 0.2% in the 

streamwise, pitchwise and spanwise directions, respectively. For low 

FSTI cases, the test section immediately follows the contraction. For 

high FSTI, a passive grid is installed at the contraction exit followed 

NASA/CR—2012-217656 98
 3 Copyright © 2003 by ASME 

by a 1 m long rectangular settling chamber. At the inlet to the test 

section the high FSTI mean flow and turbulence are spatially uniform 

to within 3% and 6% respectively. The free-stream turbulence is 

nearly isotropic with rms intensities of 8.8%, 8.9% and 8.3% in the 

streamise, pitchwise and spanwise directions. The integral length 

scales of these components are 3 cm, 1.6 cm and 1.4 cm. The integral 

scales were computed from the power spectra of each component. 

 The test section, shown in Fig. 1, consists of the passage between 

two airfoils. Details are listed in Table 1 and more information is 

available in Volino [7]. A large span to chord ratio of 4.3 was chosen 

to insure two-dimensional flow at the spanwise centerline of the 

airfoils, where all measurements were made. Upstream of each airfoil 

are flaps, which control the amount of bleed air allowed to escape 

from the passage. The flaps, along with a tailboard on the pressure 

side of the passage, are adjusted to produce the correct leading edge 

flow and pressure gradient along the airfoils. The flow in the passage 

matches that in a multi-blade cascade. 

 Experimental conditions match those of the ten baseline cases of 

Volino [7], who considered high and low FSTI cases at five Reynolds 

numbers (Re=25,000, 50,000, 100,000, 200,000, and 300,000). The 

Reynolds number range is representative of conditions from cruise to 

takeoff. The FSTI levels in an engine may vary considerably, but the 

values in the present work are believed to span the range of most 

interest. 

 Prior to the detailed experiments of the present study, various 

devices were used in preliminary attempts at flow control. The 

devices included trip wires of various diameters, rectangular bars of 

various widths and thicknesses, and delta wing vortex generators of 

various heights, spacing, and angles with respect to the flow. All of 

these devices were tried at several streamwise locations along the 

suction surface. Documentation included streamwise pressure profiles 

and velocity profiles acquired near the trailing edge. Large devices of 

any type eliminated separation (as indicated by the pressure profiles), 

but caused large increases in losses (as indicated by large increases in 

momentum deficit at the trailing edge). As the size of the devices was 

reduced, it was found that all devices which were just large enough to 

induce boundary layer reattachment at Re=25,000 caused about the 

same increase in losses at higher Re. This was somewhat unexpected, 

as it was thought that the delta wings vortex generators might present 

less blockage, and more effectively promote mixing and inhibit 

separation than the trips or bars. Reasons for this unexpected finding 

are discussed with the results below. Since no device appeared to have 

a clear advantage, rectangular bars were chosen for further study 

because of their simplicity. The bars were of uniform rectangular 

cross section and extended along the airfoil span, as shown in Fig. 2. 

It should be noted that the tests described above were not exhaustive, 

and do not preclude the possible usefulness of vortex generators or 

other types of devices. 

 The results of the preliminary tests with the bars indicated that the 

streamwise width of a bar and the location of its leading edge were 

unimportant. The bar height and the location of its trailing edge were 

critical. Hence, it appeared that the backward facing step at the 

trailing edge was most important for flow control. Bars were most 

effective when the trailing edge was near the location of the suction 

surface velocity peak. If the trailing edge was much farther 

downstream, it was located under the separation bubble and was 

ineffective. If the trailing edge was upstream in the favorable pressure 

gradient region, the stabilizing effect of the accelerating flow appeared 

to lessen the bar’s effectiveness. 

 In the present study, rectangular bars were fabricated from multiple 

layers of vinyl tape. The trailing edge of the bar was located at 

s/Ls=0.51, near the suction surface velocity peak. All bars were 6 mm 

Table 1: Test section parameters 

Axial 

Chord 

[mm] 

True 

Chord 

[mm] 

Pitch 

[mm] 

Span 

[mm] 

Suction 

side, Ls

[mm] 

Inlet 

flow 

angle 

Exit 

flow 

angle 

153.6 170.4 136.0 660.4 228.6 35° 60° 

tailboard

flap

flap bleed air

bleed air

35o

60o

Fig. 1 Schematic of the test section 

Fig. 2 Scale drawing of suction side airfoil showing 

location of bar 

wide in the streamwise direction. Bar heights of 0.4 mm, 0.8 mm and 

1.6 mm were used. The bar heights were all less than 1% of Ls. They 

compare to local boundary layer thickness at the bar location of about 

3.8 mm, 2.7 mm, 2.0 mm, 1.4 mm and 1.2 mm in the baseline 

Re=25,000 through 300,000 cases respectively. For each bar height, 

all 10 cases of the baseline study were re-documented, for a total of 30 

new experimental cases. 

Measurements 

 Pressure surveys were made for each case using a pressure 

transducer (0-870 Pa range Validyne transducer) and a Scanivalve. 

Stagnation pressure was measured with a pitot tube upstream of the 

passage inlet, and eleven pressure taps were located on each airfoil 

along their spanwise centerlines. Locations of the taps on the suction 

side are listed in Table 2 along with measured local FSTI components, 

and the ReK product at these stations based on a non-separating, 

inviscid solution. The uncertainty in the suction side pressure 

coefficients was 7% at the lowest Re, and below 4% in other cases. 

Most of this uncertainty was due to bias error. Stochastic error was 

minimized by averaging pressure transducer readings over a 10 second 

period. 

NASA/CR—2012-217656 99
 4 Copyright © 2003 by ASME 

Table 2: Measurement stations locations, local acceleration 

(inviscid soln.), and measured local free-stream turbulence 

Station s/Ls ReK Low 

FSTI

′/Uu ∞

 [%] 

Low 

FSTI

′/Uv ∞

 [%] 

High 

FSTI

′/Uu ∞

 [%] 

High 

FSTI

′/Uv ∞

 [%] 

1 0.111 1.58 0.44 5.2 

2 0.194 1.20 0.39 4.6 

3 0.278 0.86 0.37 4.0 

4 0.361 0.75 0.38 3.5 

5 0.444 0.62 0.39 3.2 

6 0.528 -0.02 0.41 2.8 

7 0.611 -0.81 0.47 0.05 2.9 5.9 

8 0.694 -0.95 0.47 0.12 3.0 6.2 

9 0.777 -0.58 0.48 0.14 3.4 6.6 

10 0.861 -0.53 0.54 0.11 3.8 6.8 

11 0.944 -0.18 0.51 0.11 4.0 6.8 

 Velocity profiles on the suction surface were measured at 

streamwise stations corresponding to pressure taps 7-11, as given in 

Table 2. These stations are downstream of the bar. Profiles at Stations 

1-6 are fully documented for the baseline cases in Volino [7, 18], and 

show that the upstream boundary layer closely follows a laminar 

solution, even in the high FSTI cases. Profiles were measured near but 

not at the spanwise centerline of the airfoil to insure that the pressure 

taps did not interfere with the velocity measurements. Profiles were 

acquired with a hot-wire anemometer (AA Lab Systems model AN1003) and a single sensor boundary layer probe (TSI model 1218-

T1.5). The sensor diameter is 3.8 μm, and the active length is 1.27 

mm. At each measurement location, data were acquired for 26 

seconds at a 20 kHz sampling rate (219 samples). All raw data were 

saved. The high sampling rate provides an essentially continuous 

signal, which is needed for intermittency and spectral post-processing. 

The long sampling time results in low uncertainty in both statistical 

and spectral quantities. Data were acquired at 60 wall normal 

locations in each profile, extending from the wall to the free-stream, 

with most points concentrated in the near wall region. The closest 

point was within 0.1 mm of the wall, which corresponds to 

y/Ls=0.0004 and between 0.01 and 0.2 boundary layer thicknesses. 

Flow direction in a separation bubble cannot be determined with a 

single-sensor hot-wire, but velocity magnitude can be measured and 

was found to be essentially zero within the bubbles of the present 

cases. Determining the direction was not, therefore, considered 

essential. Uncertainties in the mean velocity are 3-5% except in the 

very near wall region where near-wall corrections (Wills [19]) were 

applied to the mean velocity. Uncertainties in the momentum and 

displacement thicknesses computed from the mean profiles are 10%. 

Uncertainty in the shape factor, H, is 8%. 

 The uncertainty in the fluctuating streamwise velocity is below 

10%, except in the very near wall region, where spatial averaging 

effects, due to the finite length of the hot-wire sensor, become 

important in some cases. For the present cases, as explained in Volino 

[7] based on the work of Ligrani and Bradshaw [20, 21], spatial 

averaging should not be significant for the Re=25,000 and 50,000 

cases, even near the wall. For the higher Re cases, spatial averaging 

should not be significant for y>1 mm (y/Ls>0.004), but may cause 

errors as high as 30% closer to the wall. It is not certain that the errors 

are this large, however. The estimates are based on the results of 

Ligrani and Bradshaw [20, 21], who considered a boundary layer with 

Reθ=2600. The momentum thickness Reynolds numbers in the present 

cases are all below 1300. This may indicate less developed turbulence 

in the present study, which could imply fewer small scale eddies and 

lower averaging errors. 

 The intermittency, γ, is the fraction of time the flow is turbulent 

within the transition region. It was determined at each measurement 

location based on the instantaneous streamwise velocity signal, using 

the technique described in Volino et al. [17]. The uncertainty in γ is 

10%. As explained in Volino [17], turbulent flow is defined here to 

include a range of large and small scale eddies, turbulence production, 

and dissipation. A boundary layer may be characterized by significant 

u′ fluctuations but still be non-turbulent if these fluctuations are 

induced by an external source that does not also cause near wall 

turbulence production. Such is often the case under high FSTI

conditions. Free-stream eddies buffet the boundary layer, inducing 

non-turbulent boundary layer fluctuations but very little momentum 

transport. Transition to turbulence is characterized not so much by 

large increases in u′ levels, which may remain essentially constant, 

but by the appearance of higher frequencies. The higher frequencies 

signal the generation of turbulence in the near wall region and are used 

to distinguish between turbulent and non-turbulent flow. Further 

discussion is available in Volino [17]. 

RESULTS 

Velocity and Pressure Profiles

 Pressure coefficients for the low FSTI, Re=25,000 cases are shown 

in Fig. 3. Also shown is the inviscid solution for the present geometry. 

In all cases there is good agreement with the inviscid solution on the 

pressure side. On the suction side, the baseline case shows good 

agreement with the inviscid solution in the favorable pressure gradient 

region, but a large separation bubble in the adverse pressure gradient 

region. Separation is indicated by the nearly constant Cp values, 

which are well above the invisid solution. The Cp values remain high 

to the trailing edge, showing no sign of reattachment. With the 0.4 

mm thick bar, there is an increase in Cp over the baseline value at 

s/Ls=0.53. The pressure tap at this location is immediately 

downstream of the bar, and the flow over the tap is probably affected 

by the close proximity of the bar. The Cp values in this case remain 

high to the trailing edge, indicating that the boundary layer does not 

reattach. The same is true for the 0.8 mm bar case. With the 1.6 mm 

bar, Cp drops below the baseline values near the end of the favorable 

pressure gradient region. The larger bar is apparently enough of an 

obstruction to slow the near wall flow upstream of the bar. 

Downstream of the 1.6 mm bar, Cp values are high, as in the other 

cases, but at the most downstream pressure tap Cp drops to near the 

inviscid solution value, indicating boundary layer reattachment. 

 The velocity profiles for the low FSTI, Re=25,000 cases are shown 

in Fig. 4. The top row of the figure shows dimensionless mean 

velocity profiles at Stations 7-11. The baseline case shows a boundary 

layer near separation at Station 7, a small separation bubble at Station 

8, and an increasingly larger bubble at Stations 9-11. The mean 

profiles of the 0.4 mm and 0.8 mm bar cases are virtually 

indistinguishable from the baseline case. The 1.6 mm bar case shows 

a clear separation bubble at Station 7, suggesting that the boundary 

layer has separated from the trailing edge of the bar. This bubble 

continues to grow, and at Station 8 the separation bubble is larger with 

the 1.6 mm bar than in the other cases. At Station 9 all cases appear 

similar. This agreement is really a crossing, as the shear layer in the 

1.6 mm bar case is on the verge of reattaching, while the separation 

bubble is growing in the other cases. At Station 10, the near wall 

velocity in the 1.6 mm bar case has begun to rise, indicating incipient 

reattachment. By Station 11, the boundary layer has clearly reattached 

in the 1.6 mm bar case, although the mean profile has not recovered to 

NASA/CR—2012-217656 100
 5 Copyright © 2003 by ASME 

0 0.2 0.4 0.6 0.8 1 1.2 0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

baseline

s/Ls

Cp

0.4 mm bar

0.8 mm bar

1.6 mm bar

inviscid solution

Low FSTI

Re=25,000

Fig. 3 Cp profiles, Low FSTI, Re=25,000 cases 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 4 Station 7-11 profiles, low FSTI, Re=25,000 cases: (a) 

mean velocity, (b) u′, (c) intermittency 

a fully-developed turbulent shape. Dimensionless u′ profiles are 

shown in the second row of Fig. 4. As with the mean profiles, the 

baseline case and the 0.4 mm and 0.8 mm bar bases are 

indistinguishable through Station 10. All show a small u′ peak 

growing in the shear layer over the separation bubble. As explained in 

Volino [7, 18], this peak is not indicative of transition. It is caused 

primarily by low frequency fluctuations which are amplified when 

they act across the region of high mean velocity gradient in the shear 

layer. At Station 11, there is a slight increase in u′ in the 0.8 mm bar 

case over the baseline case, and the high values extend into the near 

wall region. These near wall fluctuations suggest the beginning of 

transition and reattachment, but they are not large enough in this case 

to significantly affect the mean profile. The 1.6 mm bar case shows a 

larger u′ peak than the other cases at Stations 7 and 8. The peak is in 

the shear layer and is similar to the peaks at Stations 9 and 10 of the 

other cases. It does not indicate transition. The peak becomes larger 

at Station 9, and extends into the near wall region, which is a sign of 

incipient transition. By Station 10 u′ is much larger with a clear 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 5 Station 7-11 profiles, high FSTI, Re=25,000 cases: (a) 

mean velocity, (b) u′, (c) intermittency 

double peak. This rise in u′ corresponds to the beginning of 

reattachment observed in the mean profile. The third row of Fig. 4 

shows the local intermittency. Its is zero in all cases through Station 9, 

but begins to rise at Stations 10 and 11 of the 1.6 mm bar case. The 

intermittency peak is in the shear layer, indicating that this is where 

transition begins. Intermittency only reaches about 13% at Station 11. 

This is consistent with the mean velocity profile, which shows the 

boundary layer is reattached but not yet a fully-developed turbulent 

profile. As the turbulence is intermittent, it is likely that the boundary 

layer is only intermittently reattached. The high u′ peak at Station 11 

is also consistent with a transitional boundary layer. As a boundary 

layer becomes fully turbulent, the dimensionless u′ peak will decrease 

in magnitude to about 0.1, and move close to the wall. 

 The behavior in the 1.6 mm bar case is interesting. The bar was not 

large enough to immediately trip the boundary layer to turbulent, but it 

did move the separation point upstream. This caused transition to 

move upstream, and led to at least a partial reattachment by the trailing 

edge, which did not occur in the other cases. The 0.8 mm bar trip case 

is also very interesting. The bar in this case was so small that it had no 

immediate measurable affect on the mean or u′ profiles. Well 

downstream at Station 11, however, the effect of this bar became 

visible in the u′ profile. Apparently this bar introduced a very small 

disturbance in the flow, which was too small to detect at first, but grew 

as it moved downstream. 

 The velocity profiles of Fig. 4 and the pressure profiles of Fig. 3 are 

in good agreement. Both show transition and reattachment at the same 

locations, and the measured static pressures agree with the local freestream velocities of Fig. 4. The agreement between the pressure and 

velocity results was apparent in all cases. For brevity, the pressure 

profiles are not presented in the cases which follow. 

 Figure 5 shows the velocity profiles for the high FSTI, Re=25,000 

cases. The format is that same as in Fig. 4. As in the low FSTI case, 

the baseline, 0.4 mm trip and 0.8 mm trip cases are nearly 

indistinguishable at the upstream stations. By Station 11, some 

differences are apparent in the mean profiles for these cases. The 

separation bubble is less distinct in the cases with the bars, but the 

boundary layer still does not appear fully reattached. Intermittency 

rises slightly above zero at Station 11 of the 0.8 mm bar case, while 

remaining essentially at zero with the smaller bar and in the baseline 

case. The u′ profiles show a large peak in the shear layer which 

grows in the streamwise direction. As shown in Volino [7], this peak 

NASA/CR—2012-217656 101
 6 Copyright © 2003 by ASME 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 6 Station 7-11 profiles, low FSTI, Re=50,000 cases: (a) 

mean velocity, (b) u′, (c) intermittency 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 7 Station 7-11 profiles, high FSTI, Re=50,000 cases: (a) 

mean velocity, (b) u′, (c) intermittency 

is caused by the action of the high FSTI on the shear layer, and does 

not indicate significant momentum transport. As in the low FSTI

cases of Fig. 4, the 1.6 mm bar case shows significant differences from 

the other cases in Fig. 5. The u′ peak is significantly higher in this 

case at Stations 7 and 8. At Station 9, u′ values are higher in the near 

wall region, the intermittency rises above zero, and the mean profile 

appears to be reattached. At Stations 10 and 11 the intermittency 

continues to increase. The mean profile adjusts toward a more 

turbulent shape between Stations 9 and 11. The u′ peak decreases 

somewhat by Station 11, but still shows the relatively high values of a 

transitional boundary layer, rather than the somewhat lower values of a 

fully-turbulent boundary layer. As in the low FSTI, Re=25,000 case, 

the 1.6 mm bar is not large enough to immediately trip the boundary 

layer to turbulent, but it causes transition to move upstream and leads 

to a reattachment that did not occur in the baseline or smaller bar 

cases. 

 The velocity profiles of the low FSTI, Re=50,000 case are shown in 

Fig. 6. The effects of the bars are clear. At Station 7, the 1.6 mm bar 

has caused a relatively large separation bubble compared to the other 

cases and a small u′ peak in the shear layer over this bubble. The 

smaller bar cases are indistinguishable from the baseline case, with 

mean profiles only on the verge of separation and u′ near zero. By 

Station 8, the 1.6 mm bar case has undergone a sudden transition, with 

γ=1, high u′ levels in both the shear layer and near wall regions, and 

the beginning of reattachment as indicated by non-zero mean velocity 

near the wall. The other cases are indistinguishable, exhibiting a small 

separation bubble and showing no sign of transition or reattachment. 

At Station 9, the boundary layer is clearly reattached in the 1.6 mm bar 

case, and u′ values are beginning to rise in the 0.8 mm bar case. At 

Station 10 the intermittency indicates fully-turbulent flow and the 

boundary layer has reattached in the 0.8 mm bar case. The 0.4 mm bar 

case is still separated with γ=0, but u′ has begun to rise near the wall. 

By Station 11 the 0.4 mm bar case has become turbulent and the 

boundary layer has started to reattach. The baseline case remains nonturbulent with a large separation bubble at Station 11. As observed in 

the Re=25,000 cases, the 1.6 mm bar is not large enough to 

immediately trip the boundary layer to turbulent, but it does move 

separation upstream, which causes transition and reattachment to move 

significantly upstream. The smaller bars appear to have no immediate 

effect on the boundary layer, but they must introduce small 

disturbances that grow in the streamwise direction and have a 

significant effect in moving transition and reattachment upstream. The 

0.8 mm bar must introduce a larger disturbance than the 0.4 mm bar, 

since transition and reattachment occur one station farther upstream 

with the 0.8 mm bar. 

 Figure 7 shows the high FSTI, Re=50,000 cases. With the larger 

bars, transition and reattachment move upstream. With the 1.6 mm 

bar, the intermittency is already non-zero by Station 7, and the 

boundary layer is fully turbulent and attached by Station 8. With the 

0.8 mm bar, γ rises above zero at Station 8 and is near fully turbulent 

by Station 10. The mean profile appears to indicate reattachment by 

Station 9. The intermittency rises above zero in the 0.4 mm bar case at 

Station 9, and continues to rise at Stations 10 and 11. The mean 

profile shows reattachment at Station 10. In the baseline case, the 

intermittency begins to rise at Station 10, and the boundary layer is 

reattached at Station 11. At Station 11 the mean profiles are 

indistinguishable in the cases with bars, and fuller than in the baseline 

case. In all cases, transition begins upstream of the location in the 

corresponding low FSTI case of Fig. 6, but the transition length is 

longer. Volino and Hultgren [22] also observed that transition begins 

farther upstream with high FSTI, but is more abrupt in low FSTI cases. 

 Figure 8 shows the low FSTI, Re=100,000 cases. The 1.6 mm bar 

immediately trips the boundary layer to turbulent and eliminates the 

separation bubble. The 0.8 mm bar causes a small u′ peak above the 

baseline values at Station 7. The intermittency jumps from 0 to 1 

between Stations 7 and 8, and the separation bubble is effectively 

eliminated. As in the lower Re cases, the 0.4 mm bar has no visible 

effect at Station 7, and the mean and u′ profiles are indistinguishable 

from the baseline case. The boundary layer separates, but by Station 9 

the shear layer has become fully turbulent and begun to reattach. In 

the baseline case, u′ does not begin to show elevated near wall values 

until Station 10, and transition and reattachment occur at Station 11. 

The mean profiles at Station 11 show the fullest profile and thinnest 

boundary layer in the 0.4 mm bar case. The larger bars result in 

thicker boundary layers. The mean profile in the baseline case has not 

yet recovered to a fully turbulent shape. As will be discussed below, 

the thinner attached boundary layer in the 0.4 mm bar case suggests 

that this case will have lower losses than the other cases. 

 The high FSTI, Re=100,000 cases are shown in Fig. 9. As in the 

low FSTI case, the 1.6 mm bar trips the boundary layer to turbulent 

and eliminates separation at this Re. The intermittency is non-zero at 

Station 7 of the 0.8 mm bar case, and it continues to rise through 

NASA/CR—2012-217656 102
 7 Copyright © 2003 by ASME 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 8 Station 7-11 profiles, low FSTI, Re=100,000 cases: (a) 

mean velocity, (b) u′, (c) intermittency 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 9 Station 7-11 profiles, high FSTI, Re=100,000 cases: 

(a) mean velocity, (b) u′, (c) intermittency 

Station 10, indicating an extended transition zone. Because transition 

begins so far upstream, the separation bubble is eliminated and 

transition occurs in an attached boundary layer. With the 0.4 mm bar, 

the intermittency indicates that transition does not begin until Station 

8, so a small separation bubble forms, as in the baseline case. The 

boundary layer is reattached by Station 9, however, and transition is 

nearly complete by Station 10. In the baseline case, transition begins 

at Station 9, and the boundary layer is reattached at Station 10. 

Examining the mean profiles, the 1.6 mm bar causes an immediate 

thickening of the boundary layer, and the separation bubble in the 

baseline case also causes a thicker boundary layer. By Station 11, the 

mean profiles for these two cases agree closely. The boundary layers 

are thinner in the cases with the smaller bars. 

 In the low FSTI, Re=200,000 cases of Fig. 10, the intermittency 

profiles show that the 0.8 mm and 1.6 mm bars immediately trip the 

boundary layer to turbulent and eliminate the separation bubble. The 

1.6 mm bar, which is the same thickness as the boundary layer at 

Station 7 of the baseline case, results in a substantially thicker 

boundary layer than in all of the other cases. The 0.4 mm bar case 

shows a small separation bubble at Station 8, but is fully turbulent and 

reattached by Station 9. The baseline case exhibits a clear separation 

bubble at Station 9, and is fully turbulent and reattached by Station 10. 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 10 Station 7-11 profiles, low FSTI, Re=200,000 cases: 

(a) mean velocity, (b) u′, (c) intermittency 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 11 Station 7-11 profiles, high FSTI, Re=200,000 cases: 

(a) mean velocity, (b) u′, (c) intermittency 

The mean profiles at Station 11 show that the growth of the bubble in 

the baseline case results in a thicker boundary layer than in the 0.8 and 

0.4 mm bar cases. 

 Figure 11 shows the high FSTI, Re=200,000 cases. As in the low 

FSTI cases of Fig. 10, the 0.8 and 1.6 mm bars trip the boundary layer 

to turbulent. Transition has already started, as indicated by the nonzero intermittency, in the 0.4 mm bar case at Station 7. In all of these 

cases, there is no separation. In the baseline case, γ does not rise 

above zero until Station 9, and there may be a small separation bubble 

at Station 8. At Station 11, the mean, u′ and intermittency profiles of 

the baseline, 0.4 mm and 0.8 mm bar cases are all in good agreement, 

while the 1.6 mm bar case exhibits a noticeably thicker boundary 

layer. 

 The low FSTI, Re=300,000 cases are shown in Fig. 12. As in the 

Re=200,000 cases, the 0.8 mm and 1.6 mm bars trip the boundary 

layer to turbulent. The 0.4 mm bar appears to have no effect at Station 

7, where the mean and u′ profiles agree with the baseline case and the 

intermittency is zero. By Station 8, however, the boundary layer in the 

0.4 mm bar case has become fully turbulent, while in the baseline case 

it is still laminar and has separated. By Station 9 the shear layer in the 

baseline case is transitional, and it is fully turbulent and reattached by 

Station 10. At Station 11 the mean profiles for the four cases are all 

NASA/CR—2012-217656 103
 8 Copyright © 2003 by ASME 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 12 Station 7-11 profiles, low FSTI, Re=300,000 cases: 

(a) mean velocity, (b) u′, (c) intermittency 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 1 0

0.05

γ

y/Ls

c) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 13 Station 7-11 profiles, high FSTI, Re=300,000 cases: 

(a) mean velocity, (b) u′, (c) intermittency 

different, with the 0.4 mm bar and baseline cases having the thinnest 

boundary layers. 

 Figure 13 shows the high FSTI, Re=300,000 cases. As in the low 

FSTI cases, the 0.8 mm and 1.6 mm bars trip the boundary layer to 

fully turbulent, and the 0.4 mm bar causes transition to start by Station 

7 and finish by Station 8. Transition has started at Station 8 of the 

baseline case and is complete near Station 10. In all of these cases 

transition begins far enough upstream to prevent separation. The mean 

profiles at all stations show that the boundary layer is thinnest in the 

baseline case and that the thickness increases with the bar size. With 

the 0.4 mm and 0.8 mm bars the boundary layer is only slightly thicker 

than in the baseline case, but it is substantially thicker in the 1.6 mm 

bar case. 

 Some consistent trends run through the data from all cases. If a bar 

is large enough, it will immediately trip the boundary layer to fully 

turbulent and prevent separation. As Reynolds number increases, the 

boundary layer thickness decreases as does the thickness of the bar 

required for tripping. If a bar is small enough, it initially appears to 

have no effect on the boundary layer. The boundary layer appears to 

proceed over the bar with no measurable change in the mean velocity 

or u′ from the corresponding baseline case. The bars must, however, 

introduce some small disturbance into the boundary layer. The 

boundary layer is unstable to small disturbances in the adverse 

pressure gradient region, so the small disturbances grow and 

eventually cause transition. Larger bars must impart larger (albeit 

sometimes still undetectable) perturbations than the smaller bars, 

resulting in transition locations that move upstream as bar size is 

increased. The optimal bar size depends on the Reynolds number and 

is discussed further below. 

 The present results shed some light on the transition mechanism 

under both high and low FSTI conditions. Volino [18] examined 

spectra of the fluctuating velocity in the boundary layers and shear 

layers of the baseline cases. He observed sharp peaks in the spectra of 

the low FSTI cases at frequencies that matched the most unstable 

frequencies for Tollmien-Schlicthing (TS) waves in the boundary layer 

just upstream of separation. He therefore concluded that transition in 

the shear layer might be through a TS mechanism in these cases. In 

the high FSTI cases, Volino [18] observed broadband peaks in the 

spectra, and the relatively long transition regions noted above. Volino 

and Hultgren [22] made similar observation, and concluded that the 

high FSTI separated flow transition was through a bypass mode, very 

similar to high FSTI transition in an attached boundary layer. Volino 

[18], however, noted that the broadband peaks in the high FSTI case 

spectra were centered at the same frequencies as in the low FSTI cases, 

suggesting a similar transition mechanism under high and low FSTI

conditions. He concluded that disturbances which began to grow in 

the boundary layer prior to separation were causing a TS type 

transition in the shear layer over the separation bubble in both the high 

and low FSTI cases. High FSTI has a strong effect in moving 

transition upstream, but the bars in the present cases had an equally 

strong or stronger effect in both the high and low FSTI cases. This 

confirms that the free-stream turbulence is not solely responsible for 

bypass transition in the high FSTI cases. 

 The magnitude of the disturbances induced by the bars is too small 

to be quantified based on the mean or u′ results presented above. 

Perhaps more can be learned from boundary layer spectra. Analysis of 

spectra based on u′ fluctuations for the present cases show some 

interesting but inconclusive results. Volino [18] found that u′ spectra 

are often characterized by low frequency fluctuations that are induced 

by the free-stream and have no direct effect on transition. In the early 

stages of transition, these low frequency fluctuations can hide the very 

low amplitude fluctuations important for transition. Volino [18] found 

that spectra of the turbulent shear stress are less affected by the low 

frequency unsteadiness and can provide a better means for detecting 

the early stages of transition. Acquisition and analysis of turbulent 

shear stress data for the cases of the present study may prove useful for 

explaining and quantifying the transition mechanism. 

Shape Factor and Momentum Thickness 

 The shape factor and momentum thickness are useful parameters for 

evaluating the state of the boundary layer with respect to separation, 

transition and losses. They provide a means for summarizing the 

information presented in the velocity profiles of Figs. 4 through 13. In 

the present cases, the boundary layer has a shape factor, H, of about 

2.4 at the end of the favorable pressure gradient region. This is the 

expected value for this laminar, accelerated boundary layer. If the 

boundary layer separates, the displacement thickness grows rapidly, 

while the momentum thickness remains nearly constant. The result is 

a very high shape factor. If the boundary layer reattaches, the 

displacement thickness drops, and the momentum thickness begins to 

grow. The boundary layer eventually recovers to a fully turbulent 

shape, with a shape factor of about 1.6 in the present cases. 

 Stage losses in a multi-blade turbine cascade can be determined 

through measurement of the momentum deficit in the wake 

NASA/CR—2012-217656 104
 9 Copyright © 2003 by ASME 

downstream of the blade row. With the single-passage facility of the 

present study, wake measurements are not meaningful, since there is 

flow on only one side of the airfoils on each side of the passage. If a 

boundary layer separates and does not fully reattach, or reattaches near 

the trailing edge, momentum thickness will be relatively low at the 

trailing edge, and high losses will be generated in the wake 

downstream of the passage. While the losses in such a case cannot be 

quantified in the present study, it is safe to assume that they would be 

unacceptably high, and that there would be an unacceptable loss of lift 

from the airfoils. For those cases in which the boundary layer 

reattaches and recovers to a fully developed turbulent shape, the 

suction side profile loss is likely the dominant loss mechanism in the 

passage (Howell et al. [9]). As explained by Howell et al. [9], for a 

given shape factor and passage exit angle, the momentum thickness of 

the suction side boundary layer at the trailing edge is proportional to 

the suction side profile loss. 

 Figure 14 provides an example of the development of the shape 

factor and momentum thickness, using the low FSTI, Re=100,000 

cases. In the baseline case, H increases from 2.4 to a high value of 

about 5 as the boundary layer separates. Reattachment occurs near the 

trailing edge, and H drops to about 2.2, which is still above the 

turbulent value of 1.6, indicating that recovery from the separation is 

not complete. In the 0.4 mm bar case, the boundary layer separates 

and H reaches a value of 3.7. The boundary layer then reattaches, and 

H gradually drops to a fully turbulent value by the trailing edge. With 

the 0.8 mm bar, transition occurs far enough upstream to prevent 

separation, and H drops continuously from a laminar value to a 

turbulent value as transition occurs. In the 1.6 mm bar case the 

boundary layer is tripped to turbulence, and H quickly reaches its 

turbulent value. The 1.6 mm bar causes an immediate thickening of 

the boundary layer, and the momentum thickness remains higher than 

in the other cases at all streamwise locations. The 0.4 and 0.8 mm bars 

appear to have no immediate effect on θ. When transition and 

reattachment occur, however, θ begins to rise. When reattachment 

occurs in the baseline case, it causes θ to increase to a higher value 

than in the 0.4 and 0.8 mm bar cases. Near the trailing edge θ is 

lowest in the 0.4 mm bar case. This would presumably be the case 

with the lowest profile losses. The larger bars force transition to occur 

farther upstream than necessary, resulting in a longer turbulent region 

and higher losses. In the baseline case the separation bubble becomes 

relatively thick, resulting in a thick boundary layer after reattachment. 

The 0.4 mm bar case provides a good example of the controlled 

diffusion described by Hourmouziadis [1]. With the 0.4 mm bar, the 

separation bubble is relatively thin, and the turbulent region is 

relatively short, resulting in lower losses. 

 Figure 15 shows the shape factor and momentum thickness at 

Station 11 (s/Ls=0.94), near the trailing edge, for all the low FSTI

cases. In the Re=25,000 cases, the shape factor indicates that the 

boundary layer only reattaches in the 1.6 mm bar case. The shape 

factor in this case is still above the expected turbulent value, indicating 

that recovery from the separation in not complete. Comparison of 

momentum thicknesses is not meaningful at this Re. For the 

Re=50,000 cases, the 0.8 and 1.6 mm bars cause reattachment, while 

recovery from the separation is only partially complete in the 0.4 mm 

bar case and the shear layer remains separated in the baseline case. 

The 0.8 and 1.6 mm bar cases are, therefore, preferable at this Re, and 

both have about the same momentum thickness at Station 11. At 

Re=100,000, already described in Fig. 14, the 0.4 mm bar produces the 

lowest losses. The 0.4 mm bar case is also best at Re=200,000, with 

slightly lower losses than with the 0.8 mm bar or in the baseline case. 

At Re=300,000, transition occurs sufficiently far upstream in the 

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1 0

1

2

3

4

5

s/Ls

H

a) baseline

0.4 mm bar

0.8 mm bar

1.6 mm bar

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1 0

0.5

1

s/Ls

θ [mm]

b)

Fig. 14 Shape factor and momentum thickness versus 

streamwise location, low FSTI, Re=100,000: (a) H, (b) θ

0 0.5 1 1.5 2 2.5 3 3.5

x 105

0

2

4

6

Re

H

a) baseline

0.4 mm trip

0.8 mm trip

1.6 mm trip

0 0.5 1 1.5 2 2.5 3 3.5

x 105

0

0.5

1

1.5

2

Re

θ [mm]

b)

Low FSTI

Fig. 15 Station 11 shape factor and momentum thickness 

vs Re, low FSTI cases: (a) H, (b) θ

0 0.5 1 1.5 2 2.5 3 3.5

x 105

0

1

2

3

Re

H

a)

baseline

0.4 mm trip

0.8 mm trip

1.6 mm trip

0 0.5 1 1.5 2 2.5 3 3.5

x 105

0

0.5

1

1.5

2

Re

θ [mm]

b)

High FSTI

Fig. 16 Station 11 shape factor and momentum thickness 

versus Re, high FSTI cases: (a) H, (b) θ

NASA/CR—2012-217656 105
 10 Copyright © 2003 by ASME 

baseline case to keep the separation bubble small and produce lower 

losses than in any of the cases with bars. 

 The Station 11 shape factors and momentum thicknesses for the 

high FSTI cases are shown in Fig. 16. As in the low FSTI cases, only 

the 1.6 mm bar is large enough to force reattachment at Re=25,000, 

and even it does not quite result in full recovery to a turbulent profile. 

At Re=50,000, the shape factor shows that all of the bars cause 

reattachment, while the boundary layer in the baseline case has 

reattached but not fully recovered from the separation. The 0.4 and 

0.8 mm bar cases have lower momentum thickness than the 1.6 mm 

bar case. At the higher Re, reattachment is complete in all cases, and 

the 1.6 mm bar cases have significantly higher losses than the other 

cases. At Re=100,000, the 0.4 and 0.8 mm bar cases and the baseline 

case all have about the same losses. As Re increases to 200,000 and 

300,000, the baseline case emerges as the case with lowest losses, in 

agreement with the low FSTI cases of Fig. 15. 

 The optimal bar height clearly varies with the Reynolds number. As 

Re increases, the boundary layer becomes thinner and more prone to 

transition, so a smaller bar is needed. At Re=25,000, the 1.6 mm bar is 

needed, and a larger bar would be desirable to force a more complete 

reattachment. At Re=50,000, the 0.8 mm bar is best, since it is large 

enough to cause reattachment at low FSTI, but produces lower losses 

than the thicker bar at high FSTI. At Re=100,000 and 200,000, the 0.4 

mm bar is best, since it is large enough to force complete reattachment 

at low FSTI, and results in equal or slightly lower losses than the 

baseline or 0.8 mm bar cases. At Re=300,000, the baseline case is 

best, although the losses are only slightly lower than those of the 0.4 

mm bar case. If a bar is used for passive flow control, a single bar 

thickness must be chosen for optimal overall performance. The best 

size will depend on the operating range of the engine. If the operating 

range is large, a compromise between improved performance at cruise 

and higher losses at takeoff may be needed. 

Correlation of Results 

 The size of a bar necessary to trip a boundary layer to turbulence 

can be predicted using the following correlation from Gibbings [23]. 

Red = Udd/ν > 600 (1) 

where d is the bar thickness and Ud is the velocity in the untripped 

boundary layer at y=d at the streamwise location of the bar. Equation 

1 predicts that bar thicknesses of 4.7 mm, 2.3 mm, 1.3 mm, 0.68 mm, 

and 0.50 mm would be needed to immediately trip the boundary layer 

to turbulent in the Re=25,000 through 300,000 cases respectively. In 

agreement with this prediction, the results above show that the 

boundary layer was only tripped in the Re=200,000 and 300,000 cases 

with the 0.8 mm and 1.6 mm bars, and in the Re=100,000 cases with 

the 1.6 mm bar. Since an optimal bar does not immediately trip the 

boundary layer, it will be thinner than indicated by Eqn. 1. 

 The most effective bars in the present cases appear to be those 

which cause reattachment to begin between Stations 8 and 9, at s/Ls of 

about 0.74. When reattachment begins by this location, there is 

sufficient distance downstream for the reattachment and recovery from 

the separation to be completed before the trailing edge. The beginning 

of reattachment and the start of transition are related and occur at 

approximately the same location. There are a few correlations in the 

literature for prediction of the distance from separation to transition 

onset. In general they are not very robust, but some give reasonable 

estimates. Mayle [2] provides the following correlations. 

Rest = 300 Reθs

0.7 (short bubble) (2) 

Rest = 1000 Reθs

0.7 (long bubble) (3) 

1 1.5 2 2.5 3 3.5 4 4.5 5 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

d/θp

(s −s

t p

)

m/(s −

t s

p

)

b

Low FSTI, predicted stb

High FSTI, predicted stb

Low FSTI, measured stb

High FSTI, measured stb

Equation 4

Fig. 17 Correlation of transition and reattachment start 

location to bar height, bars indicate range of possible 

values resulting from finite station spacing

Eqns. 2 and 3 apply to short and long separation bubbles respectively. 

Volino [7] found that the present baseline case results lie between the 

predictions of Eqns. 2 and 3, tending toward the long bubble 

correlation at low FSTI and about midway between the two 

correlations at high FSTI. Although they differ by a factor of 3, Eqns. 

2 and 3 provide at least a rough estimate of the reattachment location. 

 The following correlation provides an estimate of the effect of bar 

height on reattachment location. The equation is based on a curve fit 

of the present data. 

(st-sp)m/(st-sp)b = (1+0.23(d/θp)

1.56)-1 (4) 

where (st-sp)b is the distance from the suction side velocity maximum 

(s/Ls=0.53) to the location of the beginning of reattachment in the 

baseline case, and (st-sp)m is this distance with a bar in place. The 

present data along with Eqn. 4 are shown in Fig. 17. The finite 

spacing of the streamwise measurement stations results in some 

uncertainty in the transition start location, as indicated by the error 

bars in Fig. 17. Volino [7] showed that the boundary layer behavior 

upstream of separation is predictable and laminar. To predict 

reattachment in a case with a bar, the laminar solution could be used to 

predict the separation location and the momentum thickness before 

separation. A correlation such as Eqn. 2 or 3 could then be used to 

estimate the distance to transition and reattachment in the baseline 

flow. This would give stb. Equation 4 could then be used to predict stm

for a given bar thickness. Alternatively, the desired stm could be 

specified and used with Eqn. 4 to predict the optimal bar thickness. 

Setting stm=169 mm (stm/Ls=0.74) and using the measured (when 

available) or predicted values for stb and θp for the baseline cases, 

optimal bar heights have been predicted for the present cases and are 

presented in Table 3. 

Discussion

 The above correlations are based only on the present data set, so it is 

doubtful that they are universally applicable. Still, they provide a start 

in the assessment of passive flow control devices. With more 

experiments with different airfoils, it may be possible to refine 

correlations such as Eqn. 4, to make them more generally applicable. 

NASA/CR—2012-217656 106
 11 Copyright © 2003 by ASME 

Table 3: Baseline case boundary layer thickness at bar 

location (s/Ls=0.53), and predicted bar heights for 

tripping (dtrip) and for incipient reattachment (dopt) at 

s/Ls=0.74 (all values in mm) 

Re Low FSTI High FSTI

×10-3 δ99.5 p θp dtrip dopt δ99.5 p θ p dtrip dopt

25 3.8 0.48 4.7 2.4 3.9 0.48 4.7 1.5 

50 2.7 0.34 2.3 1.2 2.8 0.33 2.3 0.6 

100 1.9 0.23 1.3 0.4 2.2 0.24 1.3 0.2 

200 1.3 0.16 0.7 0.2 1.6 0.16 0.7 0 

300 1.2 0.14 0.5 0.1 1.4 0.14 0.5 0 

 Alternative passive devices such as vortex generators or dimples 

should also be considered. There is no guarantee, however, that these 

devices will provide improvement over the present bars. Dimples 

have been used in some applications such as internal blade cooling, to 

provide enhanced heat transfer with lower pressure drop than 

boundary layer trips. Dimples enhance heat transfer by promoting 

turbulence and mixing, which would also tend to promote boundary 

layer attachment. In the present application, however, we do not seek 

to enhance turbulence or heat transfer, or even to fully eliminate the 

separation bubble. Since the smaller bars in the present experiments 

did not increase losses or boundary layer thickness, or have any other 

immediate measurable effect on the boundary layer, it is not clear that 

any other device will be superior. 

 While the present bars have proven effective, it is clear from Table 

3 and Figs. 15 and 16 that the optimal bar height varies with Re and 

FSTI. If applied passively, a single bar height would be selected for 

the entire operating range of the engine. If the operating range is 

large, a compromise will be necessary between improved performance 

at low Re and higher losses at high Re. Passive flow control has the 

distinct advantage of being relatively simple to implement in practice, 

but active flow control may provide a means for optimizing 

performance over a wider range of conditions. Unsteady active 

control also provides possibilities for further flexibility and improved 

performance not available with passive devices. Further consideration 

of active control is presented in Volino [24]. 

 The present cases all involve steady inlet flow. In engine flows, the 

periodic wakes from upstream airfoils will make the flow unsteady. 

While steady flow experiments are necessary for building 

understanding of the flow and flow control devices, experiments 

should eventually be performed in flows with wakes. 

CONCLUSIONS 

1. Rectangular bars have been successfully employed as flow 

control devices on the suction side of a low pressure turbine airfoil. 

Boundary layer reattachment was forced even in very low Reynolds 

number cases. 

2. Optimal bars are not large enough to immediately trip the 

boundary layer to turbulent or prevent separation, but rather induce 

very small disturbances which at first are essentially undetectable, but 

eventually promote transition in the shear layer at a downstream 

location. 

3. Bars were effective under both high and low FSTI conditions, 

indicating that the high FSTI transition is not simply a bypass 

transition induced by the free-stream. 

4. The optimal location for reattachment results in a relatively short 

turbulent region, but occurs sufficiently far upstream to prevent a large 

separation bubble and insure complete recovery from the separation 

before the trailing edge. A bar height can be selected to induce 

reattachment at the desired location. 

5. The optimal bar height varies with the Reynolds number and freestream turbulence level. Bars that were large enough to induce 

reattachment at the lowest Re produced significantly higher losses at 

the higher Re. If a wide range of Reynolds numbers are encountered 

in practice, some compromise between improved performance at low 

Re and higher losses at high Re will be necessary in the choice of an 

overall best bar height. 

ACKNOWLEDGEMENTS 

 This work was sponsored by the NASA Glenn Research Center. 

The grant monitor is Dr. David Ashpis. Additional matching support 

was provided through a U.S. Naval Academy Recognition Grant. 

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Turbine Airfoil Conditions,” ASME Journal of Turbomachinery, 

123, pp. 189-197. 

[23] Gibbings, J.C., 1959, “On Boundary-Layer Transition Wires,” 

Aeronautical Research Council, Current Papers 462. 

[24] Volino, R.J., 2003, “Separation Control on Low-Pressure Turbine 

Airfoils Using Synthetic Vortex Generator Jets,” ASME Paper 

GT2003-38729. 

NASA/CR—2012-217656 108
 1 Copyright © 2003 by ASME 

Proceedings of ASME Turbo Expo 2003 

Power for Land, Sea, and Air 

June 16-19, 2003, Atlanta, Georgia, USA

GT2003-38729 

SEPARATION CONTROL ON LOW-PRESSURE TURBINE AIRFOILS USING 

SYNTHETIC VORTEX GENERATOR JETS

Ralph J. Volino

Department of Mechanical Engineering 

United States Naval Academy 

Annapolis, Maryland 21402 

Email: volino@usna.edu

ABSTRACT 

 Oscillating vortex generator jets have been used to control boundary 

layer separation from the suction side of a low-pressure turbine airfoil. 

A low Reynolds number (Re=25,000) case with low free-stream 

turbulence has been investigated with detailed measurements including 

profiles of mean and fluctuating velocity and turbulent shear stress. 

Ensemble averaged profiles are computed for times within the jet 

pulsing cycle, and integral parameters and local skin friction 

coefficients are computed from these profiles. The jets are injected 

into the mainflow at a compound angle through a spanwise row of 

holes in the suction surface. Preliminary tests showed that the jets 

were effective over a wide range of frequencies and amplitudes. 

Detailed tests were conducted with a maximum blowing ratio of 4.7 

and a dimensionless oscillation frequency of 0.65. The outward pulse 

from the jets in each oscillation cycle causes a disturbance to move 

down the airfoil surface. The leading and trailing edge celerities for 

the disturbance match those expected for a turbulent spot. The 

disturbance is followed by a calmed region. Following the calmed 

region, the boundary layer does separate, but the separation bubble 

remains very thin. Results are compared to an uncontrolled baseline 

case in which the boundary layer separated and did not reattach, and a 

case controlled passively with a rectangular bar on the suction surface. 

The comparison indicates that losses will be substantially lower with 

the jets than in the baseline or passively controlled cases. 

NOMENCLATURE 

B blowing ratio 

Cf skin friction coefficient

Cp ( ) 2

T − ρU/PP2 e , pressure coefficient 

cμ momentum coefficient 

D diameter 

F+ (0.442Ls)/(UeT), dimensionless frequency 

FSTI free-stream turbulence intensity 

H δ

∗

/θ, shape factor 

K ) /)(/( dsdUU 2 ν ∞∞ , acceleration parameter 

Ls suction surface length 

M jet to boundary layer mass flux ratio 

P pressure 

PT upstream stagnation pressure

Re LU se /ν , exit Reynolds number 

Reθ momentum thickness Reynolds number 

S jet holes per unit span 

s streamwise coordinate, distance from leading edge 

T jet oscillation period 

t time 

U mean streamwise velocity 

U∞ local free-stream velocity 

Ue nominal exit free-stream velocity, based on inviscid solution 

Vjet jet velocity 

u′ time averaged rms streamwise fluctuating velocity 

u

~′ ensemble averaged rms streamwise fluctuating velocity 

− vu ′′ time averaged turbulent shear stress 

vu

~~− ′′ ensemble averaged turbulent shear stress 

y cross-stream coordinate, distance from wall 

δ

∗ displacement thickness 

λθ Reθ

2

K, Thwaites parameter 

ν kinematic viscosity 

ρ density 

θ momentum thickness 

Subscripts 

ave average over jet outpulse 

jet jet condition 

max maximum in jet cycle 

∞ free-stream 

INTRODUCTION 

 Modern low-pressure turbine (LPT) airfoils are subject to 

increasingly stronger pressure gradients as designers impose higher 

loading in an effort to improve efficiency and lower cost by reducing 

the number of airfoils in an engine. If the adverse pressure gradient on 

the suction side of these airfoils becomes strong enough, the boundary 

NASA/CR—2012-217656 109
 2 Copyright © 2003 by ASME 

layer will separate. Separation bubbles, particularly those which fail 

to reattach, can result in a significant loss of lift and a subsequent 

degradation of engine efficiency (e.g. Hourmouziadis [1], Mayle [2], 

and Sharma et al. [3]). The problem is particularly relevant in aircraft 

engines. Airfoils optimized to produce maximum power under takeoff 

conditions may still experience boundary layer separation at cruise 

conditions, due to the thinner air and lower Reynolds numbers at 

altitude. A component efficiency drop of 2% may occur between 

takeoff and cruise conditions in large commercial transport engines, 

and the difference could be as large as 7% in smaller engines operating 

at higher altitudes. 

 Separation on LPT airfoils is complicated by boundary layer 

transition, which can prevent separation if it occurs far enough 

upstream, or induce boundary layer reattachment if it occurs in the 

shear layer over a separation bubble. At lower Reynolds numbers 

transition will tend to occur farther downstream, hence the problems 

associated with performance at altitude. 

 Separated flow transition has been studied extensively, and in recent 

years several studies have focused on transition in the LPT. Volino [4] 

provides a review of much of that work. Separation can be affected 

through naturally occurring phenomena such as high free-stream 

turbulence intensity (FSTI) or the unsteadiness caused by wakes 

generated upstream of an airfoil. Further discussion of these effects is 

available in Volino [5]. While high FSTI and wakes help to mitigate 

separated flow problems, they clearly do not solve all problems, as 

evidenced by the known efficiency drop in modern engines at altitude. 

Howell et al. [6], for example, studied airfoils modified for higher lift, 

noting that their highly loaded airfoils might be close to a limit, and 

that even higher loading could cause unacceptable separation problems 

even in the presence of wakes. Looking beyond free-stream 

turbulence and wakes, other types of separation control could prove 

useful. Gad-el-Hak [7] provides a recent review. Techniques include 

boundary layer tripping, vortex generation, suction, and injection of 

fluid normal to the wall to either increase the boundary layer 

momentum or promote turbulence. 

 While the general literature is extensive, only a few studies have 

considered separation control under LPT conditions. Some have 

utilized passive techniques. Lake et al. [8] considered dimples and 

boundary layer trips. Van Treuren et al [9] considered vortex 

generators. Volino [5] used rectangular bars to impose disturbances in 

a boundary layer and move transition upstream. Passive flow control 

is appealing for its simplicity and the relative ease with which it might 

be implemented in gas turbine environments. It has its limitations, 

however. Volino [5] found that passive devices can successfully 

control separation even at the lowest Reynolds number of interest, but 

that these devices caused substantial increases in losses at higher Re. 

This is an important limitation for aircraft engines, where the Re range 

between takeoff and cruise is large. An active device could be turned 

off at high Re. Static passive devices are also unable to take advantage 

of the unsteadiness caused by wake passing. An active device might 

be timed to turn on and off in response to wake passing events. 

Unsteady devices might also take advantage of the calmed region 

following a transient turbulent event. 

 The literature contains several examples of active separation 

control. Lee et al. [10] used blowing in supersonic engine inlets to 

prevent or control separation. Sturm et al. [11] reported on blowing in 

a compressor cascade. Johnston and Nishi [12] used vortex-generatorjets (VGJs) to control separation in turbulent boundary layers. This 

method utilizes blowing from “small, skewed, and pitched holes” to 

create streamwise vortices similar to those created by solid vortex 

generators. Any jet injected into a flow will tend to produce some 

turbulence, and the turbulent mixing will tend to bring some high 

momentum fluid into the near wall region and inhibit separation. 

Streamwise vortices bring additional high momentum fluid into the 

near wall region. The most effective VGJs enter the boundary layer at 

a relatively shallow pitch angle (typically 30 to 45 degrees) relative to 

the wall and a high skew angle (45 to 90 degrees) relative to the main 

flow. Compton and Johnston [13] showed that the co-rotating vortices 

produced by VGJs are stronger and more effective for separation 

control than the counter-rotating vortices which form downstream of a 

normal jet. McManus et al. [14] and Raghunathan et al. [15] used 

pulsed VGJs. Sinha and Pal [16] used acoustic excitation to perturb an 

unsteady separating flow. Jacobson and Reynolds [17] used 

piezoelectrically driven cantilevers to influence the near wall 

turbulence structure on a flat plate. They noted that the devices could 

be used in separation control. Miau et al. [18] used an oscillating 

fence to promote reattachment downstream of a backward facing step. 

Sinha et al. [19] used a driven flexible wall transducer to detect 

pressure fluctuations and then produce near wall vortices upstream of 

separation. Whitehead et al. [20] used a film transducer to produce 

airfoil vibrations and reduce separation at high angles of attack. 

Oscillatory blowing has been used in several studies to control 

separation on airfoils. Amitay and Glezer [21] provide one recent 

example. Oscillatory jets are often referred to as “synthetic jets” since 

they have no net mass flow. They are typically directed normal to a 

surface, meaning that they probably do not produce such strong 

streamwise vortices as VGJs. 

Only a few active control studies have been conducted under LPT 

conditions. Huang et al. [22] and Hultgren and Ashpis [23] employed 

high voltage electrodes to produce glow discharge plasma in a 

boundary layer to control separation. Bons et al. [24, 25] used steady 

and pulsed VGJs to successfully control separation on LPT airfoils. 

They used the “Pak-B” airfoil, which is an industry supplied research 

airfoil that is prone to separation problems at low Re. It has been used 

in numerous studies, as noted by Volino [4]. Bons et al. [24] used 

spanwise rows of VGJs at several streamwise locations on the suction 

surface of the airfoil, and found that a row near the suction surface 

velocity maximum (pressure minimum) was most effective. The VGJ 

holes were oriented at 30 degrees to the surface and 90 degrees to the 

main flow. All holes were oriented in the same direction, to produce 

co-rotating vortices. Reynolds numbers as low as 60,000 (based on 

suction surface length and exit velocity) were considered. Bons et al. 

[25] found that both steady and pulsed jets were effective in 

controlling separation. The pulsed jets were fully effective even when 

the dimensionless pulsing frequency, F+, was as low as 0.1, where F+

is a ratio of the transit time for flow between the VGJ hole and the 

trailing edge to the time interval between pulses. Ensemble averaged 

velocity profiles showed a long relaxation or “calmed” period 

following each jet pulse. During this calm period the boundary layer 

remained attached long after the turbulence generated by the pulse had 

moved downstream. Calmed regions have been observed following 

turbulent spots in transitional boundary layers (e.g. Gostelow et al. 

[26] and Schulte and Hodson [27]). The mean velocity profiles in the 

calmed region gradually relax from the turbulent shape associated with 

the turbulent spot they follow, to a laminar (and in some cases 

separated) profile shape. The calmed boundary layer is very resistant 

to separation, much like a turbulent boundary layer, but it is very 

laminar-like in terms of its fluctuation levels and low losses. The 

pulsed jets were more effective than continuous jets, even when the 

pulsed jet duty cycle was as low as 1%. This was believed to indicate 

that the starting vortex formed at the beginning of each jet pulse was 

responsible for most of the flow control in the pulsed jet cases. 

 Synthetic jets hold an advantage over continuous or pulsed jets in 

that they require no net mass flow. In the LPT environment, this 

NASA/CR—2012-217656 110
 3 Copyright © 2003 by ASME 

means that no compressor bleed air is required. Use of bleed air for 

flow control or cooling comes at a cost in efficiency, although the 

small amount of air required for the pulsed jets of Bons et al. [25] 

might not be prohibitive if bleed air were already routed to the airfoils 

for cooling. Synthetic jets would not be useful for a cooled airfoil 

since ingestion of hot gas into the airfoil would be harmful. For 

uncooled LPT airfoils, however, the airfoil temperature will match the 

main flow temperature, and ingestion of hot gas should be acceptable. 

Routing of bleed air to uncooled airfoils for flow control may present a 

prohibitive addition of complexity and weight. With synthetic jets this 

problem could be avoided. In the present study, the oscillating flow of 

synthetic jets and the compound angle injection of vortex generator 

jets are combined to produce synthetic VGJs. This is believed to be 

the first application of synthetic VGJs. They are used to control the 

flow over a Pak-B airfoil. A survey of the literature indicates that the 

jet locations and angles chosen by Bons et al. [25] were likely optimal, 

so their geometry has been copied in the present study. 

 There are many parameters which could be varied in a synthetic jet 

study, including Reynolds number, FSTI, jet geometry, jet location, jet 

velocity, jet oscillation frequency, and jet waveform, to name a few. 

These are all potentially important parameters and should eventually 

be studied. The scope of the present study is more focused. A single 

experimental case is completely documented with detailed 

measurements including time resolved mean and fluctuating velocity 

and turbulent shear stress throughout the flow field. The goals of the 

study are to build an understanding of the physics of how synthetic 

VGJs control separation and to generate questions for future 

parametric studies which may lead to optimized flow control for a 

broad range of flow conditions. 

 Volino [4, 28] studied unmodified flow over the Pak-B airfoil at 

Reynolds numbers ranging from 25,000 to 300,000 under both high 

and low FSTI. In nearly all cases the boundary layer separated from 

the suction side of the airfoil. At all but the lowest Reynolds numbers 

it reattached before the trailing edge. The most severely separated 

case was the low FSTI, Re=25,000 case. This case has, therefore, been 

chosen as the test case for the present application of synthetic VGJs. 

The unmodified case from Volino [4] is used as a baseline case for 

comparison to the new results. Also used for comparison is a case 

from Volino [5] in which a passive bar was employed to force 

reattachment. The bar was located at s/Ls=0.51, extending along the 

airfoil span. Its streamwise width was 6.35 mm and its height was 1.6 

mm. The suction surface length was 228.6 mm. Bars of various 

heights were tested. The 1.6 mm bar was the smallest bar to cause 

reattachment at Re=25,000. Volino [5] found that the most effective 

bars in terms of minimizing losses were not large enough to 

immediately trip the boundary layer to turbulent. Rather, they induced 

small disturbances with grew and caused transition and reattachment 

downstream of a small separation bubble. 

EXPERIMENTS

 Experiments were conducted in a low speed wind tunnel, described 

by Volino et al. [29]. Briefly, air enters through blowers and passes 

through a honeycomb, a series of screens, two settling chambers, and a 

three-dimensional contraction before entering the test section. At the 

exit of the contraction, the mean velocity is uniform to within 1%. 

The FSTI is 0.5%±0.05%. Nearly all of this free-stream “turbulence” 

is actually streamwise unsteadiness at frequencies below 20 Hz and is 

not associated with turbulent eddies. The rms intensities of the three 

components of the unsteadiness are 0.7%, 0.2% and 0.2% in the 

streamwise, pitchwise and spanwise directions, respectively. The test 

section immediately follows the contraction. 

 The test section, shown in Fig. 1, consists of the passage between 

two airfoils. Details are listed in Table 1, and more information is 

available in Volino [4]. A large span-to-chord ratio of 4.3 was chosen 

to insure two-dimensional flow at the spanwise centerline of the 

airfoils, where all measurements were made. Upstream of each airfoil 

are flaps, which control the amount of bleed air allowed to escape 

from the passage. The flaps, along with a tailboard on the pressure 

side of the passage, are adjusted to produce the correct leading edge 

flow and pressure gradient along the airfoils. The flow in the passage 

matches that in a multi-blade cascade. The single passage 

configuration allows for a large scale passage and better probe access 

than possible with a multi-blade cascade in the same size wind tunnel. 

The wake downstream of the passage is not representative of a multiblade facility, however, since there is flow only on one side of each 

airfoil. Downstream effects that could influence the upstream flow in 

the passage are also potentially missed. Experimental conditions 

match those of the low FSTI, Re=25,000 baseline case of Volino [4] 

and the passive bar case of Volino [5]. 

 The synthetic VGJs were produced from a cavity within the suction 

side airfoil. The airfoils are machined from high density foam, which 

has a consistency much like hard wood. The surface of each airfoil 

was sanded smooth, painted, and sanded again to provide a smooth 

surface. A 1.27 cm diameter hole was drilled through the airfoil span 

at about mid-chord, as shown in Fig. 2, to form a plenum. One end of 

the plenum is plugged, and the narrow end of a funnel is inserted in the 

other. A 20.3 cm diameter loudspeaker (100W subwoofer) is attached 

at the wide end of the funnel. The funnel is sealed to the speaker and 

to the airfoil with silicone RTV to prevent air leakage. The speaker is 

driven with a 200W audio amplifier, which is in turn powered with a 

12V DC power supply and driven by a function generator. For the 

present study the function generator was set to output a sine wave. 

The amplitude of the signal from the function generator and the gain 

of the amplifier were adjusted to provide the desired input voltage to 

the speaker. Holes for the VGJs were drilled into the suction surface 

in a spanwise line at s/Ls=0.514. The holes are 0.8 mm in diameter 

(0.35% of Ls) and are spaced 8.5 mm apart (3.7% of Ls). The holes are 

drilled at a 90 degree skew angle with respect to the main flow and a 

30 degree pitch with respect to the surface, as shown in Fig. 2. Each 

hole extends from the suction surface into the cavity in the core of the 

airfoil. The length to diameter ratio of the holes is 7.5. 

Measurements 

 Pressure surveys were made using a pressure transducer (0-870 Pa 

range Validyne transducer) and a Scanivalve. Stagnation pressure was 

measured with a pitot tube upstream of the passage inlet, and eleven 

pressure taps were located on each airfoil along their spanwise 

centerlines. Locations of the taps on the suction side are listed in 

Table 2 along with measured local FSTI components and the 

acceleration parameter, K, at these stations based on a non-separating, 

inviscid solution. The uncertainty in the suction side pressure 

coefficients was 7%. Most of this uncertainty was due to bias error. 

Stochastic error was minimized by averaging pressure transducer 

readings over a 10 second period. 

 Velocity profiles on the suction surface were measured at 

streamwise stations corresponding to pressure taps 7-11, as given in 

Table 2, and at four additional stations, labeled 7.5, 8.5, 9.5 and 10.5, 

centered between the pressure taps. All stations are downstream of the 

VGJ holes. Stations 7, 8, 9, 10 and 11 correspond to stations 

documented in the baseline case in Volino [4, 28]. Profiles at Stations 

1-6 are fully documented for the baseline case in Volino [4, 28], and 

show that the upstream boundary layer closely follows a laminar 

solution. Profiles were measured near but not at the spanwise 

NASA/CR—2012-217656 111
 4 Copyright © 2003 by ASME 

tailboard

flap

flap bleed air

bleed air

35o

60o

Fig. 1 Schematic of the test section 

VGJ Holes

(a)

Cavity

Cavity

Mainflow

Into Page

Jet

Jet

30°

30°

(b)

Fig. 2 Drawing of suction side airfoil with cavity and VGJs: 

(a) full airfoil, (b) cross section ofVGJ holes 

centerline of the airfoil to insure that the pressure taps did not interfere 

with the velocity measurements. Profiles were acquired with a hotwire anemometer (AA Lab Systems model AN-1003) and a single 

sensor boundary layer probe (TSI model 1218-T1.5). The sensor 

diameter is 3.8 μm, and the active length is 1.27 mm. At each 

measurement location, data were acquired for 26 seconds at a 20 kHz 

sampling rate (219 samples). All raw data were saved. The high 

sampling rate provides an essentially continuous signal, and the long 

sampling time results in low uncertainty in both statistical and spectral 

quantities. Data were acquired at 60 wall normal locations in each 

profile, extending from the wall to the free-stream, with most points 

concentrated in the near wall region. The closest point was within 0.1 

mm of the wall, which corresponds to y/Ls=0.0004 and about 0.015 

boundary layer thicknesses. Flow direction in a separation bubble 

cannot be determined with a single-sensor hot-wire, but velocity 

magnitude can be measured and was found to be essentially zero 

within the bubbles of the present cases. Determining the direction was 

not, therefore, considered essential. Uncertainty in the mean velocity 

is 3-5% except in the very near wall region, where near-wall 

corrections (Wills [30]) were applied to the mean velocity. 

Uncertainties in the momentum and displacement thicknesses 

computed from the mean profiles are 10%. Uncertainty in the shape 

factor, H, is 8%. Local skin friction coefficients were computed from 

the near wall mean velocity profiles using the technique of Volino and 

Simon [31]. This technique accounts for streamwise pressure gradient 

effects on the mean profile. The uncertainty in Cf is 8%. The 

uncertainty in the fluctuating streamwise velocity is below 10%. As 

Table 1: Test section parameters 

Axial 

Chord 

[mm] 

True 

Chord 

[mm] 

Pitch 

[mm] 

Span 

[mm] 

Suction 

side, Ls

[mm] 

Inlet 

flow 

angle 

Exit 

flow 

angle 

153.6 170.4 136.0 660.4 228.6 35° 60° 

Table 2: Measurement station locations, local acceleration 

(inviscid soln.), and measured local free-stream turbulence 

Station s/Ls K×106 ′/Uu ∞ [%] ′/Uv ∞ [%] 

1 0.111 6.32 0.44 

2 0.194 4.80 0.39 

3 0.278 3.44 0.37 

4 0.361 3.00 0.38 

5 0.444 2.48 0.39 

6 0.528 -0.08 0.41 

7 0.611 -3.24 0.47 0.05 

8 0.694 -3.80 0.47 0.12 

9 0.777 -2.32 0.48 0.14 

10 0.861 -2.12 0.54 0.11 

11 0.944 -0.72 0.51 0.11 

explained in Volino [4] based on the work of Ligrani and Bradshaw 

[32], spatial averaging effects due to the finite length of the hot-wire 

sensor should not be significant in the present case. 

 Profiles were also acquired using a cross-sensor boundary layer 

probe (TSI 1243-20). The sensors are 51 μm diameter hot films with 

1.02 mm active lengths. The probe is used to document the 

instantaneous turbulent shear stress, − vu ′′ . Profiles were acquired at 

the same stations as with the single-sensor probe. Data were acquired 

at 25 locations in each profile, extending from 1 mm from the wall to 

the free-stream. Sampling rates and times were the same as for the 

single-sensor profiles. The vortices induced by the oscillating jets 

cause significant secondary velocity, particularly at the streamwise 

stations immediately downstream of the jet holes. The magnitude of 

these secondary velocity components remains below 20% of the local 

streamwise velocity, however, so they should not cause significant 

error in the hot-wire measurements. The uncertainty in − vu ′′ is 10%. 

 The VGJ velocities were measured using a hot-film probe (TSI 

model 1210-10A) with a 0.25 mm active sensor length. The sensor 

was placed directly over the exit of the jet hole. During outflow from 

the hole, the jet was expected to blow directly across the sensor, 

providing an accurate measure of the jet velocity. Uncertainty in the 

velocity is 5% and results mainly from uncertainty in the position of 

the sensor, which could lead to a slightly lower velocity reading than 

the velocity at the jet exit plane. During inflow of the oscillating jet, 

the flow is expected to behave more like a sink flow than a jet. The 

measured velocity does not, therefore, provide an accurate indicator of 

the velocity inside the hole during inflow. The jet velocity was 

calibrated against the rms input voltage to the speaker with the main 

flow in the wind tunnel turned off, and the calibration was used to set 

the jet velocity in later experiments. The jet velocity is fixed by the 

frequency and amplitude of the displacement of the speaker 

diaphragm, which causes a pressurization of the cavity relative to the 

pressure at the jet exits. The dynamic pressure of the flow through the 

test section is about 2.4 Pa at the jet location. Since the test section 

exits to atmosphere, the dynamic pressure results in an average 

pressure of 2.4 Pa vacuum in the cavity, and a 2.4 Pa pressure 

difference across the speaker diaphragm. This pressure is much 

smaller than the pressure experienced by the speaker when driving the 

jets, and is not expected to influence its motion. If the amplitude of 

the diaphragm motion is unchanged by the presence of flow in the test 

NASA/CR—2012-217656 112
 5 Copyright © 2003 by ASME 

section, the jets should drive approximately the same mass flow 

through the holes with the wind tunnel on or off. The jet velocity at 

the exit plane will presumably be affected by the mean flow, however. 

 The measured maximum jet exit velocity was 9.4 m/s for most of 

the cases presented below. Given the sensor length and the diameter 

of the holes, the measured velocity is an average over the middle 30% 

of the jet. The Reynolds number based on this velocity and the jet 

diameter is 500. At this Reynolds number, approximately 30 

diameters would be needed to establish fully developed laminar flow 

inside the jet holes. Since the length to diameter ratio in only 7.5 and 

the jets are unsteady, the jet velocity is not expected to have a fully 

developed parabolic laminar profile, but rather a more flat profile. 

Given the averaging due to the sensor size and the expectation of a flat 

profile, the instantaneous mean velocity of the jet is assumed to 

approximately equal the measured velocity. Ideally this assumption 

would be checked with a survey of velocity across the jet exit plane, 

but the very small jet diameter precludes an accurate survey. The 

uncertainty in instantaneous mean velocity is, therefore, higher than 

the 5% uncertainty in the measured velocity. The uncertainty is 

estimated to be between 10 and 20%. 

Data Processing 

 In addition to conventional time averaging, the velocity data were 

ensemble averaged relative to the time within each jet oscillation 

cycle. For this purpose, the speaker input voltage was digitized 

simultaneously along with the instantaneous velocity data. Data were 

ensemble averaged at 24 instances within the cycle. At each instance, 

data were averaged over 1/180th of the cycle. For each 26 second data 

trace, this results in roughly 3000 data points to average for each 

ensemble. With this many data points to average over a 26 second 

time record, the ensemble averaged results are well resolved and have 

uncertainties as low as those given above for the time averaged results. 

The start of the cycle was arbitrarily chosen as the instant when the 

speaker input voltage crossed from negative to positive. As will be 

shown below, this roughly corresponds to the beginning of the jets’ 

outward pulse. 

RESULTS 

Jet Velocity 

 Figure 3 shows a typical time trace of the measured jet velocity and 

the speaker input voltage. The frequency of the input signal was set to 

a nominal value of 10 Hz (actual value was 10.5 Hz). The maximum 

and average velocities in each outward pulse were 9.4 m/s and 5.9 m/s 

respectively. The jet velocity can be expressed as a blowing ratio, B, 

defined as 

B=ρjetVjet / ρ∞U∞

where ρjet=ρ∞ since the jet fluid comes from the boundary layer, and 

U∞ (=2 m/s) is the local free-stream velocity at the jet location. For 

the present cases Bmax=4.7 during each cycle and Bave=3.0 during the 

outward pulse. The mass flux of the jets can be compared to the 

velocity deficit in the boundary layer as 

M=[ρjetVjet (πDjet2

/4)S]/[ρ∞U∞δ

*

] 

where S=0.118 holes/mm is the number of holes per unit span. The 

displacement thickness, δ

*

 (=1.08 mm) is the baseline case value at the 

jet location. This gives Mmax=0.258 and Mave=0.162. This could be 

interpreted to mean that 16% of the boundary layer in terms of 

displacement thickness is sucked off during the inflow half of each jet 

5 5.05 5.1 5.15 5.2 5.25 5.3 5.35 5.4 5.45 5.5 −2

0

2

4

6

8

10

time [s]

Volts or Velocity [m/s]

Jet velocity

Speaker Input Voltage

Fig. 3 Time trace of jet velocity and speaker input voltage, 

F+

=0.65 

cycle, and then re-injected into the boundary layer during the outward 

pulse. 

 The momentum coefficient, cμ, is defined as the ratio of the jet 

momentum to the free-stream dynamic pressure. Using the definition 

of Bons et al. [25], 

cμ=[ρjetVjet2

 (πDjet2/4)S]/[ρ∞U∞

2

Chord]

where the axial chord is 153.6 mm. The maximum value of cμ in each 

cycle is 0.0085. The average value of Vjet2 for the outpulse is 44 m2/s2. 

This gives an average cμ=0.0042. 

 Figure 3 shows that the speaker input voltage and the jet velocity 

are slightly out of phase. The jets lag the input voltage by about 

0.006 s, which is a dimensionless lag ∆t/T of 0.063, where T is the jet 

oscillation period of 0.095 s. The lag is expected, as the jets respond 

dynamically to the pressurizing of the cavity in the airfoil by the 

speaker, and there is no reason to expect the speaker voltage and the 

jet velocity to be exactly in phase. The finite distance from the 

speaker to the jet holes (of the order 0.5 m) and the finite speed of 

sound (340 m/s) will also lead to a time lag of the order 1 ms. The 

time lag increases with distance from the speaker along the airfoil 

span. At high jet frequencies, the time lag causes the jets along the 

span to be significantly out of phase with each other. Measurements, 

however, show that the jet amplitude along the span is uniform, 

regardless of the frequency. Variation in phase along the span could 

lead to difficulty in practice if attempts were made to time the jet 

pulsing to other cyclic events such as wake passing. It is not an issue 

in the present study. At the relatively low frequency of 10 Hz, the 

phase lag was not significant, and the jets were uniform in both phase 

and amplitude along the span. 

Pressure Profiles

 Pressure profiles were acquired for several jet amplitudes and jet 

frequencies. The general finding was that the jets were effective over 

a broad frequency range, so long as the amplitude was sufficiently 

high. Figure 4 shows Cp profiles for a range of jet amplitudes, with 

the jet frequency set to 10.5 Hz (F+=0.65). In all cases there is good 

agreement between the data and an inviscid solution for the Pak-B 

airfoil on the pressure side and the upstream portion of the suction 

side. In the adverse pressure gradient region on the suction side, 

differences are clear. Without the jets the boundary layer separates 

NASA/CR—2012-217656 113
 6 Copyright © 2003 by ASME 

0 0.2 0.4 0.6 0.8 1 0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

no jets

s/Ls

Cp

1.9

4.7

6.6

8.1

10.0

Bmax

inviscid solution

Fig. 4 Cp profiles, F+

=0.65, various blowing ratios 

0 0.2 0.4 0.6 0.8 1 0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

no jets

s/Ls

Cp

0.33

0.65

1.3

2.6

5.2

10.4

F+

inviscid solution

Fig. 5 Cp profiles, Bmax≈5, various F+

0 0.2 0.4 0.6 0.8 1 0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

baseline

inviscid solution

s/Ls

Cp

passive bar

jets

Fig. 6 Cp profiles, comparison of baseline, passive bar, 

and present jet case with Bmax=4.7 and F+

=0.65 

and does not reattach, as indicated by the region of constant Cp values. 

With Bmax=1.9 the boundary layer still does not reattach, but there is 

some sign that Cp is starting to drop at the last pressure tap. For the 

cases with Bmax=4.7 and above, the boundary layer does not appear to 

separate. Significant case to case differences in Cp are present right at 

the suction peak, but these are likely due to the injection of the jets at 

this location and their effect on the flow over the adjacent pressure tap. 

The differences diminish rapidly and are essentially gone by the next 

downstream measurement station. Bons et al. [24] demonstrated 

effective flow control with Bmax as low as 0.4 in their study. The 

significantly lower Re in the present study may explain the need for 

stronger jets. Figure 5 shows Cp profiles for several different jet 

frequencies with Bmax held approximately constant at about 5. There 

does not appear to be any clear separated region in any of the cases 

with jets. The suction peak is higher in all the cases with jets than in 

the baseline case. This is an expected result, suggesting that the lift 

will be higher when the boundary layer is attached. The F+=0.65 case 

appears to agree most closely with the inviscid solution. The broad 

range of effective frequencies agrees with the results of Bons et al. 

[25]. 

 The objective of the present study is not to establish the optimal jet 

conditions for the present case, but to investigate in detail a case in 

which the jets provide effective flow control. The dimensionless 

frequency F+=0.65 was chosen since it appeared to provide slightly 

better results than the other cases in Fig. 5, and Bmax=4.7 was chosen 

since it was the lowest effective blowing ratio tested. Figure 6 shows 

the Cp profile for the chosen case with jets along with the baseline 

case and the passive bar case of Volino [5]. 

Velocity Profiles 

 Figure 7 shows mean velocity, u′ and turbulent shear stress profiles 

for the present case, the baseline case, and the passive bar case. In the 

baseline case, the mean profiles show the boundary layer is on the 

verge of separating at Station 7, is clearly separated at Station 8, and 

the separation bubble grows through Station 11. In the bar case, the 

boundary layer separates from the bar. The separation bubble is 

visible at Station 7, and it grows through Station 9. At Station 10, the 

near wall velocities begin to rise, indicating the beginning of 

reattachment, and the mean profile shows clear reattachment at Station 

11. With the VGJs, the behavior is much different. There is no clear 

separation bubble. The mean profile appears to have an attached, 

laminar-like shape. The momentum deficit with the jets appears to be 

significantly lower at Station 11 than in the case with the bars, 

indicating lower losses with the jets. 

 The u′ values in the baseline and bar cases are very low at Station 

7, as expected since the boundary layer is still laminar. A peak 

appears downstream in the shear layer over the separation bubble. In 

the bar case, u′ begins to rise in the near wall region at Station 9, 

signaling imminent reattachment. At Stations 10 and 11, u′ in the bar 

case rises to the high values typical of a transitional boundary layer. 

In the jet case, u′ is high at all stations. At the upstream locations it is 

much higher than would be expected for a turbulent boundary layer. 

As will be shown below, however, much of the contribution to u′ is 

from 10 Hz unsteadiness associated with the jets and is not turbulence. 

The turbulent shear stress profiles show considerable momentum 

transport in the jet case at all stations, which helps to explain how the 

boundary layer remains attached. Unlike in a turbulent boundary 

layer, the − vu ′′ peak is well away from the wall. In the baseline case 

the shear layer does not transition to turbulent, and − vu ′′ remains 

near zero. In the bar case − vu ′′ profiles were not acquired. 

NASA/CR—2012-217656 114
 7 Copyright © 2003 by ASME 

0 0 0 0 0 1 0

0.05

U/Ue

y/Ls

(a) 7 8 9 10 11

0 0 0 0 0 0

0.05

u′/Ue

y/Ls

(b) 7 8 9 10 11

0.2

0 0 0 0 0 0

0.05

−u′v′/Ue

2

y/Ls

(c) 7 8 9 10 11

0.008

baseline passive bar jets

Fig. 7 Station 7-11 dimensionless time averaged profiles; 

comparison of baseline, passive bar and present jet cases: 

(a) mean velocity, (b) u′ , (c) turbulent shear stress 

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1 0

1

2

3

4

5

6

s/Ls

H

(a)

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1 0

0.5

1

1.5

2

s/Ls

θ [mm]

(b)

baseline

passive bar

jets

Fig. 8 Time averaged shape factor and momentum 

thickness verses streamwise location; comparison of 

baseline, passive bar and present jet cases: (a) H, (b) θ

 Figure 8 shows shape factor and momentum thickness as computed 

from the mean profiles of Fig. 7. The shape factor, H, provides a 

measure of the state of the boundary layer with respect to separation 

and transition. The shape factor in the baseline and bar cases rises 

rapidly after separation, as δ

*

 increases while θ remains nearly 

constant. In the baseline case the boundary layer never reattaches. In 

the bar case, transition and reattachment occur, causing δ

*

 to fall and θ

to rise. The shape factor begins to drop toward a turbulent value of 

about 1.6, but does not reach this value, indicating that the recovery 

from the separation is not complete. In the jet case, the shape factor 

remains at a laminar value of about 2. It never rises to a separated 

flow value, nor does it drop to a turbulent value. The momentum 

thickness provides a measure of the losses in a boundary layer. If the 

boundary layer reattaches before the trailing edge, the suction side 

boundary layer losses will be the dominant losses in an LPT passage 

(Howell et al. [6]). In the baseline case the boundary layer does not 

reattach, so although θ remains low, high losses would be expected in 

the wake downstream of the airfoils. For the bar and jet cases, 

however, the boundary layer is attached at the trailing edge. 

Momentum thickness is about 20% higher in the bar case, indicating 

that the jets are better able to control separation, while causing lower 

losses. 

Ensemble Averaged Velocity Profiles 

 The time averaged profiles of Fig. 7 indicate that the jets are 

effective in controlling the boundary layer, but they do not explain the 

mechanism by which the jets work. Figure 9 shows ensemble 

averaged mean velocity profiles. Profiles are shown for nine 

streamwise stations at 24 time increments within the jet oscillation 

cycle. Figures 10 and 11 show the corresponding u

~′ and vu

~~− ′′

profiles. Examining the profiles at Station 7, the mean profile initially 

appears to be laminar and attached, and both the u

~′ and vu

~~− ′′ values 

are near zero, indicating again that the flow is laminar. Given the 

phase lag shown in Fig. 3 between the speaker input and the jets, and 

the finite convection time between the jet hole and Station 7, one 

would expect that the disturbance created by the jet outpulse should 

arrive at Station 7 at t/T=0.18. In fact, however, the disturbance is not 

seen in the mean profile until t/T=0.333. In agreement, the u

~′ and 

vu

~~− ′′ values also rise above zero at t/T=0.333. This may suggest that 

the rising jet velocity must reach a sufficiently high amplitude before it 

can significantly affect the boundary layer. Comparing the observed 

phase lag and the jet velocity of Fig. 3 suggests that the jet velocity 

must be about 8 m/s, corresponding to an instantaneous B=4, for the 

jets to be effective. Continuing forward in time at Station 7, the jets 

cause a large disturbance in the mean profiles that continues until 

t/T=0.667. The time t/T=0.667 corresponds very closely with the end 

of the jet outpulse when the phase lag and convection time from the jet 

holes to Station 7 are considered. The large local minima and maxima 

in the mean velocity profiles indicate the jets are not merely adding 

turbulence to the boundary layer, but are inducing some flow structure, 

most likely streamwise vortices. These vortices likely cause spanwise 

variation in the velocity, particularly at the stations nearest the jets. In 

the present study, data were only acquired at one spanwise location. 

Spanwise surveys should be considered in future work. The u

~′ and 

vu

~~− ′′ profiles exhibit large peaks in the regions where the mean 

velocity gradients are highest in Fig. 9. The vu

~~− ′′ values have the 

appropriate sign, corresponding to the sign of / dyUd ~ in the mean 

profiles. It should be noted that the magnitude of u

~′ is smaller in the 

ensemble averaged profiles of Fig. 10 than in the time averaged 

profiles of Fig. 7. This indicates that much of the contribution to u′

in Fig. 7 is due to 10 Hz oscillations and not to turbulence. After 

t/T=0.667, the mean velocity profile resumes a laminar shape. There 

is no tendency toward boundary layer separation. The boundary layer 

did not separate in the baseline case at this station (Fig. 7), but the 

mean profile in the baseline case did appear closer to separation than 

in the present case. 

 Moving to the downstream stations, the leading edge of the 

disturbance, as observed in the mean profiles and the u

~′ and vu

~~− ′′

profiles moves to later values of t/T, as expected since the disturbance 

takes some time to convect downstream. The leading edge of the 

disturbance appears to move at about 90% of the local free-stream 

velocity, which corresponds with the expected leading edge celerity of 

a turbulent spot (e.g. Gostelow et al. [26], Schulte and Hodson [27]). 

The trailing edge of the disturbance appears to move at about 45% of 

the local free-stream velocity, agreeing with the expected trailing edge 

celerity of a turbulent spot. Because the leading and trailing edge 

celerities are different, the boundary layer is only disturbed by the jets 

during approximately 30% of the cycle at Station 7, but is disturbed 

during approximately 70% of the cycle at Station 11. The beginning 

of the disturbance at each station is seen simultaneously in the mean 

NASA/CR—2012-217656 115
 8 Copyright © 2003 by ASME 

0

0.05 0.042

7 7.5 8 8.5 9 9.5 10 10.5 11 t/T

Station

0

0.05 0.083

0

0.05 0.125

0

0.05 0.167

0

0.05 0.208

0

0.05 0.25

0

0.05 0.292

0

0.05 0.333

0

0.05 0.375

0

0.05 0.417

0

0.05 0.458

0

0.05 0.5

y/Ls

0

0.05 0.542

0

0.05 0.583

0

0.05 0.625

0

0.05 0.667

0

0.05 0.708

0

0.05 0.75

0

0.05 0.792

0

0.05 0.833

0

0.05 0.875

0

0.05 0.917

0

0.05 0.958

0 1 0 1 0 1 0 1 0 1 0 1 0 1 0 1 0 1

0

0.05 1

U/U

e

Fig. 9 Ensemble averaged dimensionless mean velocity ( UU e / ~ ) profiles 

NASA/CR—2012-217656 116
 9 Copyright © 2003 by ASME 

0

0.05 0.042

7 7.5 8 8.5 9 9.5 10 10.5 11 t/T

Station

0

0.05 0.083

0

0.05 0.125

0

0.05 0.167

0

0.05 0.208

0

0.05 0.25

0

0.05 0.292

0

0.05 0.333

0

0.05 0.375

0

0.05 0.417

0

0.05 0.458

0

0.05 0.5

y/Ls

0

0.05 0.542

0

0.05 0.583

0

0.05 0.625

0

0.05 0.667

0

0.05 0.708

0

0.05 0.75

0

0.05 0.792

0

0.05 0.833

0

0.05 0.875

0

0.05 0.917

0

0.05 0.958

0 0.2 0 0.2 0 0.2 0 0.2 0 0.2 0 0.2 0 0.2 0 0.2 0 0.2

0

0.05 1

u′/U

e

Fig. 10 Ensemble averaged rms value of dimensionless streamwise fluctuating velocity ( Uu e / ~′ ) profiles 

NASA/CR—2012-217656 117
 10 Copyright © 2003 by ASME 

0

0.05 0.042

7 7.5 8 8.5 9 9.5 10 10.5 11 t/T

Station

0

0.05 0.083

0

0.05 0.125

0

0.05 0.167

0

0.05 0.208

0

0.05 0.25

0

0.05 0.292

0

0.05 0.333

0

0.05 0.375

0

0.05 0.417

0

0.05 0.458

0

0.05 0.5

y/Ls

0

0.05 0.542

0

0.05 0.583

0

0.05 0.625

0

0.05 0.667

0

0.05 0.708

0

0.05 0.75

0

0.05 0.792

0

0.05 0.833

0

0.05 0.875

0

0.05 0.917

0

0.05 0.958

0 0 0 0 0 0 0 0 0

0

0.05 1

0.006 −u′v′/U

e

2

Fig. 11 Ensemble averaged dimensionless turbulent shear stress ( 2 Uvu e / ~~ − ′′ ) profiles 

NASA/CR—2012-217656 118
 11 Copyright © 2003 by ASME 

profile and the u

~′ and vu

~~− ′′ profiles of Figs. 9-11. At the trailing 

edge of the disturbance, return of the mean velocity profile to a smooth 

shape corresponds closely with the return of vu

~~− ′′ to near zero (see 

for example the profiles at Station 8 at t/T=0.833 or Station 10.5 at 

t/T=0.417). The u

~′ profiles, in contrast, take somewhat longer to 

return to an undisturbed condition. At Station 11, for example, there is 

at least a small near wall u

~′ peak at all times, while the vu

~~− ′′ values 

are essentially zero between t/T=0.625 and 0.792. 

 The flow structure at the downstream stations appears to be less 

distinct than at Stations 7 and 7.5. Inflection points are still present in 

the mean profiles, but the local minima and maxima in the mean 

profiles are less sharp. Still, the mean profile shapes and the multiple 

peaks in the u

~′ and vu

~~− ′′ profiles, particularly those far from the wall 

(see for example the Station 10 profiles of Figs. 10 and 11 between 

t/T=0.75 and 0.958) are evidence that the flow structures induced by 

the jets persist downstream. 

 After the disturbance caused by the jets passes, the boundary layer 

eventually separates at locations between Stations 8 and 11. At 

Station 8.5, for example, the trailing edge of the disturbance passes at 

about t/T=0.9. By t/T=0.1 the near wall profile appears separated. The 

separation appears to coincide with the return of u

~′ to near zero. The 

separation appears to persist until about t/T=0.6, after the start of the 

next disturbance event. The reattachment within the disturbance event 

corresponds to the motion of high u

~′ and vu

~~− ′′ into the near wall 

region at t/T=0.667. The same sequence of events is visible at 

downstream stations. At Station 10.5, for example, the trailing edge of 

the disturbance has passed by t/T=0.4, but the boundary layer remains 

attached until about t/T=0.75. Reattachment is visible at t/T=1, and a 

near wall peak emerges at the same time in vu

~~− ′′ . It should be noted 

that the separation bubble remains thin at all stations, and never begins 

to approach the thickness observed in the baseline flow of Fig. 7. 

 The period between the passage of the disturbance and boundary 

layer separation is believed to be a “calmed” region. Calmed regions 

have been observed in previous studies to follow turbulent spots and 

wake induced turbulent strips. The duration of the calmed region 

increases at the downstream stations, since the trailing edge celerity of 

the calmed region, shown in Fig. 9 to be about 0.3 the local freestream velocity (in agreement with previous studies of calmed 

regions), is slower than the trailing edge celerity of the disturbed 

region. Hence at Station 8.5 the calmed region extends for ∆t/T of 

about 0.2, while at Station 10.5 it extends for ∆t/T of about 0.4. The 

presence of the calmed region may help to limit the separation bubble 

thickness. 

 Figure 12 shows the ensemble u

~′ data in a different format. 

Contours of near wall u

~′ at y=0.095 mm (y/Ls=0.0004) are shown in a 

time-space plot. The horizontal axis shows the dimensionless 

streamwise location, and the vertical axis indicates the dimensionless 

time within the cycle. Two complete cycles are shown to better 

illustrate the periodicity of the event. The data in the bottom half of 

the figure is shown again in the top half. In this format, it is clear that 

the jet outpulse causes a high u

~′ event to appear at s/Ls=0.6 (Station 7) 

and t/T=0.6. This event then proceeds downstream in a widening 

wedge of turbulence. The lower and upper slopes of the wedge 

indicate its leading and trailing edge celerities, respectively. Within 

the wedge, a local u

~′ peak is visible at s/Ls=0.7 (Station 8). This peak 

indicates that the flow structure produced by the jets has penetrated 

very near the wall. Slightly farther downstream at s/Ls=0.75 (Station 

8.5) a local minimum is visible in the u

~′ contours. This may indicate 

that the flow structure induced by the jets has lifted off the wall 

slightly. The mean profiles of Fig. 9, support this, showing lower near 

0.005

0.01

0.015

0.02

0.65 0.7 0.75 0.8 0.85 0.9

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

t/T

s/Ls

Fig. 12 Time-space plot of ensemble averaged Uu e / ~′ at 

y/Ls=0.0004 

−2

−1

0

1

2

3

4

5

x 10−3

0.65 0.7 0.75 0.8 0.85 0.9

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

t/T

s/Ls

Fig. 13 Time-space plot of ensemble averaged 2 Uvu e / ~~ − ′′ at 

y/Ls=0.0128 

wall mean velocity gradients at Station 8.5, which indicates that high 

speed fluid is not as effectively brought into the near wall region at 

this station. Farther downstream in Fig. 12, the u

~′ level in the 

turbulent wedge rises again as a fully turbulent boundary layer begins 

to develop. Outside of the wedge the u

~′ level is very low, indicating 

laminar flow. 

 Figure 13 shows ensemble averaged vu

~~− ′′ contours at y/Ls=0.0128 

in the format of Fig. 12. As in Fig. 12, the wedge of turbulence is 

clear. Within the wedge momentum transport is high, but outside the 

wedge the flow appears to be laminar. 

 Figure 14 shows a time-space plot of the local ensemble averaged 

free-stream velocity. The free-stream velocity varies both spatially 

and temporally. Spatially, the shape of the airfoil passage causes the 

flow to decelerate. Temporally, when the jets cause a turbulent event 

to moves down the surface, it causes the boundary layer to thicken and 

accelerates the free-stream. During the laminar and calmed periods 

the boundary layer is thinner and the free-stream velocity is lower. 

Figure 15 shows the local Thwaites parameter, λθ, as computed from 

the free-stream velocity data of Fig. 14. In a laminar boundary layer, 

NASA/CR—2012-217656 119
 12 Copyright © 2003 by ASME 

1.05

1.1

1.15

1.2

1.25

1.3

0.65 0.7 0.75 0.8 0.85 0.9

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

s/Ls

t/T

1.1

1.1

1.15

1.15

1.15

1.2

1.2

1.2

1.2

1.2

1.25

1.25

1.25

1.25

1.25

1.25

1.3

1.3

1.3

1.3

1.3

1.2

1.3

Fig. 14 Time-space plot of ensemble averaged 

dimensionless free-stream velocity UU e / ~

∞

−2.5

−2

−1.5

−1

−0.5

0

0.5

1

1.5

0.65 0.7 0.75 0.8 0.85 0.9

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

s/Ls

t/T

−2

−1.5

−1.5

−1.5

−1.5

−1

−1

−1

−0.5

−0.5

−0.5

−0.5

0

0

0

0

0

0

0

0

0

0.5

−0.5

−0.5

0.5

0.5

0.5

−0.5

−0.5

0.5

0.5

−2

1

1

−2

−2

0

0

1

1

1

−2

−2

1

Fig. 15 Time-space plot of ensemble averaged Thwaites 

acceleration parameter, λθ

separation is expected when λθ is less than –0.082. With the exception 

of the turbulent strip, where the boundary layer is locally accelerated, 

λθ is below –0.082 at most times and locations on the surface, and at 

some times is below –2. Hence it is not a surprise that the boundary 

layer tends to separate when not controlled by turbulence or a calmed 

region. 

Integral Parameters 

 Local displacement and momentum thickness values can be 

computed from the ensemble mean velocity profiles of Fig. 9. 

Momentum thickness is shown in Fig. 16. Both δ

∗

 and θ grow in the 

streamwise direction and are about twice as large in the turbulent 

region than in the laminar flow region. The slow growth of the 

momentum thickness in the non-turbulent flow indicates that losses 

should be low in this region, as might be expected based on the low u

~′

and vu

~~− ′′ values of Figs. 10-11. At s/Ls=0.6 (Station 7) δ

*

 and θ in 

the laminar flow are both roughly 67% of their values in the baseline 

case of Volino [4]. The lower values in the present case may be due to 

the suppression of the large separation bubble downstream, which 

changes the local pressure gradient at Station 7. The suction during 

the inflow portion of the jet cycle could also be an explanation, but δ

*

and θ are uniformly low in the laminar flow at Station 7. If the suction 

were causing a thinner boundary layer, one would expect the effect to 

be stronger during the time of strongest inflow into the jet holes. 

Figure 17 shows the local shape factor H, computed from the δ

*

 and θ

values. Within the turbulent region, the shape factor remains between 

2 and 2.4, which is well above the expected value of 1.6 for a fullyturbulent boundary layer subject to the strong adverse pressure 

gradient of the present case. In the non-turbulent flow, H reaches 

values as high as 3.4, which is consistent with the observed flow 

separation. 

Skin Friction Coefficient 

 It is clear from the mean profiles of Fig. 9 that the local wall shear 

stress varies greatly during the jet oscillation cycle. At the upstream 

stations at t/T=0.6, for example, the jets bring high speed fluid very 

close to the wall, producing a very high mean velocity gradient at the 

wall. At other times at all stations, the boundary layer appears very 

laminar-like, indicating a relatively low wall shear stress. When the 

boundary layer is separated the wall shear goes to zero. Figure 18 

shows the local skin friction coefficients, Cf. The white regions 

indicate where the flow is separated and the local skin friction is 

essentially zero. At s/Ls=0.6 (Station 7) there is a laminar boundary 

layer with Cf=0.005 between jet outward pulses. The corresponding 

momentum thickness Reynolds is about 60. The skin friction drops in 

the streamwise direction and the boundary layer separates at s/Ls=0.7 

(Station 8). At t/T=0.6 and s/Ls=0.6, the skin friction is much higher, 

following the jet outpulse. The momentum thickness Reynolds 

number is as high as 180, and Cf reaches values as high as 0.014. 

Moving downstream along this turbulent region, Cf rapidly drops to 

0.001 at t/T=0.8 and s/Ls=0.75 (Station 8.5). The turbulence at this 

time and position is also low (Fig. 10). As stated above, the vortices 

produced by the jets appear to quickly bring high speed fluid near the 

wall, resulting in high skin friction and turbulence, but by s/Ls=0.75 

this effect may weaken, resulting in lower Cf and u

~′ . Moving farther 

downstream to t/T=1.1 and s/Ls=0.85, Cf rises again to as high as 

0.007. The corresponding Reθ is about 200 at this time and location. 

Cf=0.007 is about what one would expect for a fully turbulent 

boundary layer with Reθ=200, based on the standard correlation 

Cf=0.0256/Reθ

0.25 (Schlichting [33]). 

 Figure 18 clearly shows the calmed region described above. It is the 

triangular shaped region centered at t/T=0.4 and s/Ls=0.9 that in the 

figure is bounded below by the strip of high Cf and above by the 

separated flow region. Comparing Figs. 12, 13, 16 and 18 shows that 

the calmed region has low skin friction, low momentum thickness and 

low turbulence. Losses should, therefore, be low for the calmed 

region, as expected. 

 A comparison of the momentum thickness and skin friction in Figs. 

16 and 18 is interesting. The wedge of high θ overlaps both the upper 

portion of the separated flow region and the strip of high Cf. As 

shown in the profiles of Figs. 10 and 11, downstream of s/Ls=0.75 

(Station 8.5), the turbulence and mixing induced by the jet outpulse 

initially affects the outer part of the boundary layer and does not 

immediately eliminate the separation bubble. Figure 16 shows that 

this outer region mixing causes a rise in the momentum thickness. It is 

only somewhat later and farther downstream that this mixing moves 

into the near wall region, causing reattachment and high Cf. Kaszeta et 

al. [34] reported a similar result for flow over the Pak-B airfoil subject 

to wake passing events. They observed a time lag between the wake 

arrival and near wall transition. 

NASA/CR—2012-217656 120
 13 Copyright © 2003 by ASME 

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

0.65 0.7 0.75 0.8 0.85 0.9

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

s/Ls

t/T

0.6

0.6

0.8

0.8

0.8

0.8

0.8

0.8

1

1

1

1

1

1.2

1.2

1.2

1.2

1

1

1

1

1.2

1.2

1.2

1.4

1.4

1.4

1.4

1.4

1.4 1.4

1.6

1.6 1.6

1.6

1 1.2

1.6

1.6 1.6

1.4

1.8

1.8

.6

1.8

.8

1

1

2

2

2

2

2

Fig. 16 Time-space plot of ensemble averaged momentum 

thickness, θ, in mm

1.8

2

2.2

2.4

2.6

2.8

3

3.2

0.65 0.7 0.75 0.8 0.85 0.9

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

s/Ls

t/T

2

2

2

2

2.2 2.2

2.2

2.2

2.2

2.2

2.2

2.4

2.4

2.4

2.4

2.4

2.4

2.4

2.4

.4

2.6

2.6

2.6

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2.6

2.6

2.8

2.8

2.8

2.8

2.8

2.8

.8

3 3

3

3

3

3

3.2

3.2

3.2

3.2

2.4

2.6

2

2.8

1.8

1.8

3.4

3.4

2.2

2.2

1.8

1.8

2.4

2.4

3

Fig. 17 Time-space plot of ensemble averaged shape 

factor, H 

1

2

3

4

5

6

7

8

9

x 10−3

0.65 0.7 0.75 0.8 0.85 0.9

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

s/Ls

t/T

0.001

0.001

0.001

0.001

0.001

0.001

0.002

0.002

0.002

0.002

0.0006

0.0006

0.0006

0.0006

0.0006

0.0006

0.001

0.001

0.002 0.002

0.002

.002

0.002

0.003

0.003

0.003

0.003

0.003

0.003

0.003

0.003

0.004

0.004

0.004

0.004

0.005

0.005

0.005

0.004

0.004

0.004

0.004

0.006

0.006 0.007

0.007

0.005

0.005

.008

0.008

0.002

0.009

0.009 0.004

0.01

0.01

0.003

0.006

0.006

Fig. 18 Time-space plot of ensemble averaged skin friction 

coefficient, Cf; white areas indicate separated flow

 Kaszeta et al. [34] also observed a relation between the temporal 

acceleration and deceleration of the flow associated with the wake 

passing, and its relation to transition and the thickening and thinning 

of the boundary layer. Similarly, a comparison of Figs. 15, 16 and 18 

show the relation between the local acceleration, θ and Cf. The region 

of overlap between high θ and separated flow corresponds to the 

region of strong deceleration centered at s/Ls=0.9 and t/T=0.85. One 

could argue that the deceleration inhibits reattachment in spite of the 

mixing in the outer part of the boundary layer. At slightly later t/T, the 

flow is accelerated. Under the combination of acceleration and higher 

vu

~~− ′′ the boundary layer reattaches. 

Open Questions 

 The first objective of the present study was to assess the 

effectiveness of synthetic VGJs on separation control for an LPT 

airfoil and provide some description of the mechanisms through which 

the jets work. This has been addressed above. The second objective 

was to use these results to generate questions for further consideration. 

These are discussed below. 

 What is the effect of jet frequency on flow control? In the present 

case, with F+=0.65, the airfoil boundary layer at any given time was 

approximately half disturbed by the jets and half laminar or calmed. 

The effects of one jet outpulse are present in the boundary layer at all 

times. Separation occurred, but the separation bubble was small and 

did not appear to result in any harmful effects. How thick a separation 

bubble is tolerable? If F+ were increased, the duration of each event 

would be reduced, but more events would be present in the boundary 

layer at any given time. It is not clear what effect this would have. In 

other studies with synthetic jets, some have found that F+ of the order 

1, as in the present study, is most effective (e.g. Seifert and Pack [35]). 

Others, such as Amitay and Glezer [21] report that under some 

conditions jets with F+ of the order 10 are more effective. Lowering 

the jet frequency might be beneficial. Losses appear to be generated 

primarily in the flow disturbed by the jets, while the flow between 

these events is laminar or calmed with low losses. Reducing F+ could 

presumably result in a smaller fraction of the airfoil covered by 

disturbed flow at any time. At some point, however, if the jets are too 

widely spaced, the flow will become uncontrolled at times and revert 

to the large separation bubble of the baseline case. Based on the 

trailing edge celerity of the disturbed flow, if F+ were reduced below 

0.45, there would be instances within the jet cycle when no 

disturbances would be present in the boundary layer. If F+ were 

reduced below 0.3, there would be times when neither disturbed or 

calmed flow would be present. In flow around a single airfoil, Seifert 

and Pack [35] found that jets with F+ between 0.5 and 1.5 were most 

effective at all Re, but that with F+=0.25 the jets were ineffective. 

Bons et al. [25], in contrast, found that their VGJ’s were effective at 

F+ as low as 0.1. They suggested that the more controlled nature of 

the LPT flow, where adjacent airfoils provide covered turning, might 

explain the lower effective F+ in their study. It should be noted that 

the Bons et al. [25] experiments were conducted at a Reynolds number 

of 60,000, where separation effects are not as severe as in the present 

case with Re=25,000. 

 What is the effect of jet amplitude on flow control? The pressure 

profiles of Fig. 4 indicate that the jet amplitude must be sufficiently 

high for the jets to be effective. It is expected that using a higher 

amplitude than necessary will result in higher losses, but the extent to 

which the losses would increase is not known. Volino [5] using 

passive bars found that cases with a small separation bubble followed 

by reattachment had lower losses than cases in which the boundary 

layer was tripped to turbulent to prevent separation. Low amplitude 

jets could potentially produce a similar effect. 

NASA/CR—2012-217656 121
 14 Copyright © 2003 by ASME 

 What is the effect of jet waveform? In the present case the speaker 

was driven with a sine wave input. A square wave with a short duty 

cycle, as in the study of Bons et al. [25], might be better. By keeping 

the duty cycle short, the amount of calmed flow relative to disturbed 

flow could be increased. This might allow a reduction of losses 

without a sacrifice of separation control. 

 What is the effect of jet inflow? The outward flow portion of each 

jet cycle appears to dominate the flow control. It is possible, however, 

that the inflow may help reduce the boundary layer thickness in the 

undisturbed flow periods and help reduce separation effects. A direct 

comparison of cases at the same Re using synthetic VGJs and pulsed 

VGJs could help answer this question. 

 What is the effect of jet geometry? The VGJs are clearly effective. 

Through the generation of streamwise vortices they appear to provide 

more mixing than would be produced with normal jets that simply 

produced turbulence. The relative magnitudes of the effects of 

turbulence and streamwise vortices in enhancing mixing are uncertain. 

Might the generation of turbulent spots with a normal jet be sufficient? 

Which would result in lower losses? 

 What are the effect of Reynolds number and free-stream turbulence 

level? The baseline cases of Volino [4] show that at higher Re and 

higher FSTI the boundary layer is more likely to transition and reattach 

even without flow control. The generation of calmed attached flow 

between jet events could still prove beneficial at higher Re, however. 

There would be a tradeoff between the losses generated by the jets 

themselves and the reduction in losses the jets might provide by 

reducing the separation bubble thickness and producing low-loss 

calmed regions. It might be possible to control the flow at higher Re

with significantly lower blowing ratios than in the present case. 

Natural transition in the undisturbed flow between jet events and its 

interaction with calmed regions would be an added complication at 

higher Re not seen in the present study. If the Re were sufficiently 

high so that separation did not occur in the uncontrolled case, it is 

unlikely that the jets would provide any benefit. Schulte and Hodson 

[27] noted that the presence of calmed regions produced by unsteady 

wake passing could not significantly lower the losses in an already 

attached boundary layer. 

CONCLUSIONS 

1. Synthetic vortex generator jets proved effective for controlling 

boundary layer separation on an LPT airfoil at very low Reynolds 

numbers. The separation bubble was effectively eliminated, and losses 

were lower than in a similar case with passive flow control. 

2. The VGJs prevent separation by bringing high momentum fluid 

into the near wall region and by promoting momentum transport 

through turbulent mixing. 

3. The disturbance produced by the VGJs behaves in many ways 

like the disturbance associated with a turbulent spot or a wake induced 

turbulent strip. The leading edge celerity of the disturbance is 

approximately 0.9U∞, and the trailing edge celerity is about 0.45U∞. A 

calmed region with a trailing edge celerity of 0.3U∞ follows the 

disturbance. The calmed region is resistant to separation. 

4. The adverse pressure gradient in the present case was strong 

enough so that the boundary layer did separated after the passage of 

the calmed flow. The separation bubble remained thin, however. The 

appearance of a disturbance did not immediately induce reattachment. 

The disturbance appeared initially in the outer part of the boundary 

layer. After some lag time the disturbance spread into the near wall 

region and caused reattachment. 

5. While much has been learned regarding the effectiveness and 

physics of synthetic VGJs in LPT flows, many questions remain 

regarding their applicability under different flow conditions and their 

optimal design. These questions have been discussed. 

ACKNOWLEDGEMENTS 

 This work was sponsored by the NASA Glenn Research Center. 

The grant monitor is Dr. David Ashpis. Additional matching support 

was provided through a U.S. Naval Academy Recognition Grant. Mr. 

Dale Boyer of the Technical Support Department at the Naval 

Academy fabricated the test section and the airfoils. Midshipman Dirk 

Heron conducted preliminary tests which led into the present study. 

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[14] McManus, K., Legner, H., and Davis, S., 1994, “Pulsed Vortex 

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Paper 94-2218. 

[15] Raghunathan, S., Watterson, J., Cooper, R., and Lee, S., 1999, 

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[16] Sinha, S.K., and Pal, D., 1993, “Optimizing the Use of Acoustic 

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NASA/CR—2012-217656 122
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[17] Jacobson, S.A., and Reynolds, W.C., 1998, “Active Control of 

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Dynamics of LPT Blade Separation Control Using Pulsed Jets,” 

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Melwani, N., 1997, “Investigation of the Calmed Region Behind 

a Turbulent Spot,” ASME Journal of Turbomachinery, 119, pp. 

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[27] Schulte, V., and Hodson, H.P., 1998, “Prediction of the Becalmed 

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[31] Volino, R.J., and Simon, T.W., 1997, “Velocity and Temperature 

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Pressure Gradients,” ASME Journal of Heat Transfer, 119, pp. 

433-439. 

[32] Ligrani, P.M., and Bradshaw, P., 1987, “Spatial Resolution and 

Measurement of Turbulence in the Viscous Sublayer Using 

Subminiature Hot-Wire Probes,” Experiments in Fluids, 5, pp. 

407-417. 

[33] Schlichting, H., 1979, Boundary Layer Theory, 7th ed. McGrawHill, New York. 

[34] Kaszeta, R.W., Simon, T.W., and Ashpis, D.E., 2001, 

“Experimental Investigation of Transition to Turbulence as 

Affected by Passing Wakes,” ASME Paper 2001-GT-195. 

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Separation at High Reynolds Numbers,” AIAA Journal, 37, pp. 

1062-1071. 

NASA/CR—2012-217656 123
1 Copyright © 2004 by ASME 

Proceedings of ASME Turbo Expo 2004 

Power for Land, Sea, and Air 

June 14-17, 2004, Vienna, Austria

GT2004-53360 

SEPARATED FLOW TRANSITION MECHANISM AND PREDICTION WITH HIGH AND 

LOW FREESTREAM TURBULENCE UNDER LOW PRESSURE TURBINE CONDITIONS

Ralph J. Volino and Douglas G. Bohl

Department of Mechanical Engineering 

United States Naval Academy 

Annapolis, Maryland 21402 

Email: volino@usna.edu

ABSTRACT 

 A correlation for separated flow transition has been developed for 

boundary layers subject to initial acceleration followed by an 

unfavorable pressure gradient. The correlation is based on the 

measured growth of small disturbances in the pre-transitional 

boundary layer. These disturbances were identified and quantified 

through spectral analysis of the wall normal component of velocity. 

Cases typical of low pressure turbine airfoil conditions, with Reynolds 

numbers (Re) ranging from 25,000 to 300,000 (based on suction 

surface length and exit velocity) were considered at low (0.5%) and 

high (8.7% inlet) freestream turbulence levels. In some cases, twodimensional rectangular bars were placed at the beginning of the 

adverse pressure gradient region as passive flow control devices. The 

dimensionless magnitude of the initial disturbance which begins to 

grow at the suction peak depends on the freestream turbulence level 

and the size of any bar applied to the surface. The growth rate 

depends on the Reynolds number. When the pre-transitional 

disturbances grow to a sufficient magnitude, transition begins. The 

new correlation is based on the physics observed in the turbulence 

spectra, but allows transition prediction using only the Reynolds 

number, freestream turbulence level and bar height. The correlation 

has been checked against experimental data from the literature, and 

allows transition location prediction to within the uncertainty of the 

experimental measurements. The correlation represents an 

improvement over previous correlations which accounted for Reynolds 

number or freestream turbulence effects, but not both. 

NOMENCLATURE 

Cp ( ) 2

T − ρU/PP2 e , pressure coefficient 

d bar height 

FSTI freestream turbulence intensity, based on rms u′ , v′ , or 

combined u′ and v′ , and normalized using U∞, [%] 

K ) /)(/( dsdUU 2 ν ∞∞ , acceleration parameter 

Ls suction surface length 

P pressure 

PT upstream stagnation pressure

Re LU se /ν , exit Reynolds number 

Rept Up(st-sp)/ν, suction peak to transition Reynolds number

Rest Us(st-ss)/ν, separation to transition Reynolds number

Reθ momentum thickness Reynolds number 

s streamwise coordinate, distance from leading edge 

U freestream velocity 

U∞ local freestream velocity 

Ue nominal exit freestream velocity, based on inviscid solution 

u mean streamwise velocity 

u′ streamwise fluctuating velocity 

v′ wall normal fluctuating velocity 

− ′vu ′ time averaged turbulent shear stress 

y cross-stream coordinate, distance from wall 

γ intermittency, fraction of time flow is turbulent 

ν kinematic viscosity 

ρ density 

θ momentum thickness 

Subscripts 

b baseline flow 

m modified flow 

p suction surface pressure minimum, velocity maximum 

s separation location 

t transition start location 

INTRODUCTION 

 Boundary layer separation is a known problem on some modern 

low-pressure turbine (LPT) airfoils, due to the strong adverse pressure 

gradients created when designers impose higher loading in an effort to 

improve efficiency and lower cost by reducing airfoil count in engines. 

Separation bubbles, particularly those which fail to reattach, can result 

in a significant loss of lift and a subsequent degradation of engine 

efficiency (e.g. Hourmouziadis [1], Mayle [2], and Sharma et al. [3]). 

The problem is particularly relevant in aircraft engines. Airfoils 

optimized to produce maximum power under takeoff conditions may 

still experience boundary layer separation at cruise conditions, due to 

NASA/CR—2012-217656 124
2 Copyright © 2004 by ASME 

the lower density and therefore lower Reynolds numbers at altitude. A 

component efficiency drop of 2% may occur between takeoff and 

cruise conditions in large commercial transport engines, and the 

difference could be as large as 7% in smaller engines operating at 

higher altitudes [4, 5]. Prediction and control of suction side 

separation, without sacrifice of the benefits of higher loading, is 

therefore, crucial for improved engine design. 

 Separation on airfoils is complicated by boundary layer transition. 

Separated flow transition in the LPT has been the focus of several 

recent studies. Volino [6] provides a review and describes as follows 

the transition process on the suction side of a typical LPT airfoil. The 

strong acceleration on the leading section of the airfoil keeps the 

boundary layer thin and laminar, even in the presence of elevated 

freestream turbulence. The favorable pressure gradient keeps the 

boundary layer stable to small disturbances, but immediately 

downstream of the suction peak, the boundary layer becomes unstable. 

In most cases Volino [6] observed that the boundary layer separated 

just downstream of the suction peak. Transition then occurred in the 

shear layer over the separation bubble and caused the boundary layer 

to reattach. The transition location moved upstream as the freestream 

turbulence intensity (FSTI) or Reynolds number was increased. At the 

highest FSTI and Re, transition occurred far enough upstream to 

prevent separation. At the lowest Re, transition did not occur in the 

shear layer and the boundary layer did not reattach. 

 Separated flow transition under low FSTI conditions appears to 

occur through instabilities which lead to the growth of small 

disturbances. Hence it is a natural transition, as opposed to a bypass 

transition. The shear layer is subject to both global and convective 

instability, so breakdowns due to Kelvin-Helmholtz (KH) or TollmienSchichting (TS) type disturbances are possible. One or the other may 

be dominant in particular cases. Hatman and Wang [7] and Lou and 

Hourmouziadis [8] suspected KH type instabilities in their 

experiments. Volino [9] documented turbulence spectra in pretransitional shear layers and found that the frequencies of the spectral 

peaks closely matched the most unstable frequencies expected for TS 

waves in the boundary layer just upstream of separation. Under high 

FSTI conditions, Volino and Hultgren [5] found that the shear layer 

spectra were more broad banded and did not exhibit the sharp peaks 

observed in low FSTI cases. They concluded that transition was 

through a bypass mode induced by the freestream disturbance. Under 

favorable pressure gradients, transition under high FSTI can occur 

even when stability analysis indicates that the boundary layer is stable 

to small disturbances, clearly indicating a bypass mechanism. Recent 

evidence, however, suggests that with adverse pressure gradients, 

natural transition is important even with high FSTI. Hughes and 

Walker [10] used wavelet analysis to document TS frequencies in an 

adverse pressure gradient boundary layer. These TS waves were clear 

even in the presence of passing wakes, which raised the FSTI to about 

8%. Volino [9] noted that although his spectral peaks were more 

broad banded with high FSTI, they were centered at the same 

frequencies as the sharp peaks in otherwise similar low FSTI cases, 

and could be observed to grow from initially very low magnitudes in 

the pre-transition region. 

 Volino [11] provides further evidence of the importance of natural 

transition in a study of passive flow control using rectangular bars on 

the suction surface of an LPT airfoil. The bars were thin, ranging from 

about 0.1 to 1 boundary layer thickness in height, and were placed at 

the beginning of the adverse pressure gradient region. The thickest 

bars tripped the boundary layer to turbulent and prevented separation. 

Thinner bars appeared to have no immediate effect, and both the mean 

velocity and rms turbulence quantities just downstream of the bars 

were indistinguishable from those in baseline cases without bars. 

Farther downstream, however, transition and reattachment occurred 

sooner in the cases with bars than in the unmodified flow, with 

transition moving upstream as bar height increased. The bars appeared 

to impart a small disturbance that was too small to detect at first, but 

that grew until it eventually induced transition. This suggests a natural 

transition mechanism. The same behavior was observed under both 

low and high FSTI, indicating that although high FSTI does help to 

promote transition, it is not purely a freestream induced bypass 

transition. 

 Studies such as those noted above provide important clues about 

separated flow transition, but the transition mechanism is still not fully 

understood. The rate at which disturbances grow has not been 

quantified or related in a direct way to the prediction of transition. 

Existing correlations, such as those of Roberts [12], Davis et al. [13], 

Mayle [2], and Hatman and Wang [7], provide reasonable estimates of 

transition start location, but they do not always agree with each other 

and are not always as accurate or robust as desired. Improved 

computational models and correlations are desired, and incorporation 

of the physics from recent observations may help. 

 The objective of the present study is to quantify the growth of 

disturbances in separated boundary layers under a range of conditions 

that are representative of LPT airfoil suction side flows, and to use this 

information to develop a correlation for transition prediction. Data 

from the studies of Volino [6, 9, 11] along with new experimental data 

will be used. After a correlation is developed, it will be compared to 

other data sets from the literature. 

EXPERIMENTS

 Experiments were conducted in a low speed wind tunnel, described 

by Volino et al. [14]. Briefly, air enters through blowers and passes 

through a honeycomb, a series of screens, two settling chambers, and a 

three-dimensional contraction before entering the test section. At the 

exit of the contraction, the mean velocity is uniform to within 1%. 

The FSTI is 0.5%±0.05%. Nearly all of this freestream “turbulence” 

is actually streamwise unsteadiness at frequencies below 20 Hz and is 

not associated with turbulent eddies. The rms intensities of the three 

components of the unsteadiness are 0.7%, 0.2% and 0.2% in the 

streamwise, pitchwise and spanwise directions, respectively. For low 

FSTI cases, the test section immediately follows the contraction. For 

high FSTI, a passive grid is installed at the contraction exit followed 

by a 1 m long rectangular settling chamber. At the inlet to the test 

section (10 cm upstream of the plane containing the leading edges of 

the airfoils) the high FSTI mean flow and turbulence are spatially 

uniform to within 3% and 6% respectively (FSTI=8.7%±0.5%). The 

freestream turbulence is nearly isotropic with rms intensities of 8.8%, 

8.9% and 8.3% in the streamwise, pitchwise and spanwise directions. 

The integral length scales of these components are 3 cm, 1.6 cm and 

1.4 cm. The integral scales were computed from the power spectra of 

each component. 

 The test section, shown in Fig. 1, consists of the passage between 

two airfoils. Details are listed in Table 1 and more information is 

available in Volino [6]. A large span to chord ratio of 4.3 was chosen 

to insure two-dimensional flow at the spanwise centerline of the 

airfoils, where all measurements were made. Upstream of each airfoil 

are flaps, which control the amount of bleed air allowed to escape 

from the passage. The flaps, along with a tailboard on the pressure 

side of the passage, are adjusted to produce the correct leading edge 

flow and pressure gradient along the airfoils. The flow in the passage 

matches that in a multi-blade cascade. The geometry of the passage 

corresponds to that of the Pak-B airfoil, which is an industry supplied 

research airfoil that is representative of a modern LPT design. It has 

been used in several studies, as noted in Volino [6]. 

NASA/CR—2012-217656 125
3 Copyright © 2004 by ASME 

tailboard

flap

flap bleed air

bleed air

35o

60o

Fig. 1 Schematic of the test section 

 Data representing 40 experimental conditions are used. Baseline 

data are from Volino [6, 9], who considered high and low FSTI cases 

at five Reynolds numbers (Re=25,000, 50,000, 100,000, 200,000, and 

300,000). Reynolds numbers are based on the exit velocity from the 

passage and the suction surface length, Ls. The Reynolds number 

range is representative of conditions from cruise to takeoff. The FSTI 

levels in an engine may vary considerably, but the values in the 

present work are believed to span the range of most interest. Data 

from cases with passive-bar flow control are from Volino [11]. The 

bars were of uniform rectangular cross section and extended along the 

airfoil, as shown in Fig. 2. The trailing edge of the bar was located at 

s/Ls=0.51, near the suction surface velocity peak. All bars were 6 mm 

wide in the streamwise direction. Bar heights of 0.4 mm, 0.8 mm and 

1.6 mm were used. The bar heights were all less than 1% of Ls. They 

compare to local boundary layer thickness at the bar location of about 

3.8 mm, 2.7 mm, 2.0 mm, 1.4 mm and 1.2 mm in the baseline 

Re=25,000 through 300,000 cases respectively. For each bar height, 

all 10 cases of the baseline study were re-documented, for a total of 30 

new experimental cases. 

Measurements 

 Pressure surveys were made for each case using pressure taps 

located along the spanwise centerline of each airfoil. Locations of the 

taps on the suction side are listed in Table 2 along with measured local 

FSTI components, and U∞/Ue and the ReK product at these stations 

based on a non-separating, inviscid solution. The uncertainty in Cp is 

7% at Re=25,000, and 4% in other cases. Further details of these 

measurements and their uncertainties are available in Volino [6, 11]. 

 Profiles of the streamwise velocity component in the suction side 

boundary layer were measured for all cases at streamwise stations 

corresponding to pressure taps 7-11, as given in Table 2. These 

stations are in the adverse pressure gradient region. Profiles at 

Stations 1-6 were fully documented for the baseline cases, and show 

that the upstream boundary layer closely follows a laminar solution, 

even in the high FSTI cases. Further details of these measurements are 

available in Volino [6, 11]. 

 The intermittency, γ, is the fraction of time the flow is turbulent 

within the transition region. It was determined at each measurement 

location based on the instantaneous streamwise velocity signal, using 

the technique described in Volino et al. [14]. Turbulent flow is 

defined here to include a range of large and small scale eddies, 

turbulence production, and dissipation. A boundary layer may be 

characterized by significant u′ fluctuations but still be non-turbulent if 

Fig. 2 Scale drawing of suction side airfoil showing 

location of bar 

Table 1: Test section parameters 

Axial 

Chord 

[mm] 

True 

Chord 

[mm] 

Pitch 

[mm] 

Span 

[mm] 

Suction 

side, Ls

[mm] 

Inlet 

flow 

angle 

Exit 

flow 

angle 

inlet

e

U

U

153.6 170.4 136.0 660.4 228.6 35° 60° 1.64 

Table 2: Measurement stations locations, local acceleration 

(inviscid soln.), and measured local freestream turbulence 

Station s/Ls

Ue

U∞

inviscid

ReK 

inviscid 

Low 

FSTI

∞

′

U

u 2

 [%] 

Low 

FSTI

∞

′

U

v 2

[%] 

High 

FSTI

∞

′

U

u 2

 [%] 

High 

FSTI

∞

′

U

v 2

[%] 

1 0.111 0.923 1.58 0.44 5.2 

2 0.194 1.007 1.20 0.39 4.6 

3 0.278 1.123 0.86 0.37 4.0 

4 0.361 1.190 0.75 0.38 3.5 

5 0.444 1.247 0.62 0.39 3.2 

6 0.528 1.277 -0.02 0.41 2.8 

7 0.611 1.240 -0.81 0.47 0.05 2.9 5.9 

8 0.694 1.161 -0.95 0.47 0.12 3.0 6.2 

9 0.777 1.100 -0.58 0.48 0.14 3.4 6.6 

10 0.861 1.053 -0.53 0.54 0.11 3.8 6.8 

11 0.944 1.020 -0.18 0.51 0.11 4.0 6.8 

these fluctuations are induced by an external source that does not also 

cause near wall turbulence production. Such is often the case under 

high FSTI conditions. Freestream eddies buffet the boundary layer, 

inducing non-turbulent boundary layer fluctuations but very little 

momentum transport. Transition to turbulence is characterized not so 

much by large increases in u′ levels, which may remain essentially 

constant, but by the appearance of higher frequencies. The higher 

frequencies signal the generation of turbulence in the near wall region 

and are used to distinguish between turbulent and non-turbulent flow. 

 The measurements of most significance for the present study are 

profiles of the wall normal fluctuating velocity, v′ . One objective of 

the present study is to track the growth of small disturbances in the 

pre-transitional flow. Volino [9] found that small disturbances at 

particular frequencies can be masked by the broadband unsteadiness in 

a non-turbulent boundary layer, particularly under high FSTI 

conditions. The broadband unsteadiness tends to be much higher in 

u′ than in v′ , making disturbances in v′ easier to detect. Wall 

normal velocity was measured for the baseline cases by Volino [6]. 

New data were acquired in the present study for the passive bar cases. 

A hot wire anemometer (AA Lab Systems model AN-1003) and a 

NASA/CR—2012-217656 126
4 Copyright © 2004 by ASME 

boundary layer cross-wire probe (TSI model 1243-T1.5) were used to 

measure profiles of the wall normal velocity and turbulent shear stress 

at Stations 7-11 for each case. Profiles were measured near but not at 

the spanwise centerline of the airfoil to insure that the pressure taps 

did not interfere with the velocity measurements. At each 

measurement location, data were acquired for 26 seconds at a 20 kHz 

sampling rate (219 samples). All raw data were saved. The high 

sampling rate provides an essentially continuous signal, which is 

needed for intermittency and spectral post-processing. Data were 

acquired at 25 locations in each profile, beginning 1 mm from the wall 

and extending to the freestream. Just upstream of separation, y=1 mm 

is equivalent to y

+≈10 when Re=25,000 and y+≈60 when Re=300,000. 

Uncertainties in 2 v′ and the turbulent shear stress, − vu ′′ , are 10%. 

 Power spectra of v′ were computed for the data from all 

measurement locations. Frequencies are resolved from 4.88 Hz to 10 

kHz in 4.88 Hz increments using a 4096 point Fast Fourier Transform 

to compute the spectra. As a check, the spectra were integrated with 

respect to frequency and found to equal the corresponding time 

averaged Reynolds stresses. The average uncertainties in the spectra 

are 10%. Measurements with the cross-wire probe are subject to 

spatial averaging errors due to the length of the sensors (1.27 mm 

active length) and the spacing between the two sensors (1 mm). 

Ligrani and Bradshaw [15] and Ligrani et al. [16] investigated spatial 

averaging. Their results, when applied to the present cases, indicate 

that spatial averaging should not result in significant error in average 

quantities at locations farther from the wall than 1 mm. As explained 

by Ligrani and Bradshaw [15], however, spatial averaging effects will 

not be uniform across the frequency spectrum, and will be most severe 

for the smallest scales (highest frequencies) in the flow. Applying the 

spectral results of Ligrani and Bradshaw [15] to the present study, 

errors due to spatial averaging will rise above 10% at frequencies 

above 150, 300, 600, 1200 and 1800 Hz for the Re=25,000, 50,000, 

100,000, 200,000, and 300,000 cases respectively. Below these 

frequencies the errors should be under 10%. These frequencies are all 

above the frequencies of the spectral peaks in the results presented 

below, so the peaks should not be significantly attenuated. 

RESULTS 

Velocity and Pressure Profiles

 Examples of the pressure and velocity profiles for the various cases 

are shown in Figs. 3-6. Figure 3 shows pressure profiles for the low 

FSTI, Re=50,000 cases. On the pressure side of the airfoil, the data 

from all cases agree with the inviscid solution for flow through the 

passage. This is expected since the boundary layer on most of the 

pressure side is subject to a favorable pressure gradient, and is 

attached. Agreement with the inviscid solution is also good on the 

upstream portion of the suction side. In the baseline case, the data dips 

below the inviscid solution at the suction peak (indicating a loss of 

lift), and then exhibits a plateau of nearly constant value in the adverse 

pressure gradient region. This plateau indicates that the boundary 

layer is separated and does not reattach. The behavior is essentially 

the same with the 0.4 mm thick bar, indicating that this bar is too small 

to induce transition and reattachment before the trailing edge. With 

the 0.8 mm thick bar there is still a large separation bubble, but the 

pressure coefficient drops toward the inviscid solution at the last 

measurement location. This indicates reattachment just before the 

trailing edge. With the largest bar, there is still a separation, but the 

reattachment moves farther upstream. 

 Figure 4 shows the velocity profiles for the cases presented in Fig. 

3. The top row shows the mean velocity at Stations 7-11. At Station 

7, the baseline, 0.4 mm and 0.8 mm bar cases are virtually 

0 0.2 0.4 0.6 0.8 1 1.2 0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

baseline

s/Ls

Cp

0.4 mm bar

0.8 mm bar

1.6 mm bar

inviscid solution

Low FSTI

Re=50,000

Fig. 3 Cp profiles, low FSTI, Re=50,000 cases 

0 0 0 0 0 1 0

0.05

u/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

rms u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 0

0.05

rms v′/Ue

y/Ls

c) 7 8 9 10 11

0.2

0 0 0 0 0 0

0.05

−u′v

′

/Ue

2

y/Ls

d) 7 8 9 10 11

0.01

0 0 0 0 0 1 0

0.05

γ

y/Ls

e) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 4 Station 7-11 profiles, low FSTI, Re=50,000 cases: (a) 

mean velocity, (b) rms u′/Ue, (c) rms v′/Ue, (d) -u′v′/Ue

2

, (e) 

intermittency 

indistinguishable, while the 1.6 mm bar case shows a separation 

bubble. The rms streamwise velocity, u′ , is very low at Station 7, 

with the exception of a small peak in the shear layer for the 1.6 mm 

bar case. The same is true for rms v′ . The turbulent shear stress is 

essentially zero for all cases. The intermittency is also near zero, 

indicating laminar flow in all cases. At Station 8, the baseline and 

smaller bar cases are still essentially indistinguishable. The boundary 

layers have separated in these cases and there is a small peak in u′ in 

the shear layer, but v′ , the turbulent shear stress and the intermittency 

all remain near zero. With the largest bar, the boundary layer has 

undergone transition and reattached. At Station 9, a small disturbance 

appears in the turbulence quantities of the 0.8 mm bar case, but the 

NASA/CR—2012-217656 127
5 Copyright © 2004 by ASME 

intermittency remains low. The baseline and 0.4 mm bar cases remain 

indistinguishable, with a peak in u′ in the shear layer, but v′ and 

− vu ′′ values near zero. By Station 10, the boundary layer is turbulent 

in the 0.8 mm bar case, and has reattached. In the 0.4 mm bar case, 

the intermittency is still zero, but u′ has begun to rise in the near wall 

region, and v′ and − vu ′′ have risen as well. This indicates the 

beginning of transition, and the boundary layer is turbulent and 

reattached in the 0.4 mm bar case by Station 11. In the baseline case, 

the shear layer does not transition or reattach before the trailing edge. 

 Several conclusions can be drawn from the information in Figs. 3 

and 4. The reattachment locations observed in the mean velocity 

profiles agree with the locations in the pressure profiles. The 

beginning of reattachment coincides closely with the beginning of 

transition (rise of γ above zero). The rise in u′ in the shear layer does 

not indicate turbulence in the sense defined above, but only 

streamwise unsteadiness. This is evidenced by the near zero values of 

v′ and − vu ′′ corresponding to the u′ peaks. High u′ does not 

necessarily imply significant momentum transport, and it is 

momentum transport which characterizes turbulent mixing and 

promotes reattachment of the boundary layer. The v′ and − ′vu ′

values begin to rise simultaneously, signaling the beginning of 

transition. None of the bars were large enough to immediately trip the 

boundary layer to turbulent at this Reynolds number, and the smaller 

bars appeared to have no immediate effect on the profiles. The smaller 

bars must have imparted a small disturbance in the flow, however, 

because the transition location moved upstream from its location in the 

baseline case. 

 Figures 5 and 6 show pressure and velocity profiles for the high 

FSTI, Re=50,000 cases. The beginning of transition and reattachment 

in each case are between one and two stations farther upstream than in 

the corresponding low FSTI cases of Figs. 3 and 4. The u′ peaks are 

much higher than in the low FSTI cases, as is the freestream v′ . The 

near wall v′ and − vu ′′ , however, remain low until transition starts, in 

agreement with the low FSTI cases. As in the low FSTI cases, high 

u′ does not necessarily imply significant turbulent mixing. The effect 

of the bars in the low and high FSTI cases is similar. The smaller bars 

appear to have no immediate effect on the profiles, but they must 

impart a small disturbance which grows and causes transition to move 

upstream of its location in the baseline case. Hence, the high FSTI 

does promote transition, but it is not a pure bypass induced only by the 

freestream. Similar effects can be seen at other Reynolds numbers, as 

documented in Volino [11]. 

 Volino [11] correlated the reattachment location as a function of bar 

height as 

(st-sp)m/(st-sp)b = (1+0.23(d/θp)

1.56)-1 (1) 

where d is the bar height, θp is the baseline flow momentum thickness 

at the suction peak (which corresponds to the bar location), the 

subscripts p and t indicate suction peak and transition start, and the 

subscripts b and m denote baseline and modified cases. Equation (1) is 

based on a single study. More experiments with similar bars under 

different flow conditions are needed to test its general applicability. 

Turbulence Spectra 

 Although the small disturbances which eventually lead to transition 

are difficult to detect in the velocity profiles of Figs. 4 and 6, they can 

be seen in the turbulence spectra. The u′ , v′ and − vu ′′ spectra have 

been inspected, and disturbances are most consistently clear in the v′

spectra. Figure 7 shows the v′ spectra for baseline low FSTI cases at 

0 0.2 0.4 0.6 0.8 1 1.2 0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

baseline

s/Ls

Cp

0.4 mm bar

0.8 mm bar

1.6 mm bar

inviscid solution

High FSTI

Re=50,000

Fig. 5 Cp profiles, high FSTI, Re=50,000 cases 

0 0 0 0 0 1 0

0.05

u/Ue

y/Ls

a) 7 8 9 10 11

0 0 0 0 0 0

0.05

rms u′/Ue

y/Ls

b) 7 8 9 10 11

0.2

0 0 0 0 0 0

0.05

rms v′/Ue

y/Ls

c) 7 8 9 10 11

0.2

0 0 0 0 0 0

0.05

−u′v

′

/Ue

2

y/Ls

d) 7 8 9 10 11

0.01

0 0 0 0 0 1 0

0.05

γ

y/Ls

e) 7 8 9 10 11

baseline 0.4 mm bar 0.8 mm bar 1.6 mm bar

Fig. 6 Station 7-11 profiles, high FSTI, Re=50,000 cases: 

(a) mean velocity, (b) rms u′/Ue, (c) rms v′/Ue, (d) -u′v′/Ue

2

, 

(e) intermittency 

Re of 50,000 and 200,000. Data at each station are from the y location 

corresponding to the highest spectral peak. At Re=50,000, there are no 

clear peaks in the spectra at Stations 7-9, indicating that any important 

disturbances which may be present are as small as the very low 

background unsteadiness in the flow. By Station 10, a clear peak 

emerges, centered at about 80 Hz, along with some hints of a higher 

harmonic. By Station 11 this peak has grown by over 2 orders of 

magnitude. Note that although the peak grows by 4 orders of 

magnitude between Stations 9 and 11, the absolute value is still very 

low. The rms v′ profile of Fig. 4 is still near zero for this case at 

Station 11, and the shear layer remains laminar and separated. Hence, 

the spectra provide a means of tracking small pre-transitional 

NASA/CR—2012-217656 128
6 Copyright © 2004 by ASME 

100 101 102 103 10−12

10−10

10−8

10−6

10−4

10−2

PSD(v′

2) [m2/s2/Hz]

f [Hz]

Re=50,000, Low FSTI, Baseline

(a) Station 7

8

9

10

11

100 101 102 103

10−10

10−8

10−6

10−4

10−2

PSD(v′

2) [m2/s2/Hz]

f [Hz]

Re=200,000, Low FSTI, Baseline

(b)

Station 7

8

9

10

11

Fig. 7 Low FSTI, baseline case, boundary layer spectra: 

(a) Re=50,000, (b) Re=200,000

disturbances before they become apparent in the turbulence statistics. 

Figure 7b shows similar behavior at Re=200,000. A peak centered at 

about 600 Hz emerges at Station 9, more than three orders of 

magnitude larger than the background unsteadiness. It is still too 

small, however, to have a significant effect on the rms v′ profile. A 

significant rise in the v′ profile does not occur until Station 10, after 

transition has occurred. Transition is clearly visible in the spectra as a 

rise in energy of several orders of magnitude at all frequencies. This is 

a reminder that turbulence is broad banded, with a full range of scales, 

in contrast to the more narrow band disturbances which lead to 

transition. Similar results are found at all Reynolds numbers. The 

frequency of the peak in each cases matches the most unstable 

frequency for TS waves just before separation, as explained in Volino 

[9] using an analysis from Walker [17]. 

 Figure 8 shows the effect of the bars on the spectra at Station 8 of 

the Re=50,000 cases. The bars result in peaks at the same frequency 

as in the baseline case. The 0.4 mm and 0.8 mm thick bars cause a rise 

in the peaks of two and four orders of magnitude respectively above 

the baseline value. The absolute values of these peaks are still quite 

100 101 102 103 10−12

10−10

10−8

10−6

10−4

10−2

baseline

0.4 mm bar

0.8 mm bar

1.6 mm bar

PSD(v′

2) [m2/s2/Hz]

f [Hz]

Re=50,000, Low FSTI, Station 8

Fig. 8 Low FSTI, Station 8, boundary layer spectra 

small, however, in comparison to the turbulent values present in the 

1.6 mm bar case. 

Growth of Spectral Peaks 

 To quantify the growth of disturbances such as those shown in Figs. 

7 and 8, the data from the y location corresponding to the largest 

spectral peak at each station were selected, and the energy of the 

spectral peak was computed by integrating each spectrum over 

frequencies from 0.8 to 1.2 times the peak frequency of interest. So, 

for example, at Re=100,000, the peak frequency was centered at about 

273 Hz, so the spectra were integrated from 220 Hz to 327 Hz. These 

energies were then normalized using the exit velocity from the 

passage, Ue, and plotted verses the dimensionless distance from the 

suction peak (s-sp)/Ls. The distance from the suction peak was chosen 

because it is at the suction peak that the boundary layer becomes 

unstable and disturbances should start to grow. Figure 9 shows the 

resulting plot in log-linear coordinates for all of the low FSTI cases. 

At first glance, Fig. 9 may appear to be just a random jumble of data 

points, but on closer inspection, some patterns emerge. In each 

experimental case, the data have an upward slope, showing the growth 

of energy in the streamwise direction. The slopes appear to be greater 

at the higher Reynolds numbers. At any given Re, the cases with 

thicker bars have higher magnitude data, but the slopes appear to be 

about the same. The data from individual cases do lie along perfectly 

straight lines, and the trends noted above might not be noticed in an 

examination of one case by itself. When all the data are examined 

together, as in Fig. 9, the trends emerge. It should also be noted that 

the goal is to extract enough quantitative information from the data to 

develop a correlation for transition prediction. The exact steps in the 

data processing (e.g. the frequency range chosen for integration of the 

spectra) and the scatter of the data in intermediate steps, as apparent in 

Fig. 9, are not critical. 

 Figure 10 shows the data from Fig. 9 for the low FSTI baseline 

cases. There appears to be a background disturbance or “noise” level 

with a dimensionless magnitude of about 10-8.5. When the 

disturbances of interest become large enough, they rise above the 

noise. Comparison to the mean velocity profile data (e.g. Fig. 4) 

indicates that when the disturbances reach a magnitude of about 10-3.83, 

transition begins. The lines in Fig. 10 are based on fits to all 40 data 

sets. The slopes are assumed to depend on the Reynolds number as 

NASA/CR—2012-217656 129
7 Copyright © 2004 by ASME 

0 0.1 0.2 0.3 0.4 0.5 0.6 10−9

10−8

10−7

10−6

10−5

10−4

10−3

10−2

10−1

100

(s−sp)/Ls

Energy in v′ spectral peak / U2

e

Re=25,000

Re=50,000

Re=100,000

Re=200,000

Re=300,000

baseline

0.4mm bar

0.8mm bar

1.6mm bar

baseline

0.4mm bar

0.8mm bar

1.6mm bar

baseline

0.4mm bar

0.8mm bar

1.6mm bar

baseline

0.4mm bar

baseline

0.4mm bar

Low FSTI

Fig. 9 Normalized v′ spectral peak energy vs streamwise 

location for all low FSTI cases 

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45

10−12

10−10

10−8

10−6

10−4

10−2

Energy in v′ spectral peak / U2

e

(s−sp)/Ls

Transition Starts

Noise Level

Re=25,000

50,000

100,000

200,000

300,000

Low FSTI, Baseline

Symbols=data, Lines=correlation

Fig. 10 Normalized v′ spectral peak energy vs streamwise 

location for baseline low FSTI cases 

log10[(energy in v′ spectral peak)/Ue

2

]/[(s-sp)/Ls] = 0.504Re1/3 (2) 

When the data for the baseline cases are extrapolated back to (s-sp)=0, 

they all appear to begin at a common origin with magnitude 10-12.2. 

This magnitude is presumed to be proportional to the size of the initial 

disturbances which begin to grow at the suction peak. Since all of the 

experiments were done in the same facility, it is reasonable to believe 

that the dimensionless initial disturbance may be the same in all cases. 

 Figure 11 shows the data from Fig. 9 for the low FSTI, Re=50,000 

cases. The correlation lines all have the same slope, as given by Eq. 

(2). The intercepts at (s-sp)=0 depend on the bar height as 

Low FSTI intercept = [ ] ( ) 833 10 .]d/0.23[112.2-3.83 1.56 + θ p − (3) 

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45

10−12

10−10

10−8

10−6

10−4

10−2

Transition Starts

Noise Level

Energy in v′ spectral peak / U2

e

(s−sp)/Ls

baseline data

0.4 mm bar

0.8 mm bar

1.6 mm bar

Symbols=data, Lines=correlation

Low FSTI, Re=50,000

Fig. 11 Normalized v′ spectral peak energy vs streamwise 

location for low FSTI, Re=50,000 cases 

Equation (3) utilizes the baseline origin and transition magnitude of 

Fig. 10, and the form of Eq. (1), which relates the transition location in 

cases with bars to the location in the corresponding baseline case.

 Figure 12 shows the data for the high FSTI baseline cases in the 

format of Fig. 10. The background disturbance or “noise” level is 

higher than in the low FSTI cases. Since the range through which the 

disturbances grow between the noise and transition is smaller than in 

the low FSTI cases, it is doubtful that trends concerning the growth 

rate could have been extracted from the high FSTI data alone. When 

the slopes of Eq. (2) are applied to the high FSTI data, however, the fit 

is not implausible; suggesting that the growth rate of the disturbances 

depends mainly on the Reynolds number and not the FSTI. As in the 

low FSTI cases, the correlation lines for the baseline cases appear to 

extrapolate to a common origin, suggesting that the dimensionless 

initial disturbance caused by the high FSTI is the same for all cases. 

The high FSTI raises this initial disturbance level from 10-12.2 in the 

low FSTI cases to 10-8.9 for the high FSTI cases. Figure 13 shows data 

for the high FSTI Re=25,000 cases. The slope of the correlation lines 

is given by Eq. (2). The intercepts at (s-sp)=0 are based on the 

baseline value of Fig. 12 and the form of Eq. (3) as 

High FSTI intercept = [ ] ( ) 833 10 .]d/0.23[1-8.93.83 1.56 + θ p − (4) 

 The ratio of the intercept values for the high and low FSTI cases is 

10-8.9/10-12.2=1995. The intercept value is related to the size of the 

initial disturbance in the boundary layer, which is believed to be 

strongly influenced by the freestream turbulence, particularly the wall 

normal component. The v′ component of the FSTI at the suction peak 

is estimated to be 0.1% and 4.5% for the low and high FSTI cases 

respectively, based on the inlet FSTI and the acceleration of the flow 

between the inlet and the suction peak. The ratio of the v′ component 

of the freestream turbulence energy between the high and low FSTI 

cases, (4.5%/0.1%)2

=2025, is roughly equivalent to the ratio of the 

intercepts. The rough agreement may be fortuitous, but suggests the 

following formula for the intercept as a function of bar thickness, v′

component FSTI, and boundary layer thickness at the suction peak. 

Intercept = ( ) 833 10 .]d/0.23[1)](FSTIlog212.2-[3.83 1.56

p

2 ++ 10 + θ − (5) 

NASA/CR—2012-217656 130
8 Copyright © 2004 by ASME 

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 10−10

10−9

10−8

10−7

10−6

10−5

10−4

10−3

10−2

10−1

Energy in v′ spectral peak / U2

e

(s−sp)/Ls

Transition Starts

Noise Level

Re=25,000

50,000

100,000

200,000

300,000

High FSTI, Baseline

Symbols=data, Lines=correlation

Fig. 12 Normalized v′ spectral peak energy vs streamwise 

location for baseline high FSTI cases 

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 10−10

10−9

10−8

10−7

10−6

10−5

10−4

10−3

10−2

10−1

Transition Starts

Noise Level

Energy in v′ spectral peak / U2

e

(s−sp)/Ls

baseline data

0.4 mm bar

0.8 mm bar

1.6 mm bar

Symbols=data, Lines=correlation

High FSTI, Re=25,000

Fig. 13 Normalized v′ spectral peak energy vs streamwise 

location for high FSTI, Re=25,000 cases 

Simplifying Eq. (5) and combining with Eq. (2), the magnitude of the 

energy at any location is given as 

Magnitude = 

( ) 31 Lss0420833 sp 10

1.56 /

p

2 + 10 + θ ⋅−+− Re/)(..]d/0.23[1)](FSTIlog[-6.37

(6) 

Transition Correlations 

 To find the transition start location, the magnitude in Eq. (6) is set to 

10-3.83, and solved for the transition location as 

31

561

p

2

10

spt d2301

376981 FSTI Lss /

. Re

)/(.

)(log.(. /)( −

+

− =− θ

 (7) 

Utilizing Up/Ue=1.277 from the inviscid Cp profile, Eq. (7) can be 

expressed in terms of a Reynolds number as 

32

561

p

2

10

pt d2301

376532 FSTI /

. Re

)/(.

)(log.(. Re

+ θ

− = (8) 

where Rept is a Reynolds number based on the local freestream 

velocity at the suction peak and the distance from the suction peak to 

the start of transition. To potentially make Eq. (8) more generally 

applicable, it can be expressed in terms of the local values at the 

suction peak. For the present cases, 

Reθp = 0.393 Re1/2 (9) 

This results in 

34

561

p

2

10

pt p d2301

376808 FSTI /

. Re

)/(.

)(log.(. Re θ + θ

− = (10) 

 Although Eqs. (7), (8), and (10) were developed based on the 

turbulence spectra and a particular series of choices in processing the 

spectral data, the final correlations are only dependent on Reynolds 

number, FSTI and dimensionless bar height. It should, therefore, be 

possible to predict transition location without knowledge of the 

turbulence spectra. 

Comparison of Correlations to Experimental Data 

 Figure 14 compares the transition locations of the present 

experimental cases to Eq. (7). The error bars on the data indicate the 

uncertainty in the transition start location resulting from the finite 

spacing of the measurement stations. The agreement between the 

correlations and the data is not surprising, given that the correlations 

are based on these experimental data. Still, given the scatter apparent 

in Figs. 9-13, the good agreement in nearly all cases is reassuring. 

 Data from the baseline cases is compared to the present correlation 

and other correlations from the literature in Fig. 15. Mayle [2] 

presents the following correlations for short and long separation 

bubbles. 

Rest = 300 Reθs

0.7 short bubble (11)

Rest=1000 Reθs

0.7 long bubble (12) 

where Rest is a Reynolds number based on the freestream velocity at 

separation and the distance from separation to transition start. The 

correlation of Davis et al. [13] is 

Rest = 25000 log10[coth(0.1732 FSTI)] (13) 

The FSTI levels used with Eq. (13) for Fig. 15 are suction peak values 

based on both u′ and v′ (0.3% and 3.74% for the low and high FSTI 

cases). Hatman and Wang [7] identify several transition modes and 

present correlations for each of them. Their laminar separation mode 

transition correlation can be cast in terms of Rest as 

Rest = 0.0816 Res + 26805 (14) 

where Res=Usss/ν. Separation occurred at approximately the same 

location in all of the baseline cases of the present study, and 

(ss-sp)/Ls = 0.0833 (15) 

Us/Up = 0.971 (16) 

Reθs = 0.481 Re1/2 (17) 

NASA/CR—2012-217656 131
9 Copyright © 2004 by ASME 

0 0.5 1 1.5 2 2.5 3 3.5

x 105

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

(a)

Re

(s −s

t p)/Ls

baseline

0.4 mm bar

0.8 mm bar

1.6 mm bar

symbols=data

lines=correlation

Low FSTI

0 0.5 1 1.5 2 2.5 3 3.5

x 105

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

(b)

Re

(s −s

t p)/Ls

baseline

0.4 mm bar

0.8 mm bar

1.6 mm bar

symbols=data

lines=correlation

High FSTI

Fig. 14 Comparison of suction peak to transition start 

distance for experimental data and correlation: 

(a) low FSTI, (b) high FSTI

Using Eqs. (15)-(17), Eqs. (11)-(14) are cast in the coordinates of Eq. 

(8) and shown in Fig. 15. The present correlations agree with the data 

somewhat better than the other correlations, but all give reasonable 

predictions. The Hatman and Wang [7] correlation is based on low 

FSTI data, so it is not surprising that it only agrees with the low FSTI 

data of the present study. The Mayle [2] correlations do not take the 

FSTI into account explicitly. The long bubble correlation is intended 

for separation bubbles which do not close and therefore alter the 

pressure distribution along most of the airfoil surface. The short 

bubble correlation is intended for cases in which the boundary layer 

reattaches and the pressure distribution is only significantly altered in 

the vicinity of the bubble. To the extent that separation bubbles tend 

to be longer under low FSTI conditions, the Mayle [2] correlations can 

capture FSTI effects indirectly in some cases. The respective 

agreement of the long and short bubble correlations with the low and 

high FSTI data of the present study, however, may be somewhat 

fortuitous. The Davis et al. [13] correlation includes the effect of 

FSTI, but does not directly include the effect of Reynolds number. 

The present correlation appears to be the only one to explicitly include 

both effects. 

0 0.5 1 1.5 2 2.5 3 3.5

x 105

0

1

2

3

4

5

6

7

8

9

10 x 104

Low FSTI data

High FSTI data

Present low FSTI correlation

Present high FSTI correlation

Mayle long bubble

Mayle short bubble

Davis et al. low FSTI

Davis et al. high FSTI

Hatman & Wang

Rept

Re

Fig. 15 Comparison of suction peak to transition start Re 

for baseline data and various correlations 

Comparison to Data from Literature 

 The present correlation is based on the data of the present study. To 

better test it, data are used from other studies in the literature. The 

cases chosen for comparison all included flow over flat plates subject 

to a favorable pressure gradient followed by an adverse pressure 

gradient, with separation occurring shortly downstream of the suction 

peak. Cases which did not include these features did not have a single 

streamwise position which could be identified as the location where 

small disturbances would begin to grow, or they included elements of 

attached flow transition which the present correlation was not intended 

to predict. This limiting to a particular class of flows is not unduly 

restrictive on the utility of the present correlation, as it includes the 

flows of most interest for LPT boundary layers. Cases were selected 

in which FSTI, Reθs, Reθp, Rest, and Rept were either presented or could 

be extracted from the available data. The studies include those of Lou 

and Hourmouziadis [8], who include data for four steady flow cases; 

18 cases from Yaras [18]; three cases designated TL10, C and D from 

Howell [19]; and six cases from Volino and Hultgren [5], who 

considered a flat plate flow subject to the same pressure gradient as the 

cases of the present study. The results from these studies and 

comparisons to correlations are presented in Figs. 16-19. Equations 

(10)-(14) are used for the predictions. 

 The FSTI levels used in the comparisons were measured or 

extrapolated to values at the suction peak. Extrapolations were done, 

when necessary, using measured upstream FSTI values and the 

measured U∞ distributions. The present correlation was developed 

using FSTI levels based on v′ , but in many studies, only the u′

component is measured. As shown in Table 2, the u′ and v′

components can be different. In the absence of a turbulence 

generating grid, the v′ component tends to be lower than the u′

unsteadiness. For the cases of the present study and of Volino and 

Hultgren [5], results are presented based on an overall FSTI which 

includes the measured u′ and v′ components, and based on an FSTI 

which includes only the v′ component. For the Yaras [18], Lou and 

Hourmouziadis [8], and Howell [19] cases, only u′ FSTI are 

available. Results are presented based on these u′ FSTI, and also 

using assumed levels of v′ FSTI. For the low FSTI cases of Refs. [8], 

NASA/CR—2012-217656 132
10 Copyright © 2004 by ASME 

0 50 100 150 200 250 300 350 400 450 500 0

1000

2000

3000

4000

5000

6000

Present Study, Low FSTI

Present Study, High FSTI

Yaras [17], Low FSTI

Yaras [17], High FSTI

Lou & Hourmouziadis [6], Low FSTI

Howell [18], Low FSTI

Volino & Hultgren [8], Low FSTI

Volino & Hultgren [8], High FSTI

Present Correlation

First symbol : v′ FSTI

Second symbol : u′ or overall FSTI

Reθp

Rept/[8.801*(6.37−log10(FSTI2))]

Fig. 16 Comparison of experimental data from various 

studies to present correlation 

0 100 200 300 400 500 600 700 800 0

2

4

6

8

10

12 x 104

Present Study, Low FSTI

Present Study, High FSTI

Yaras [17], Low FSTI

Yaras [17], High FSTI

Lou & Hourmouziadis [6], Low FSTI

Howell [18], Low FSTI

Volino & Hultgren [8], Low FSTI

Volino & Hultgren [8], High FSTI

Mayle [2], Short Bubble

Mayle [2], Long Bubble

Reθs

Rest Fig. 17 Comparison of experimental data from various 

studies to correlations from Mayle [2] 

[18], and [19], v′ FSTI of 0.1% was assumed. This assumption 

presumes similar behavior to the present study, in which streamwise 

unsteadiness results in rms u′ of about 0.5% in the freestream, while 

v′ is significantly lower at 0.1%. For the higher FSTI cases of Yaras 

[18], the freestream turbulence was assumed isotropic, so the u′ and 

v′ FSTI were assumed equal. It should be noted that results based on 

assumed v′ levels for Refs. [8], [18], and [19] are presented only to 

demonstrate the potential differences that may exist between u′ and 

v′ based results, and to show the possibility of better prediction if 

results are based on v′ FSTI. Results based on assumed v′ FSTI are 

not used for assessment of correlations. Evaluation of the accuracy of 

correlations should be based on comparison to the u′ FSTI results for 

Refs. [8], [18], and [19], since the u′ FSTI were actually measured. 

 Figure 16 compares the data from all cases to the present 

correlation, Eq. (10). The open symbols in Figs. 16-19 represent cases 

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 0

1

2

3

4

5

6

7

8

9 x 104

Present Study, Low FSTI

Present Study, High FSTI

Yaras [17], Low FSTI

Yaras [17], High FSTI

Lou & Hourmouziadis [6], Low FSTI

Howell [18], Low FSTI

Volino & Hultgren [8], Low FSTI

Volino & Hultgren [8], High FSTI

Davis et al. [12] Correlation

First symbol : v′ FSTI

Second symbol : u′ or overall FSTI

FSTI [%]

Rest Fig. 18 Comparison of experimental data from various 

studies to correlation from Davis et al. [13] 

0 1 2 3 4 5 6 7 8 9 10

x 105

0

2

4

6

8

10

12 x 104

Present Study, Low FSTI

Present Study, High FSTI

Yaras [17], Low FSTI

Yaras [17], High FSTI

Lou & Hourmouziadis [6], Low FSTI

Howell [18], Low FSTI

Volino & Hultgren [8], Low FSTI

Volino & Hultgren [8], High FSTI

Hatman and Wang [5] Correlation

Res

Rest Fig. 19 Comparison of experimental data from various 

studies to correlation from Hatman and Wang [7] 

with low FSTI (no turbulence grid), and the solid symbols are for 

higher FSTI cases. Agreement between the data and present 

correlation is good in all cases, and extends to Reynolds numbers over 

twice those of the present study. The agreement is better when the 

FSTI is based on v′ . In terms of predicting the transition start 

location, the agreement between the correlation and data is generally 

within the experimental uncertainty resulting from the finite 

streamwise measurement station spacing. 

 Figure 17 compares the data from all cases to the long and short 

bubble correlations of Mayle [2] (Eq. 11-12). The trend of the data 

and correlations with Reynolds number agree, and the two correlation 

lines tend to bracket the data. If the long and short bubble correlations 

are used for the low and high FSTI cases, respectively, they are able to 

predict the transition start location about as accurately as the present 

NASA/CR—2012-217656 133
11 Copyright © 2004 by ASME 

correlation in many cases. There are, however, cases in which the data 

fall midway between the long and short bubble correlations. 

 Figure 18 compares the data to the Davis et al. [13] correlation (Eq. 

13). The correlation under predicts the distance from separation to 

transition in most cases. It tends to do best at the higher Reynolds 

numbers, and misses the Reynolds number dependence shown in Figs. 

16 and 17. 

 Figure 19 compares the data to the Hatman and Wang [7] 

correlation (Eq. 14). Agreement between the data and correlation is 

good in the low FSTI cases, and prediction of the transition start 

location is as good as with the present correlation. With higher FSTI, 

however, the Hatman and Wang [7] correlation over predicts the 

distance from separation to transition. 

CONCLUSIONS 

1. A correlation for separated flow transition has been developed 

based on the measured growth of small disturbances in the separated 

shear layer. The correlation has been demonstrated effective over a 

range of Reynolds numbers and freestream turbulence levels through 

comparison to data from several studies from the literature. 

2. Disturbances appear to begin growing at the beginning of the 

adverse pressure gradient region, where the boundary layer becomes 

unstable. The size of the initial disturbance at this location depends on 

the freestream turbulence level and the height of any obstruction 

placed on the surface. In dimensionless form, the size of the 

disturbance does not appear to depend on the Reynolds number. 

3. The growth rate of the disturbances depends on the Reynolds 

number, and appears independent of the FSTI or size of the initial 

disturbance. 

4. The transition location depends both on Reynolds number and 

FSTI. The present correlation represents a physics based improvement 

over previous correlations that did not account for both of these 

effects. 

5. The wall normal component of the freestream turbulence is 

preferred for the correlation, and results in better predictions than the 

u′ based FSTI. Measurement of freestream v′ should be included in 

future experimental studies. 

ACKNOWLEDGEMENTS 

 This work was sponsored by the NASA Glenn Research Center. 

The grant monitor is Dr. David Ashpis. Additional matching support 

was provided through a U.S. Naval Academy Recognition Grant. The 

second author received matching support from the Office of Naval 

Research as a postdoctoral fellow. Tabulated experimental results for 

Ref. [18] were provided by Stephen Roberts of Carleton University. 

 

REFERENCES 

[1] Hourmouziadis, J., 1989, “Aerodynamic Design of Low Pressure 

Turbines,” AGARD Lecture Series 167. 

[2] Mayle, R.E., 1991, “The Role of Laminar-Turbulent Transition in 

Gas Turbine Engines,” ASME Journal of Turbomachinery, 113, 

pp. 509-537. 

[3] Sharma, O.P., Ni, R.H., and Tanrikut, S., 1994, “Unsteady Flow 

in Turbines,” AGARD Lecture Series 195, Paper No. 5. 

[4] Bons, J.P., Sondergaard, R., and Rivir, R.B., 2001, 

“Turbine Separation Control Using Pulsed Vortex 

Generator Jets,” ASME Journal of Turbomachinery, 123, pp. 

198-206.

[5] Volino, R.J., and Hultgren, L.S., 2001, “Measurements in 

Separated and Transitional Boundary Layers Under Low-Pressure 

 Turbine Airfoil Conditions,” ASME Journal of Turbomachinery, 

123, pp. 189-197. 

[6] Volino, R.J., 2002, “Separated Flow Transition Under Simulated 

Low-Pressure Turbine Airfoil Conditions: Part 1 – Mean Flow 

and Turbulence Statistics,” ASME Journal of Turbomachinery, 

124, pp. 645-655. 

[7] Hatman, A., and Wang, T., 1999, “A Prediction Model for 

Separated Flow Transition,” ASME Journal of Turbomachinery, 

121, pp. 594-602. 

[8] Lou, W., and Hourmouziadis, J., 2000, “Separation Bubbles 

Under Steady and Periodic-Unsteady Main Flow Conditions,” 

ASME Journal of Turbomachinery, 122, pp. 634-643. 

[9] Volino, R.J., 2002, “Separated Flow Transition under Simulated 

Low-Pressure Turbine Airfoil Conditions: Part 2 - Turbulence 

Spectra,” ASME Journal of Turbomachinery, 124, pp. 656-664. 

[10] Hughes, J.D. and Walker, G.J., 2001, “Natural Transition 

Phenomena on an Axial Compressor Blade,” ASME Journal of 

Turbomachinery, 123, pp. 392-401. 

[11] Volino, R.J., 2003, “Passive Flow Control on Low-Pressure 

Turbine Airfoils,” ASME Journal of Turbomachinery, 125, pp. 

754-764. 

[12] Roberts, W.B., 1980, “Calculation of Laminar Separation 

Bubbles and Their Effect on Airfoil Performance,” AIAA Journal, 

18, pp. 25-31. 

[13] Davis, R.L., Carter, J.E., and Reshotko, E., 1985, “Analysis of 

Transitional Separation Bubbles on Infinite Swept Wings,” AIAA 

Paper 85-1685. 

[14] Volino, R.J., Schultz, M.P., and Pratt, C.M., 2001, “Conditional 

Sampling in a Transitional Boundary Layer Under High FreeStream Turbulence Conditions,” ASME Journal of Fluids 

Engineering, 125, pp. 28-37. 

[15] Ligrani, P.M., and Bradshaw, P., 1987, “Spatial Resolution and 

Measurement of Turbulence in the Viscous Sublayer Using 

Subminiature Hot-Wire Probes,” Experiments in Fluids, 5, pp. 

407-417. 

[16] Ligrani, P.M., Westphal, R.V., and Lemos, F.R., 1989, 

“Fabrication and Testing of Subminiature Multi-Sensor Hot-Wire 

Probes,” Journal of Physics E: Scientific Instruments, 22, pp. 

262-268. 

[17] Walker, G.J., 1989, “Transitional Flow on Axial Turbomachine 

Blading,” AIAA Journal, 27, pp. 595-602. 

[18] Yaras, M.I., 2002, “Measurement of the Effects of Freestream 

Turbulence on Separation-Bubble Transition,” ASME Paper GT2002-30232. 

[19] Howell, R.J., 1999, “Wake – Separation Bubble Interactions in 

Low Reynolds Number Turbomachinery,” Ph.D. Thesis, Whittle 

Laboratory, Cambridge University. 

NASA/CR—2012-217656 134
1 Copyright © 2005 by ASME 

EXPERIMENTS WITH THREE DIMENSIONAL PASSIVE FLOW CONTROL DEVICES ON 

LOW-PRESSURE TURBINE AIRFOILS

Douglas G. Bohl and Ralph J. Volino

Department of Mechanical Engineering 

United States Naval Academy 

Annapolis, Maryland 21402 

Email: volino@usna.edu

ABSTRACT 

 The effectiveness of three dimensional passive devices for flow 

control on low pressure turbine airfoils was investigated 

experimentally. A row of small cylinders was placed at the pressure 

minimum on the suction side of a typical airfoil. Cases with Reynolds 

numbers ranging from 25,000 to 300,000 (based on suction surface 

length and exit velocity) were considered under low freestream 

turbulence conditions. Streamwise pressure profiles and velocity 

profiles near the trailing edge were documented. Without flow control 

a separation bubble was present, and at the lower Reynolds numbers 

the bubble did not close. Cylinders with two different heights and a 

wide range of spanwise spacings were considered. Reattachment 

moved upstream as the cylinder height was increased or the spacing 

was decreased. If the spanwise spacing was sufficiently small, the 

flow at the trailing edge was essentially uniform across the span. The 

cylinder size and spacing could be optimized to minimize losses at a 

given Reynolds number, but cylinders optimized for low Reynolds 

number conditions caused increased losses at high Reynolds numbers. 

The effectiveness of two-dimensional bars had been studied previously 

under the same flow conditions. The cylinders were not as effective 

for maintaining low losses over a range of Reynolds numbers as the 

bars. 

NOMENCLATURE 

Cp ( ) 2

T ρUpp2 e − / , pressure coefficient 

D cylinder diameter 

H shape factor, δ

*

/θ

Ls suction surface length 

P center to center spacing of cylinders 

p pressure 

pT upstream stagnation pressure

Re LU se

/ν , exit Reynolds number 

s streamwise coordinate, distance from leading edge 

Ue nominal exit freestream velocity, based on inviscid solution 

u mean streamwise velocity 

u rms streamwise fluctuating velocity 

y cross-stream coordinate, distance from wall 

z spanwise coordinate 

δ

∗

 displacement thickness 

ν kinematic viscosity 

ρ density 

θ momentum thickness 

θb

 momentum thickness in baseline case 

INTRODUCTION 

 Boundary layer separation is a known problem on some modern 

low-pressure turbine (LPT) airfoils, due to the strong adverse pressure 

gradients created when designers impose higher loading in an effort to 

improve efficiency and lower cost by reducing airfoil count in engines. 

Separation bubbles, particularly those which fail to close, can result in 

a significant loss of lift and a subsequent degradation of engine 

efficiency (e.g. Hourmouziadis [1], Mayle [2], and Sharma et al. [3]). 

The problem is particularly relevant in aircraft engines. Airfoils 

optimized to produce maximum power under takeoff conditions may 

still experience boundary layer separation at cruise conditions, due to 

the lower density and therefore lower Reynolds numbers at altitude. A 

component efficiency drop of 2% may occur between takeoff and 

cruise conditions in large commercial transport engines, and the 

difference could be as large as 7% in smaller engines operating at 

higher altitudes [4, 5]. Prediction and control of suction side 

separation, without sacrifice of the benefits of higher loading, is 

therefore, crucial for improved engine design. 

 Separation on airfoils is complicated by boundary layer transition. 

Separated flow transition in the LPT has been the focus of several 

recent studies. Volino [6] provides a review and describes as follows 

the transition process on the suction side of a typical LPT airfoil. The 

strong acceleration on the leading section of the airfoil keeps the 

boundary layer thin and laminar, even in the presence of elevated 

freestream turbulence. In most cases Volino [6] observed that the 

boundary layer separated just downstream of the suction peak. If 

transition then occurred in the shear layer over the separation bubble, 

it caused the boundary layer to reattach. 

 A few recent studies have focused on control of transition and 

reattachment in the LPT. Some have used active devices. Huang et al. 

Proceedings of GT2005

ASME Turbo Expo 2005: Power for Land, Sea and Air

June 6-9, 2005, Reno-Tahoe, Nevada, USA

GT2005-68969

NASA/CR—2012-217656 135
2 Copyright © 2005 by ASME 

[7] and Hultgren and Ashpis [8] employed high voltage electrodes to 

produce glow discharge plasma in a boundary layer to control 

separation. Bons et al. [4] used steady and pulsed vortex generator 

jets. Volino [9] used oscillating vortex generator jets with no net mass 

flow. Sieverding et al. [10] used adjustable devices built into the 

suction surface. 

 While active flow control provides a means for adjusting to 

changing flow conditions and in some cases the benefits of calmed 

regions [4, 9], passive flow control holds the advantage of simplicity. 

Van Treuren et al. [11], Lake et al. [12]. Murawski and Vafai [13], 

Byerley et al. [14], Volino [15], Sieverding [10], Vera et al. [16], and 

Zhang and Hodson [17] used various passive devices under LPT 

conditions to control separation and in many cases reduce losses. 

Most employed a relatively simple modification, such as a small trip 

wire or bar (essentially roughness), on the suction surface of an airfoil. 

 Successful flow control results in a thin, attached boundary layer at 

the trailing edge of an airfoil, thereby reducing losses. The consensus 

of the studies listed above is that a device on the suction surface 

should be placed at or slightly downstream of the pressure minimum. 

This is a logical result, since the effects of a device farther upstream 

would be damped by the favorable pressure gradient, and a device too 

far downstream would lie under the separation bubble and be 

ineffective. Volino [15] used rectangular bars and found that the 

optimal bars were not large enough to immediately trip the boundary 

layer to turbulent, but instead allowed a small separation bubble to 

form. The bars introduced small disturbances that grew and caused 

transition and reattachment to move upstream of their location in the 

uncontrolled case, as explained in Volino and Bohl [18]. The optimal 

bar height depended on the flow conditions. As Re or freestream 

turbulence is lowered, the separation bubble becomes larger, so a 

larger bar is needed to produce enough of a disturbance to move 

transition sufficiently far upstream. A flow control device producing 

too small a disturbance will allow a larger separation bubble than 

desired, resulting in a thicker boundary layer downstream of 

reattachment and higher losses. Similarly, too large a disturbance will 

move transition farther upstream than necessary, resulting in a longer 

turbulent region and higher losses. Volino [15], Sieverding et al. [10], 

and Zhang and Hodson [17] all found that under steady flow 

conditions, devices optimized for low Re tend to increase losses at 

high Re. Devices optimized for high Re can be too small to be 

effective at low Re. Unsteady wakes from upstream airfoils promote 

transition and reattachment, and Zhang and Hodson [17] found that in 

unsteady flow optimal control was achieved using smaller devices than 

in comparable steady flow cases. This made it possible in unsteady 

flow to reduce losses with a single device over a wider range of Re. 

 Passive flow control devices of various geometries have been tested, 

but it is still uncertain if any particular device is superior. In 

preliminary testing, Volino [15] considered trip wires, rectangular 

bars, and delta wing vortex generators. All produced similar results. 

Sieverding et al. [10] found that straight trip wires were somewhat 

better than rows of spherical roughness elements, but only a limited 

number of cases were tested. Lake et al. [12] found dimples superior 

to other devices, presumably because the dimples produced less 

blockage than devices that protruded into the flow. Again, however, 

the number of cases considered was limited, and more recent evidence 

[15, 17] suggests that optimal devices should be quite small and 

produce minimal blockage even if they do extend into the flow. Zhang 

and Hodson [17] noted differences in transition location with straight 

and “wavy” trip wires and rectangular bars. The sharp backward 

facing step on a bar, for example, produced an earlier transition than a 

round trip wire of the same height. Still, this does not preclude that a 

bar and a slightly larger wire could produce comparable results. 

 It is still possible that some devices might prove better for reducing 

losses than others. Reynolds number can vary by an order of 

magnitude during engine operation given the change in ambient 

pressure between takeoff and cruise. Since passive devices by 

definition cannot be adjusted as conditions change, it is highly 

desirable to use devices that reduce losses over as large a Reynolds 

number range as possible. Given the potential payoff of a more 

efficient engine, it is worthwhile to further consider passive flow 

control devices of different geometries. 

 A row of small vertical cylinders is considered in the present study. 

The cylinders are located at the pressure minimum on the suction 

surface and the spacing between cylinders is varied. The geometry 

was chosen in the hope that in comparison to 2-dimensional bars or 

trips wires of the same height, isolated elements might produce a 

stronger disturbance due to the 3-dimensional nature of the flow 

around them, while presenting less blockage due to the gaps between 

elements. The net result would presumably be successful separation 

control with lower losses. Experimental conditions match the low 

freestream turbulence cases of Volino [15]. Details of the 

experimental conditions and the results are presented below. 

EXPERIMENTS

 Experiments were conducted in a low speed wind tunnel, described 

by Volino et al. [19]. Briefly, air enters through blowers and passes 

through a honeycomb, a series of screens, two settling chambers, and a 

three-dimensional contraction before entering the test section. At the 

exit of the contraction, the mean velocity is uniform to within 1%. 

The freestream turbulence intensity is 0.5%±0.05%. Nearly all of this 

freestream “turbulence” is actually streamwise unsteadiness at 

frequencies below 20 Hz and is not associated with turbulent eddies. 

The rms intensities of the three components of the unsteadiness are 

0.7%, 0.2% and 0.2% in the streamwise, pitchwise and spanwise 

directions, respectively. 

 The test section, shown in Fig. 1, follows the contraction and 

consists of the passage between two airfoils. Details are listed in 

Table 1 and more information is available in Volino [6]. A large span 

to chord ratio of 4.3 was chosen to insure two-dimensional flow at the 

spanwise centerline of the airfoils, where all measurements were 

made. Upstream of each airfoil are flaps, which control the amount of 

bleed air allowed to escape from the passage. The flaps, along with a 

tailboard on the pressure side of the passage, are adjusted to produce 

the correct leading edge flow and pressure gradient along the airfoils. 

The flow in the passage matches that in a multi-blade cascade. The 

geometry of the passage corresponds to that of the Pak-B airfoil, 

which is an industry supplied research airfoil that is representative of a 

modern LPT design. It has been used in several studies, as noted in 

Volino [6]. 

 Experimental conditions match the smooth airfoil baseline cases of 

Volino [6], who considered five Reynolds numbers (Re=25,000, 

50,000, 100,000, 200,000, and 300,000). Reynolds numbers are based 

on the exit velocity from the passage and the suction surface length, Ls. 

Comparison data from cases with passive-bar flow control are from 

Volino [15]. The bars were of uniform rectangular cross section and 

extended along the airfoil, as shown in Fig. 2. The trailing edge of the 

bar was located at s/Ls=0.51, near the suction surface velocity peak. 

All bars were 6 mm wide in the streamwise direction. Bar heights of 

0.4 mm, 0.8 mm and 1.6 mm were used. The bar heights were all less 

than 1% of Ls. They compare to local boundary layer thickness at the 

bar location of about 3.8 mm, 2.7 mm, 2.0 mm, 1.4 mm and 1.2 mm in 

the baseline Re=25,000 through 300,000 cases respectively. 

 A row of small vertical cylinders is used in the present experimental 

cases. The cylinders are D=6 mm in diameter with one end affixed to 

NASA/CR—2012-217656 136
3 Copyright © 2005 by ASME 

tailboard

flap

flap bleed air

bleed air

35o

60o

Fig. 1 Schematic of the test section 

the suction surface and the other extending into the flow. The 

cylinders are located in a line at the suction peak, in the same location 

as the bars described above, as shown in Fig. 2. Cylinder heights of 

0.4 mm and 1.6 mm were considered, matching the smaller and larger 

bar heights of Volino [15]. Center to center cylinder spacings ranging 

from a pitch, P, of 1D (i.e. the cylinders were touching) to 30D were 

considered. 

Measurements 

 Pressure surveys were made for each case using a pressure 

transducer (0-870 Pa range Validyne transducer) and a Scanivalve. 

Stagnation pressure was measured with a pitot tube upstream of the 

passage inlet, and eleven pressure taps were located on each airfoil 

along their spanwise centerlines. The uncertainty in Cp is 7% at 

Re=25,000, and 4% in other cases. Most of this uncertainty is due to 

bias error. Stochastic error was minimized by averaging pressure 

transducer readings over a 10 second period. The flow control 

cylinders were moved in the spanwise direction allowing 

documentation with the pressure taps directly downstream of the 

center of one cylinder and with the taps midway between the centers 

of two adjacent cylinders. 

 Profiles of the streamwise velocity component in the suction side 

boundary layer were measured for the Re=50,000 and 300,000 cases 

near the trailing edge at s/Ls=0.944. Profiles were measured at several 

spanwise locations relative to the position of the upstream cylinders. 

Data were acquired with a hot-wire anemometer (AA Lab Systems 

model AN-1003) and a single sensor boundary layer probe (TSI model 

1218-T1.5). The sensor diameter is 3.8 μm, and the active length is 

1.27 mm. At each measurement location, data were acquired for 26 

seconds at a 20 kHz sampling rate (219 samples). Data were acquired 

at 60 wall normal locations in each profile, extending from the wall to 

the free-stream, with most points concentrated in the near wall region. 

The closest point was within 0.1 mm of the wall, which corresponds to 

y/Ls=0.0004 and between 0.02 and 0.04 boundary layer thicknesses. 

Uncertainties in the mean velocity are 3-5% except in the very near 

wall region where near-wall corrections (Wills [20]) were applied to 

the mean velocity. Uncertainties in the momentum and displacement 

thicknesses computed from the mean profiles are 10%. Uncertainty in 

the shape factor, H, is 8%. The uncertainty in the fluctuating 

streamwise velocity is below 10%, except in the very near wall region 

(y<1 mm, y/Ls<0.004) of the Re=300,000 cases, where spatial 

averaging effects, due to the finite length of the hot-wire sensor, may 

become important. This is discussed in Volino [6]. 

(a) (b) 

Fig. 2 Scale drawing of suction side airfoil showing 

location of a) bar, or b) cylinders 

Table 1: Test section parameters 

Axial 

Chord 

[mm] 

True 

Chord 

[mm] 

Pitch 

[mm] 

Span 

[mm] 

Suction 

side, Ls

[mm] 

Inlet 

flow 

angle 

Exit 

flow 

angle 

inlet

e

U

U

153.6 170.4 136.0 660.4 228.6 35° 60° 1.64 

RESULTS 

Pressure Profiles

 Suction side pressure profiles for the cases with larger (1.6 mm 

high) cylinders are shown in Figs. 3-5. Also shown in each figure are 

the corresponding baseline results from Volino [6], the 1.6 mm thick 

bar case results from Volino [15], and an inviscid solution for flow 

through the passage. Figure 3 shows results at Re=25,000 with various 

cylinder spacings and the pressure taps directly downstream of one 

cylinder. The data agree with the inviscid solution in the favorable 

pressure gradient region. The near wall flow slows as it approaches a 

cylinder or bar, causing a drop in the measured Cp below the inviscid 

solution at the fifth pressure tap. If the boundary layer separates and 

does not reattach, there is a drop in the suction peak, as indicated by 

low Cp values. In addition to this effect, if the blockage caused by a 

bar or cylinder is sufficiently high, the streamlines immediately 

downstream will be displaced as they flow over the sixth pressure tap, 

which may cause the measured local Cp to either rise or fall relative to 

the inviscid solution, depending on the geometry of a particular case. 

For controlling separation, cylinders with spacing of P=2D or more are 

ineffective at Re=25,000. There is a plateau in Cp in the adverse 

pressure gradient region extending to the trailing edge, indicating a 

separation bubble that does not reattach. The P=1D (touching) 

cylinders and the solid bar result in a large separation bubble which 

appears to be starting to reattach by the trailing edge, as indicated by 

the drop in Cp back toward the inviscid value at the last pressure tap. 

Figure 4 show the results at Re=50,000 with the pressure taps directly 

downstream of one cylinder (z/P=0, Fig. 4a) and with the taps 

downstream of the midpoint between adjacent cylinders (z/P=0.5, Fig. 

4b). In the baseline case, the boundary layer does not reattach. With 

P=1D or the solid bar, reattachment moves upstream to s/Ls=0.78. 

With P=2D reattachment occurs at s/Ls=0.86. With P>2D the effect of 

the cylinders is not observed until the last pressure tap, but even with 

P=30D there is some effect on Cp at s/Ls=0.94 and z/P=0. As the 

spacing increases, the Cp values increase from the inviscid solution 

toward the baseline case value. At z/P=0.5 and s/Ls=0.94, Cp is 

affected for spacings up to 20D, although not as strongly as it is at 

z/P=0. For P=24D and 30D, the cylinders are too far apart to affect Cp

at z/P=0.5. The results at Re=100,000 are shown in Fig. 5. In the 

baseline case, there is a clear separation bubble, and it reattaches near 

NASA/CR—2012-217656 137
4 Copyright © 2005 by ASME 

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

0.8

1

1.2

1.4

1.6

1.8

2

s/L

s

Cp

z/P=0

Large Cylinders

Re=25,000

baseline

2D bar

P/D=1

2

3

5

10

16

24

30

inviscid

Fig. 3 Pressure profiles, large cylinders, Re=25,000, z/P=0 

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

0.8

1

1.2

1.4

1.6

1.8

2

s/L

s

Cp

(a)

z/P=0

Large Cylinders

Re=50,000

baseline

2D bar

P/D=1

2

3

5

10

16

20

24

30

inviscid

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

0.8

1

1.2

1.4

1.6

1.8

2

s/L

s

Cp

(b)

z/P=0.5

Large Cylinders

Re=50,000

baseline

2D bar

P/D=2

3

5

10

16

20

24

30

inviscid

Fig. 4 Pressure profiles, large cylinders, Re=50,000, a) 

z/P=0, b) z/P=0.5 

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

0.8

1

1.2

1.4

1.6

1.8

2

s/L

s

Cp

(a)

z/P=0

Large Cylinders

Re=100,000

baseline

2D bar

P/D=1

2

3

5

10

16

20

24

30

inviscid

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

0.8

1

1.2

1.4

1.6

1.8

2

s/L

s

Cp

(b)

z/P=0.5

Large Cylinders

Re=100,000

baseline

2D bar

P/D=2

3

5

10

16

20

24

30

inviscid

Fig. 5 Pressure profiles, large cylinders, Re=100,000, a) 

z/P=0, b) z/P=0.5 

the trailing edge. With P≤3D, the separation is essentially eliminated. 

With P=5D there is a separation, but reattachment by s/Ls=0.78. For 

P≤5D, the Cp values are nearly uniform across the span. For 

10D≤P≤30D, the boundary layer appears to be at least starting to 

reattach by s/Ls=0.86, with the effect on Cp greater at z/P=0. Results 

for the Re=200,000 cases (not shown) are similar to those described 

above. The boundary layer separates in the baseline case and 

reattaches by s/Ls=0.86. Cylinders with P≤16D effectively suppress 

the separation across the span. For P>16D, the boundary layer is 

attached at s/Ls=0.78 and z/P=0, but is still separated at z/P=0.5. At 

Re=300,000 there is only a small separation bubble in the baseline 

case, and it reattaches by s/Ls=0.78. Cylinders with spacing up to 30D

appear to suppress this small bubble across the span. 

 Comparing the results at z/P=0 and z/P=0.5, the spanwise influence 

of each cylinder can be estimated. If the region affected by the 

cylinder is assumed to spread linearly in the spanwise direction as the 

flow convects downstream, the half angle for the spreading is between 

roughly 30 and 40 degrees. Changes in Cp at z/P=0.5 could be due to 

local turbulence arising from the cylinder induced disturbances. It is 

NASA/CR—2012-217656 138
5 Copyright © 2005 by ASME 

also possible, however, that these changes in Cp are due to changes in 

the mean flow resulting from the cylinder induced changes at z/P=0. 

 The smaller (0.4 mm high) cylinders and bar are ineffective at 

Re=25,000 and 50,000. The boundary layer separates and does not 

reattach. At Re=100,000, as shown in Fig. 6, the cylinders cause the 

first indication of reattachment to move upstream from near the 

trailing edge to s/Ls=0.78 or 0.86. Cylinder spacings up to 10D were 

considered, and the Cp values decrease from the baseline case values 

toward the inviscid solution as P is decreased. No significant 

spanwise variation in Cp was observed. The results at z/P=0.5 are 

essentially the same as those shown in Fig. 6 for z/P=0. Results for 

the Re=200,000 cases are shown in Fig. 7. The results are very similar 

to those of the Re=100,000 cases, but with reattachment shifted about 

0.08Ls upstream. When Re=300,000 the baseline separation bubble is 

small, and cylinders effectively eliminate it for all value of P. 

 The presumably more complex 3-dimensional flow around isolated 

(P>1D) cylinders did not produce disturbances better able to control 

separation than the touching (P=1D) cylinders or two-dimensional 

bars. In fact, the opposite was observed. The pressure profiles show 

that as the spacing between cylinders increases, their ability to 

promote reattachment decreases. This is not simply due to the finite 

spanwise influence of each cylinder, as very little spanwise variation 

was observed for cases with P<10D. Whether increasing the spacing 

results in reduced flow blockage and losses will be considered next as 

the velocity profile results are considered. 

Velocity Profiles 

 Mean and fluctuating streamwise velocity profiles for the large 

cylinder Re=50,000 cases at z/P=0 are shown in Fig. 8. Data were 

acquired near the trailing edge (s/Ls=0.94). The baseline case shows a 

thick separation bubble in the mean profile and a small peak in u in 

the shear layer over the bubble. The 2-dimensional bar case exhibits a 

fully attached turbulent mean profile shape and a typical turbulent u

profile with a near wall peak. For the cases with cylinder spacing up 

to P=5D, the mean profiles appear attached and turbulent, but the trend 

is toward a less full profile as P increases. This trend continues for the 

cases with P≥20D, with the mean profiles in these cases appearing 

only partially reattached. The peaks in the u profiles increase in 

magnitude and move farther from the wall as the cylinder spacing is 

increased. High peaks away from the wall are typical of transitioning 

and reattaching boundary layers. Transition begins in the shear layer 

over the separation bubble, so u should be high at this location. The 

high u peaks result from the switching between intermittently 

attached-turbulent-like and separated-laminar-like states. Figure 9 

shows the profiles at z/P=0.5. For P≤5D, the profiles appear 

essentially the same as those at z/P=0. For P≥20D the mean profiles 

show a separation bubble, somewhat thinner than that of the baseline 

case, and u profiles with only a small peak in the shear layer over the 

separation bubble. Figure 10 shows the velocity profiles at several 

spanwise positions for the P=30D case. The progression from 

attached flow at z/P=0 to separated flow at z/P=0.5 is clear. The 

profiles at z/P=0.13 and 0.20 are on the edge of the attached flow 

region with u peaks between the high peaks at low z/P and the small 

shear layer peaks at higher z/P. The influence of the cylinders to a 

spanwise position between z/P=0.13 and 0.20 suggests a half angle for 

the spread of the cylinder influence of about 20 degrees. The velocity 

profile data of Figs. 8-10 are consistent with the conclusions drawn 

from the corresponding pressure coefficient data of Fig. 4. 

 The information concerning separation in the mean profiles can be 

presented in terms of the shape factor, H. The shape factor rises to 

about 4 when a laminar boundary layer separates, and reaches higher 

values as a separation bubble thickens. This is due to an increase in 

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

0.8

1

1.2

1.4

1.6

1.8

2

s/L

s

Cp

z/P=0

Small Cylinders

Re=100,000

baseline

2D bar

P/D=1

2

3

5

10

inviscid

Fig. 6 Pressure profiles, small cylinders, Re=100,000, z/P=0 

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

0.8

1

1.2

1.4

1.6

1.8

2

s/L

s

Cp

z/P=0.5

Small Cylinders

Re=200,000

baseline

2D bar

P/D=1

2

3

5

10

inviscid

Fig. 7 Pressure profiles, small cylinders, Re=200,000, z/P=0 

the displacement thickness while the momentum thickness remains 

nearly constant. If the boundary layer reattaches, the displacement 

thickness drops and the momentum thickness begins to rise. The 

shape factor reaches a turbulent value of about 1.4 after recovering 

from the separation. Figure 11 shows H as a function of z/P for the 

Re=50,000 cases with the larger cylinders. In the baseline case, 

H=6.5, indicating a thick separation bubble. With the twodimensional bar, H=1.6, indicating that the boundary layer has 

reattached and is nearing fully developed turbulent conditions. With 

cylinder spacing up to P=5D, H is spanwise uniform. With P=5D, H

is about 2.3, indicating that the boundary layer has reattached but is 

not fully recovered from the separation. For P≥20D, H3 at z/P=0, 

indicating the boundary layer has just begun to reattach at this 

location. For P≥20D and higher z/P, H is between 4 and 5.5 indicating 

the boundary layer is still separated. These value are still below the 

baseline value of 6.5, however, indicating that even the widely spaced 

cylinders are effective in keeping the separation bubble thinner. 

 Velocity profiles for the Re=300,000 cases with the larger bar and 

cylinders are shown in Figs. 12-14. Figure 12 shows the profiles at 

z/P=0. The mean profiles show that the boundary layer is attached in 

NASA/CR—2012-217656 139
6 Copyright © 2005 by ASME 

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1

0

0.2

0.4

0.6

0.8

1

y/L

s

U/Ue

(a)

z/P=0

Large Cylinders

Re=50,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0.2

y/L

s

u′/Ue

(b)

z/P=0

Large Cylinders

Re=50,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

Fig. 8 Trailing edge velocity profiles, large cylinders, 

Re=50,000, z/P=0, a) U/Ue, b) u/Ue

all cases. The u peaks are near the wall, as expected for attached 

turbulent boundary layers. The baseline case has the highest u peak, 

indicating that it is the case with the least developed turbulence. The 

bar and cylinders move transition upstream, resulting in a more 

developed turbulent boundary layer by the trailing edge. Both the 

mean and u profiles collapse for the cases with P≥20D. This suggests 

that the cylinders are far enough apart in these cases so that they act as 

if isolated from each other. Both the mean and u profiles show that 

the boundary layer becomes thicker as the cylinder spacing is 

decreased. The P=1D case has a second u peak away from the wall 

that is not present in the two-dimensional bar case. The touching 

cylinders appear to generate more turbulence than the bar. Figure 13 

shows the profiles at z/P=0.5. The profiles for the P=24D and 30D

cases are very similar to the baseline results, again suggesting that 

with large enough spacing the cylinders are effectively isolated from 

each other and cannot directly influence the flow across the entire 

span. As P is reduced, the boundary layer becomes thicker, and u

assumes a more turbulent like shape. Figure 14 shows the profiles at 

several spanwise positions for the P=30D case. The cylinders appear 

to influence the boundary layer u nearly uniformly from z/P=0 to 

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1

0

0.2

0.4

0.6

0.8

1

y/L

s

U/Ue

(a)

z/P=0.5

Large Cylinders

Re=50,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0.2

y/L

s

u′/Ue

(b)

z/P=0.5

Large Cylinders

Re=50,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

Fig. 9 Trailing edge velocity profiles, large cylinders, 

Re=50,000, z/P=0.5, a) U/Ue, b) u/Ue

z/P=20. For z/P>27, the cylinders have little influence and the profiles 

are very similar to the baseline case profiles. The spanwise extent of 

the cylinder influence spreads at a half angle of about 20 degrees for a 

strong effect on the boundary layer and about 30 degrees for some 

effect. 

 Figure 15 shows the shape factor at s/Ls=0.94 for the profiles of the 

Re=300,000 cases with the larger cylinders. The boundary layer is 

attached, so the shape factor is between 1.4 and 1.7 in all cases. 

Figure 16 shows the momentum thickness for these cases. The 

momentum thickness is related to losses in the boundary layer, and in 

cases with equal shape factor and exit flow angle, the momentum 

thickness is directly proportional to profile losses (Howell et al. [21]). 

In the Re=50,000 cases, the large variation in H (Fig. 11) precludes a 

comparison of losses based on θ, but with the smaller range of H at 

Re=300,000 (Fig. 15) the comparison is appropriate. Figure 16 shows 

that the cylinders cause a rise in momentum thickness above the 

baseline case value and that the effect increases as cylinder spacing is 

decreased. Since the separation bubble is small even in the baseline 

case, the cylinders and bars add an unnecessary disturbance and 

NASA/CR—2012-217656 140
7 Copyright © 2005 by ASME 

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1

0

0.2

0.4

0.6

0.8

1

y/L

s

U/Ue

(a)

P/D=30

Large Cylinders

Re=50,000

baseline

2D bar

z/P=0.00

0.06

0.13

0.20

0.27

0.33

0.40

0.47

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0.2

y/L

s

u′/Ue

(b)

P/D=30

Large Cylinders

Re=50,000

baseline

2D bar

z/P=0.00

0.06

0.13

0.20

0.27

0.33

0.40

0.47

Fig. 10 Trailing edge velocity profiles, large cylinders, 

Re=50,000, P/D=30, a) U/Ue, b) u/Ue

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5

1

2

3

4

5

6

7

z/P

H

Large Cylinders

Re=50,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

Fig. 11 Shape factor, H, at trailing edge, large cylinders, 

Re=50,000 

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07

0

0.2

0.4

0.6

0.8

1

y/L

s

U/Ue

(a)

z/P=0

Large Cylinders

Re=300,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07

0

0.02

0.04

0.06

0.08

0.1

0.12

y/L

s

u′/Ue

(b)

z/P=0

Large Cylinders

Re=300,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

Fig. 12 Trailing edge velocity profiles, large cylinders, 

Re=300,000, z/P=0, a) U/Ue, b) u/Ue

increase losses. The more widely spaced cylinders cause less blockage 

and create less of a disturbance, therefore the losses are lower. 

 Velocity profiles for the Re=50,000 cases with the smaller bar and 

cylinders are shown in Fig. 17. Cylinder spacings up to P=5D were 

considered. No variation was observed between the results from 

different spanwise locations, so only the results from z/P=0 are shown. 

The boundary layer did not reattach for cases with P≥2D, although the 

separation bubble was slightly thinner and there was a slight increase 

in the u peak compared to the baseline case. With P=1D the boundary 

layer appears to be on the verge of reattachment, and with the twodimensional bar the boundary layer has just begun to reattach. The 

smaller cylinders are inadequate for control of the boundary layer at 

this Reynolds number, as previously indicated by the pressure profiles. 

 The effect of the smaller cylinders on the velocity profiles at 

Re=300,000 are shown in Fig. 18. No spanwise variation was 

observed, so only results from z/P=0 are shown. The mean profile for 

the P/D=1 case is noticeably different than those for the other cases, 

including the 2D bar case. As was noted above, the touching cylinders 

apparently generate more turbulence that the 2D bar, resulting in a 

thicker boundary layer. As was the case with the larger cylinders (Fig. 

NASA/CR—2012-217656 141
8 Copyright © 2005 by ASME 

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07

0

0.2

0.4

0.6

0.8

1

y/L

s

U/Ue

(a)

z/P=0.5

Large Cylinders

Re=300,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07

0

0.02

0.04

0.06

0.08

0.1

0.12

y/L

s

u′/Ue

(b)

z/P=0.5

Large Cylinders

Re=300,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

Fig. 13 Trailing edge velocity profiles, large cylinders, 

Re=300,000, z/P=0.5, a) U/Ue, b) u/Ue

12), the smaller cylinders cause a drop in the u peak from the baseline 

case value, indicating that the cases with cylinders are closer to fully 

developed turbulent behavior. Shape factors are shown in Fig. 19 as a 

function of spanwise position. There is little variation between cases, 

as expected since the separation bubble is small and the boundary 

layer has fully reattached in all cases. The momentum thickness is 

shown in Fig. 20. All of the cases with cylinders have higher 

momentum thickness than the baseline case. The case with P=1D has 

the highest values, again indicating that the touching cylinders 

generate higher losses than the two-dimensional bars or more widely 

spaced cylinders. 

DISCUSSION 

 The results presented above suggest that the half angle for the 

spreading of the disturbances from the cylinders is about 30 degrees. 

This is a rough estimate due to the finite spacing between pressure tap 

locations and the finite number of cylinder spacings investigated, but 

the half angle roughly agrees with the spreading angle for turbulent 

spots in an adverse pressure gradient, as given by D’Ovidio et al. [22]. 

This angle suggests that a cylinder spacing of between P/D=5 and 

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07

0

0.2

0.4

0.6

0.8

1

y/L

s

U/Ue

(a)

P/D=30

Large Cylinders

Re=300,000

baseline

2D bar

z/P=0.00

0.06

0.13

0.20

0.27

0.33

0.40

0.47

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07

0

0.02

0.04

0.06

0.08

0.1

0.12

y/L

s

u′/Ue

(b)

P/D=30

Large Cylinders

Re=300,000

baseline

2D bar

z/P=0.00

0.06

0.13

0.20

0.27

0.33

0.40

0.47

Fig. 14 Trailing edge velocity profiles, large cylinders, 

Re=300,000, P/D=30, a) U/Ue, b) u/Ue

P/D=10 is necessary to insure flow control across the span. The 

velocity profiles showed good spanwise uniformity for P/D<10. 

 The optimal spacing for spanwise disturbances was investigated in 

an analytical study by Tumin and Ashpis [23]. They considered 

disturbances in both favorable and adverse pressure gradient flows, 

including the favorable pressure gradient found in the upstream region 

of the present test section. Although these results do not apply directly 

to the adverse pressure gradient region of the present study, they can 

be extrapolated to the present conditions. They suggest an optimal 

spacing for maximum disturbance growth in the Re=50,000 cases of 

P/Ls0.03. For the Re=300,000 cases, the optimal spacing would be 

P/Ls0.012. With the present cylinder diameter of 6 mm, these 

spacings correspond to P/D=1 and P/D=0.4 respectively, and could 

not be achieved unless the cylinders were touching. Consistent with 

this, the present results show that the disturbance created by the 

cylinders decreases as the spacing increases. To better test the Tumin 

and Ashpis [23] results, however, smaller diameter, separated 

cylinders with the recommended spacing should be considered. 

 The results at Re=50,000 show that the larger cylinders with the 

closest spacing are needed to effectively control separation. At 

NASA/CR—2012-217656 142
9 Copyright © 2005 by ASME 

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5

1.4

1.45

1.5

1.55

1.6

1.65

1.7

1.75

z/P

H

Large Cylinders

Re=300,000

baseline

2D bar

P/D=1

2

3

5

20

24

30

Fig. 15 Shape factor, H, at trailing edge, large cylinders, 

Re=300,000 

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5

0.8

1

1.2

1.4

1.6

1.8

2

2.2

2.4

2.6

z/P

θ/θb

Large Cylinders

Re=300,000

2D bar

P/D=1

2

3

5

20

24

30

Fig. 16 Momentum thickness at trailing edge, large 

cylinders, Re=300,000 

Re=300,000, the separation bubble is small in the baseline case, and 

cylinders or bars cause transition to move upstream, increasing losses. 

Hence, the thinnest, most widely spaced (which in the limit means 

non-existent) devices are optimal. The cylinders which are best at 

Re=50,000 cause significantly higher losses at Re=300,000. This is 

the same result found by Volino [15] using two-dimensional bars, and 

agrees with other findings in the literature, as noted above. So long as 

the cylinders are close enough to provide spanwise uniformity, it 

appears that varying the cylinder (or bar) thickness and varying the 

cylinder spacing are both effective for controlling the transition 

location and moving it to an optimal location for minimizing losses. 

 The present results can address the question of whether separated 

cylinders provide an advantage over a 2-dimensional bar. A case with 

a 2-dimensional bar that is capable of controlling separation while 

keeping losses to a minimum at a low Reynolds number should be 

compared to a case with cylinders that are thicker than this optimal bar 

but produce the same reattachment and low losses. Cases with these 

same geometries should then be compared at a high Reynolds number 

to see which results in lower high-Re losses. The thin bar in the 

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1

0

0.2

0.4

0.6

0.8

1

y/L

s

U/Ue

(a)

z/P=0

Small Cylinders

Re=50,000

baseline

2D bar

P/D=1

2

3

5

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0.2

y/L

s

u′/Ue

(b)

z/P=0

Small Cylinders

Re=50,000

baseline

2D bar

P/D=1

2

3

5

Fig. 17 Trailing edge velocity profiles, small cylinders, 

Re=50,000, z/P=0, a) U/Ue, b) u/Ue

present study results in marginally reattached flow at Re=50,000. The 

larger cylinders result in spanwise uniform reattachment at Re=50,000 

when P/D=5. The momentum thicknesses at s/Ls=0.94 for these two 

cases are within 8% of each other. At Re=300,000 the thin bar results 

in a momentum thickness 7% above the baseline case value, while the 

thick, P/D=5 cylinders result in a spanwise averaged momentum 

thickness that is 46% above the baseline value. Clearly the thin 2-

dimensional bar is superior to the thicker cylinders. This comparison 

is not completely adequate, however, since the thin bar only causes 

marginal reattachment at Re=50,000, while the thick cylinders result in 

a somewhat more complete reattachment. Volino [15] also considered 

an intermediate bar with thickness twice that of the thin bar considered 

above. At Re=50,000 the intermediate bar induced complete 

reattachment with H=1.8 and θ/θb=1.23. These results are very close 

to those obtained with the large cylinders spaced at P/D=3. At 

Re=300,000 the intermediate bar resulted in H=1.59 and θ/θb=1.29. 

The thick, P/D=3 spaced cylinders resulted in an approximately equal 

shape factor, but θ/θb=1.72. With this better comparison, it is still 

clear that a thin bar is superior to larger cylinders. 

NASA/CR—2012-217656 143
10 Copyright © 2005 by ASME 

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07

0

0.2

0.4

0.6

0.8

1

y/L

s

U/Ue

(a)

z/P=0

Small Cylinders

Re=300,000

baseline

2D bar

P/D=1

2

3

5

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07

0

0.02

0.04

0.06

0.08

0.1

0.12

y/L

s

u′/Ue

(b)

z/P=0

Small Cylinders

Re=300,000

baseline

2D bar

P/D=1

2

3

5

Fig. 18 Trailing edge velocity profiles, small cylinders, 

Re=300,000, z/P=0, a) U/Ue, b) u/Ue

 The present results support the conclusions of Sieverding [10] and 

Zhang and Hodson [17] who found that 2-dimensional bars or trips are 

as good or better than 3-dimensional devices for controlling 

separation. The present results cannot be considered absolutely 

conclusive, however, since only a single geometry was considered 

under a limited number of conditions. The number of possible 

geometries and spacings for three dimensional devices is infinite, so it 

will never be possible to prove conclusively through experiments that 

2-dimensional devices are always better. Further study of devices 

such as the dimples considered by Lake et al. [12] would be useful. 

The effects of high freestream turbulence and unsteady wakes should 

also be considered. Perhaps thinner, smaller diameter cylinders with 

spacings closer to those extrapolated from Tumin and Ashpis [23] 

could provide better separation control with lower losses. This is 

merely speculation. What can be said is that the present results add to 

the evidence that simple 2-dimensional bars are preferable. 

CONCLUSIONS 

 A row of small cylinders located at the suction peak on an LPT 

airfoil were effective for separation control. The reattachment of the 

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5

1.4

1.45

1.5

1.55

1.6

1.65

z/P

H

Small Cylinders

Re=300,000

baseline

2D bar

P/D=1

2

3

5

Fig. 19 Shape factor, H, at trailing edge, small cylinders, 

Re=300,000 

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5

1

1.05

1.1

1.15

1.2

1.25

1.3

1.35

z/P

θ/θb

Small Cylinders

Re=300,000

2D bar

P/D=1

2

3

5

Fig. 20 Momentum thickness at trailing edge, small 

cylinders, Re=300,000 

boundary layer moves upstream as the cylinder height is increased or 

the cylinder spacing is decreased. The half angle of the spreading of 

the disturbance created by the cylinders was of the order 30 degrees. 

This is roughly the same as the expected spreading angle for a 

turbulent spot under the same adverse pressure gradient conditions. 

Based on this angle, the maximum allowable spacing for spanwise 

uniform separation control can be determined. By varying the cylinder 

height and spacing, an optimal reattachment location can be achieved 

for minimum losses at a given Reynolds number. Cylinders optimized 

for low Reynolds numbers resulted in higher losses at high Reynolds 

numbers. The present results add to the evidence that 3-dimensional 

passive flow control devices are not as effective as 2-dimensional bars 

for minimizing losses over a range of Reynolds numbers. 

ACKNOWLEDGEMENTS 

 This work was sponsored by the NASA Glenn Research Center. 

The grant monitor is Dr. David Ashpis. The first author received 

matching support from the Office of Naval Research as a postdoctoral 

fellow. 

NASA/CR—2012-217656 144
11 Copyright © 2005 by ASME 

 

REFERENCES 

[1] Hourmouziadis, J., 1989, “Aerodynamic Design of Low Pressure 

Turbines,” AGARD Lecture Series 167. 

[2] Mayle, R.E., 1991, “The Role of Laminar-Turbulent Transition in 

Gas Turbine Engines,” ASME Journal of Turbomachinery, 113, 

pp. 509-537. 

[3] Sharma, O.P., Ni, R.H., and Tanrikut, S., 1994, “Unsteady Flow 

in Turbines,” AGARD Lecture Series 195, Paper No. 5. 

[4] Bons, J.P., Sondergaard, R., and Rivir, R.B., 2001, “Turbine 

Separation Control Using Pulsed Vortex Generator Jets,” ASME 

Journal of Turbomachinery, 123, pp. 198-206. 

[5] Volino, R.J., and Hultgren, L.S., 2001, “Measurements in 

Separated and Transitional Boundary Layers Under Low-Pressure 

Turbine Airfoil Conditions,” ASME Journal of Turbomachinery, 

123, pp. 189-197. 

[6] Volino, R.J., 2002, “Separated Flow Transition Under Simulated 

Low-Pressure Turbine Airfoil Conditions: Part 1 – Mean Flow 

and Turbulence Statistics,” ASME Journal of Turbomachinery, 

124, pp. 645-655. 

[7] Huang, J., Corke, T., and Thomas, F., 2003, “Plasma Actuators 

for Separation Control on Low Pressure Turbine Blades,” AIAA 

Paper 2003-1027. 

[8] Hultgren, L.S., and Ashpis, D.E., 2003, “Demonstration of 

Separation Delay with Glow Discharge Plasma Actuators,” AIAA 

Paper 2003-1025. 

[9] Volino, R.J., 2003, “Separation Control on Low-Pressure Turbine 

Airfoils Using Synthetic Vortex Generator Jets,” ASME Journal 

of Turbomachinery, 125, pp. 765-777. 

[10] Sieverding, C.H., Bagnera, C., Boege, A.C., Antòn, J.A.C., and 

Luère, V., 2004, “Investigation of the Effectiveness of Various 

Types of Boundary Layer Transition Elements of Low Reynolds 

Number Turbine Bladings,” ASME Paper GT2004-54103. 

[11] Van Treuren, K.W., Simon, T., von Koller, M., Byerley, A.R., 

Baughn, J.W., and Rivir, R., 2002, “Measurements in a Turbine 

Cascade Flow Under Ultra Low Reynolds Number Conditions,” 

ASME Journal of Turbomachinery, 124, pp. 100-106. 

[12] Lake, J.P., King, P.I., and Rivir, R.B., 2000, “Low Reynolds 

Number Loss Reduction on Turbine Blades With Dimples and VGrooves,” AIAA Paper 00-738. 

[13] Murawski, C.G., and Vafai, K., 1999, “Effect of Variable Axial 

Chord on a Low-Pressure Turbine Blade,” Journal of Propulsion 

and Power, 15, pp. 667-674. 

[14] Byerley, A.R., Störmer, O, Baughn, J.W., Simon, T.W., 

VanTreuren, K.W., and List, J., 2002, “Using Gurney Flaps to 

Control Laminar Separation on Linear Cascade Blades,” ASME 

Journal of Turbomachinery, 125, pp. 114-120. 

[15] Volino, R.J., 2003, “Passive Flow Control on Low-Pressure 

Turbine Airfoils,” ASME Journal of Turbomachinery, 125, pp. 

754-764. 

[16] Vera, M., Hodson, H.P., and Vazquez, R., 2004, “The Effects of a 

Trip Wire and Unsteadiness on a High Speed Highly Loaded 

Low-Pressure Turbine Blade,” ASME Paper GT2004-53822. 

[17] Zhang, X.F., and Hodson, H., 2004, “The Combined Effects of 

Surface Trips and Unsteady Wakes on the Boundary Layer 

Development of an Ultra-High-Lift LP Turbine Blade,” ASME 

Paper GT2004-53081. 

[18] Volino, R.J., and Bohl, D.G., 2003, “Separated Flow Transition 

Mechanisms and Prediction With High and Low Freestream 

Turbulence Under Low Pressure Turbine Conditions,” ASME 

Paper GT2004-63360. 

 

[19] Volino, R.J., Schultz, M.P., and Pratt, C.M., 2001, “Conditional 

Sampling in a Transitional Boundary Layer Under High FreeStream Turbulence Conditions,” ASME Journal of Fluids 

Engineering, 125, pp. 28-37. 

[20] Wills, J.A.B., 1962, “The Correction of Hot-Wire Readings for 

Proximity to a Solid Boundary,” Journal of Fluid Mechanics, 12, 

pp. 65-92. 

[21] Howell, R.J., Ramesh, O.N., Hodson, H.P., Harvey, N.W., and 

Schulte, V., 2001, “High Lift and Aft-Loaded Profiles for LowPressure Turbines,” ASME Journal of Turbomachinery, 123, pp. 

181-188. 

[22] D’Ovidio, A., Harkins, J.A., and Gostelow, J.P., “Turbulent Spots 

in Strong Adverse Pressure Gradients: Part 2 – Spot Propagation 

and Spreading Rates,” ASME Paper 2001-GT-0406. 

[23] Tumin, A., and Ashpis, D.E., 2003, “Optimal Disturbances in 

Boundary Layers Subject to Streamwise Pressure Gradients,” 

AIAA Journal, 41, pp. 2297-2300. 

NASA/CR—2012-217656 145
1 

American Institute of Aeronautics and Astronautics

A Comprehensive CFD Study of 

Transitional Flows in Low-Pressure Turbines 

Under a Wide Range of Operating Conditions 

Y. B. Suzen∗, P. G. Huang

University of Kentucky 

Department of Mechanical Engineering 

Lexington, Kentucky 40506-0503 

R. J. Volino

United States Naval Academy 

Department of Mechanical Engineering 

Annapolis, Maryland 21402 

T. C. Corke§, F. O. Thomas¶, J. Huang#

University of Notre Dame 

Department of Aerospace and Mechanical Engineering 

Notre Dame, Indiana 46556-56372 

 J. P. Lake∗∗ P. I. King

 AFRL/PRSS Air Force Institute of Technology 

 Edwards AFB, CA 93524 WPAFB, OH 

Abstract

A transport equation for the intermittency factor is 

employed to predict the transitional flows in lowpressure turbines. The intermittent behavior of the 

transitional flows is taken into account and incorporated 

into computations by modifying the eddy viscosity, μt

, 

with the intermittency factor, γ . Turbulent quantities are 

predicted by using Menter's two-equation turbulence 

model (SST). The intermittency factor is obtained from a 

transport equation model which can produce both the 

experimentally observed streamwise variation of 

intermittency and a realistic profile in the cross stream 

direction. 

The model had been previously validated against lowpressure turbine experiments with success. In this paper, 

the model is applied to predictions of three sets of recent 

low-pressure turbine experiments on PAK-B blade to 

further validate its predicting capabilities under various 

∗

Senior Engineer Associate, Member AIAA. 



Professor, Senior Member AIAA. 



Associate Professor. 

§

Professor. 

¶

Professor. 

#Graduate Assistant. 

∗∗Research Engineer, Member AIAA. 



Associate Professor, Senior Member AIAA 

flow conditions. Comparisons of computational results 

with experimental data are provided. Overall, good 

agreement between the experimental data and 

computational results is obtained. The new model has 

been shown to have the capability of accurately 

predicting transitional flows under a wide range of lowpressure turbine conditions. 

Nomenclature

Cp pressure coefficient, )/()(2

2

− ρ ∞∞ UPP in

Cx axial chord 

FSTI freestream turbulence intensity(%) 

Kt

 flow acceleration parameter, )d/d)(/(

2

ν sUU

k turbulent kinetic energy 

Lx

 axial chord 

N 

 non-dimensional spot breakdown rate 

parameter, /νσθ

3

n t

n spot generation rate 

P static pressure 

Ptotal total pressure 

Re Reynolds number 

Rest

 − Uss est

/)( ν

Reθt

 Uet

/νθ 

s streamwise distance along suction surface 

33rd AIAA Fluid Dynamics Conference

23-26 June 2003, Orlando, Florida AIAA 2003-3591 

Copyright © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

NASA/CR—2012-217656 146
2 

American Institute of Aeronautics and Astronautics

Tu turbulence intensity(%), ′/Uu

U boundary layer streamwise velocity 

Ue local freestream velocity 

Uin inlet freestream velocity 

uτ friction velocity 

W magnitude of vorticity 

n

y distance normal to the wall 

+

y n

uy τ/ν

γ intermittency factor 

θ momentum thickness 

λθ

 

 pressure gradient parameter, 

)d/d)(/(

2

νθ sU

μ molecular viscosity 

μt eddy viscosity 

ν μ / ρ

ν t t

/ ρμ

ρ density 

σ spot propagation parameter 

subscripts 

 e freestream 

 s onset of separation 

 t onset of transition 

1 Introduction

The process of transition from laminar to turbulent flow 

is a major unsolved problem in fluid dynamics and 

aerodynamics. One area where the transition process 

plays an important role and is even more complicated 

due to the diverse flow conditions encountered is the 

low-pressure turbine applications. Transitional flows in 

these applications are affected by several factors such as 

varying pressure gradients, wide range of Reynolds 

number and freestream turbulence variations, flow 

separation, and unsteady wake-boundary layer 

interactions. Accurate simulation and prediction of 

transitional flows under these diverse conditions is key to 

design of more efficient jet engines. 

In low-pressure turbine applications, flow over the blades 

is mostly turbulent at the high Reynolds number 

conditions encountered at take off and the efficiency is at 

its design maximum. However, at lower Reynolds 

number conditions which correspond to high altitudes 

and cruise speeds the boundary layers on the airfoil 

surface have a tendency to remain laminar and hence the 

flow may separate on the suction surface of the turbine 

blades before it becomes turbulent. This laminar 

separation causes unpredicted losses, substantial drops 

in efficiency, and increase in fuel consumption (Mayle, 

1991; Rivir, 1996; Lake et al. 2000). 

In order to calculate the losses and heat transfer on 

various components of gas turbine engines, and to be 

able to improve component efficiencies and reduce losses 

through better designs, accurate prediction of 

development of transitional boundary layers is essential 

(Mayle, 1991). 

One approach proven to be successful for modeling 

transitional flows is to incorporate the concept of 

intermittency into computations. This can be done by 

multiplying the eddy viscosity obtained from a 

turbulence model, μt

, used in the diffusive parts of the 

mean flow equations, by the intermittency factor, γ

(Simon and Stephens, 1991). This method can be easily 

incorporated into any Reynolds Averaged Navier Stokes 

solver. In this approach, the intermittency factor,γ , can 

be obtained from an empirical relation such as the 

correlation of Dhawan and Narasimha (1958), or it can 

be obtained from a transport model. 

Dhawan and Narasimha (1958) correlated the 

experimental data and proposed a generalized 

intermittency distribution function across flow transition. 

Gostelow et al. (1994) extended this correlation to flows 

with pressure gradients under the effects of a range of 

freestream turbulence intensities. Solomon et al. (1995), 

following the work of Chen and Thyson (1971), 

developed an improved method to predict transitional 

flows involving changes in pressure gradients. These 

empirical methods led to a development of transport 

equations for intermittency. 

Steelant and Dick (1996) proposed a transport equation 

for intermittency, in which the source term of the 

equation is developed such that the γ distribution of 

Dhawan and Narasimha (1958) across the transition 

region can be reproduced. Steelant and Dick used their 

model, coupled with two sets of conditioned NavierStokes equations, to predict transitional flows with zero, 

favorable, and adverse pressure gradients. However, 

since their technique involved the solution of two sets of 

strongly coupled equations, the method is not compatible 

with existing CFD codes, in which only one set of 

Navier-Stokes equations is involved. Moreover, the 

model was designed to provide a realistic streamwise γ

behavior but with no consideration of the variation of γ

in the cross-stream direction. 

NASA/CR—2012-217656 147
3 

American Institute of Aeronautics and Astronautics

Cho and Chung (1992) developed a k − ε − γ turbulence 

model for free shear flows. Their turbulence model 

explicitly incorporates the intermittency effect into the 

conventional k − ε model equations by introducing an 

additional transport equation forγ . They applied this 

model to compute a plane jet, a round jet, a plane far 

wake, and a plane mixing layer with good agreements. 

Although this method was not designed to reproduce 

flow transition it provided a realistic profile of γ in the 

cross-stream direction. 

Suzen and Huang (1999) developed an intermittency 

transport equation combining the best properties of 

Steelant and Dick's model and Cho and Chung's model. 

The model reproduces the streamwise intermittency 

distribution of Dhawan and Narasimha (1958) and also 

produces a realistic variation of intermittency in the 

cross-stream direction. This model has been validated 

against T3- series experiments of Savill (1993a,1993b), 

low-pressure turbine experiments of Simon et al. (2000) 

and separated and transitional boundary layer 

experiments of Hultgren and Volino (2000) with success 

(Suzen and Huang, 1999, 2000a, 2000b; Suzen et al. 

2000, 2001, 2002, 2003). 

In this paper we concentrate on prediction of three recent 

low-pressure turbine experiments on the Pratt and 

Whitney's PAK-B blade under low Reynolds number 

conditions using the transport model for intermittency. 

Due to the fact that PAK-B blade is very sensitive to 

changes of flow conditions, it is an ideal test blade for 

validating the transition/turbulence models. The three 

sets of experiments considered are conducted by Lake et 

al. (1999, 2000), Corke et al. (2002), and Volino (2002) 

at three independent facilities. These experiments 

provide an extensive database for investigating 

transitional flows under low pressure turbine conditions 

and are employed as benchmark cases for further testing 

of the predicting capabilities of the current intermittency 

model. A summary of the experiments are given in the 

next section. In section 3, the intermittency transport 

model is presented and implementation of the model and 

the empirical correlations employed for the onset of 

transition are described. In Section 4, the predictions of 

the new intermittency model are compared against the 

experimental data. Conclusions are provided in Section 

5. 

2 Low-Pressure Turbine Experiments

In this paper, we concentrate on computation of three sets 

of low-pressure turbine experiments using the 

intermittency transport model. These experiments are 

conducted by Lake et al. (1999, 2000), Corke et al. 

(2002), and Volino (2002). In these experiments Pratt 

and Whitney’s PAK-B blade is used and the details of 

the blade are shown in Figure 1. Overall these 

experiments cover a Reynolds number range from 10,000 

to 172,000 and the freestream turbulence intensity range 

from 0.08% to 4%. The cases and data used for 

comparison in this paper are summarized in Table 1. In 

the following sections details of these experimental 

efforts are given. 

 

2.1 PAK-B Blade Cascade Experiments of 

Lake et al. (1999, 2000) 

Lake et al. (1999, 2000) conducted experiments on the 

PAK-B blade in order to identify methods for reducing 

separation losses on low-pressure turbine blades under 

low Reynolds number conditions. In the experiments, 

they investigated flows at low Reynolds numbers of 

43,000, 86,000, and 172,000 based on inlet velocity and 

axial chord and freestream turbulence intensities of 1% 

and 4%. These conditions are similar to those 

encountered at high altitude, low speed flight of 

reconnaissance unmanned aerial vehicles used by USAF. 

In Lake’s experiments, surface pressure coefficients, 

boundary layer velocity and turbulence profiles, total 

pressure loss data were obtained at FSTI=1% and 

FSTI=4%. The test set up shown in Figure 2 included 

eight blades with axial chord of 7 inches, and blade 

spacing of 6.2 inches. The blades were numbered 1 

through 8 starting from the inside bend. Boundary layer 

measurements were taken on blade 5 and surface 

pressures were measured around blades 4 and 6. In this 

paper, the PAK-B blade experiments with Reynolds 

numbers of 43,000, 86,000, and 172,000 and freestream 

turbulence intensities of 1% and 4% are computed and 

comparison of pressure distributions between 

experiments and computations are performed. 

2.2 PAK-B Blade Cascade Experiments of 

Corke et al. (2002)

Corke et al. (2002) conducted experiments on PAK-B 

blade cascade for a range of Reynolds numbers and 

turbulence intensities. The Reynolds numbers range from 

10,000 to 100,000 based on inlet velocity and axial chord 

as listed in Table 1. In their experiments the blades had 

an axial chord length of 6.28 inches. The freestream 

turbulence intensity in the tunnel was measured as 

0.08%. In order to increase the turbulence intensity, two 

grids with different mesh sizes were used. One of the 

grids had the mesh size of 2.54cm (denoted as Grid 0) 

and the other had 0.80cm (denoted as Grid 3). The decay 

of turbulence after the grids was measured using 

crosswire and they are shown in Figures 3 and 4 along 

NASA/CR—2012-217656 148
4 

American Institute of Aeronautics and Astronautics

with the computed results for Grid 0 and Grid 3 

respectively. The grids were movable in the tunnel so 

that the turbulence level of the flow that reaches the 

blades could be controlled by moving the grid that is, by 

increasing or decreasing the distance between the grid 

and the blade. Experiments were performed for Reynolds 

numbers 50,000, 75,000, and 100,000, with grids placed 

30 inches away from the blade leading edge, 

corresponding to turbulence intensities of 2.85% and 

1.6% at the leading edge for Grid 0 and Grid 3, 

respectively. For Re=100,000, Grid 0 is placed at 22 

inches and 14 inches corresponding to turbulence 

intensities of 3.62% and 5.2%, respectively. Pressure 

coefficient data is available for all cases and detailed 

boundary layer measurements are available for 

Re=50,000, 75,000, and 100,000 with FSTI=0.08% and 

2.85% cases. The cases and data used for comparisons in 

this paper are listed in Table 1. 

2.3 PAK-B Experiments of Volino (2002)

Volino (2002) investigated the boundary layer 

separation, transition, and reattachment under lowpressure turbine airfoil conditions. The experiments 

included five different Reynolds numbers ranging 

between 10,291 and 123,492 and freestream turbulence 

intensities of 0.5% and 9%. The test section consisted of 

a single passage between two PAK-B blades as shown in 

Figure 5. The axial chord length of the blades was 6.05 

inches. There are flaps located upstream of each blade to 

control the amount of bleed air allowed to escape from 

the passage. These flaps were adjusted by matching 

measured pressure distribution for a high Reynolds 

number with the inviscid pressure distribution on the 

blade. In addition to the upstream bleed flaps, a tailboard 

on the pressure side was used to set the pressure gradient. 

The compiled data include pressure surveys, mean and 

fluctuating velocity profiles, intermittency profiles, and 

turbulent shear stress profiles. It was observed that, the 

effect of high Reynolds number or high free stream 

turbulence level was to move transition upstream. 

Transition started in the shear layer over the separation 

bubble and led to rapid boundary layer reattachment. At 

the lowest Re case, transition did not take place before 

the trailing edge and the boundary layer did not reattach. 

The beginning of transition corresponded to the 

beginning of a significant rise in the turbulent shear 

stress. These experimental results provide detailed 

documentation of the boundary layer and extend the 

existing database to lower Reynolds numbers. The cases 

used for comparisons with computations in this paper are 

listed in Table 1 along with the type of data used for 

comparisons. 

3 Intermittency Transport Model

In this section, the transport model for intermittency is 

presented. The model combines the transport equation 

models of Steelant and Dick (1996) and Cho and Chung 

(1992). Details of the development and implementation 

of the transport model are given in Suzen and Huang 

(1999, 2000a, 2000b), Suzen et al. (2000). 

The model equation is given by: 

( )

( )

( )

(( ) ( ) )

















∂

∂

−+−

∂

∂

+

∂

∂

∂

∂

+









































∂

∂

∂

∂

−

∂

∂

+

−

−

=

∂

∂

+

∂

∂

j

l tt

j

jj

j j

i

k k

i

j

i

ij

kk

j

j

x x

xx

k

C

xx

u

uu

uk

C

x

u

k

C

F

sfsfuuCF

x

u

t

γ

μσγμγσγ

γγ

ε

ρ

γ

ε

τ γρ

γ

ρ

γ

ργ γρ

γ γ 1 1

21 )()(

1

2

3

2/1

2/3

2

1

'

0

 (1) 

The distributed breakdown function, f(s) has the form: 

hsg

bsa s esdsc

sf

′ +

′ + ′ + ′ + ′ +

= 3

234

)(

 (2) 

where t

′ −= sss , and s is the distance along the 

streamline coordinate, and t

s is the transition location. 

The coefficients are; 

U

n

a

σ

= 50 

b = − 4906.0 

5.0

204.0

−













=

U

n

c

σ 

d = 0.0 

5.1

0.04444

−













=

U

n

e

σ

= 10eh 

g = 50 (3) 

The shear stresses are defined as: 

ij ij

k

k

i

j

j

i

tij

k

x

u

x

u

x

u

μτ δρδ

3

2

3

2

−

















∂

∂

−

∂

∂

+

∂

∂

=

 (4) 

The blending function F is constructed using a 

nondimensional parameter, k/Wν where k is the 

turbulent kinetic energy and W is the magnitude of the 

vorticity. The blending function has the form: 

NASA/CR—2012-217656 149
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American Institute of Aeronautics and Astronautics

( ) 















−

= 3.0 1.0

4

200 1

/

tanh

γ

Wk ν

F

 (5) 

The model constants used in Equation 1 are: 

== 0.1 σσ γγ tl

 0.1 C0 = C1 = 6.1 

0 16. C2 = C3 = 15.0

The intermittency is incorporated into the computations 

simply by multiplying the eddy viscosity obtained from 

a turbulence model, μt, by the intermittency factor, γ. 

Simon and Stephens (1991) showed that by 

combining the two sets of conditioned Navier-Stokes 

equations and making the assumption that the Reynolds 

stresses in the nonturbulent part are negligible, the 

intermittency can be incorporated into the 

computations by using the eddy viscosity, μt

*

 which is 

obtained by multiplying the eddy viscosity from a 

turbulence model, μt, with the intermittency factor, γ. 

That is, 

tt = γμμ

*

 (6) 

is used in the mean flow equations. It must be noted 

that γ does not appear in the generation term of the 

turbulent kinetic energy equations. 

Computations of the experiments are performed using a 

recently developed multi-block Navier-Stokes solver, 

called GHOST. The code is developed at University of 

Kentucky, by George Huang, and is a pressure-based 

code based on SIMPLE algorithm with second order 

accuracy in both time and space. Advection terms are 

approximated by a QUICK scheme and central 

differencing is used for the viscous terms. The ‘Rhie and 

Chow’ momentum interpolation method (1983) is 

employed to avoid checkerboard oscillations usually 

associated with the non-staggered grid arrangement. This 

code is capable of handling complex geometries, moving, 

and overset grids and includes multiprocessor 

computation capability using MPI. Since multiple 

processors are used during the computations, it is more 

efficient to divide the computational domain into several 

smaller pieces with very fine grids and distribute the 

zones to processors with the consideration of load 

balancing. This code has been used extensively in a 

recent turbulence model validation effort (Hsu et al., 

2003) and computations of unsteady wake/blade 

interaction (Suzen and Huang, 2003) conducted at 

University of Kentucky. 

In using this intermittency approach, the turbulence 

model selected to obtain μt must produce fully turbulent 

features before transition location in order to allow the 

intermittency to have full control of the transitional 

behavior. Menter's (1994) SST model satisfies this 

requirement. It produces almost fully turbulent flow in 

the leading edge of the boundary layer and therefore it is 

used as a baseline model to compute μt and other 

turbulent quantities in the computations (Suzen et al., 

2000). 

The value of nσ used in evaluating the constants given 

by (3) is provided by the following correlation for zero 

pressure gradient flows (Suzen et al., 2000); 

( )

32 7/411

ˆ 1/ . Tu108

− σ = Unvn σ ×= (7) 

When flows are subject to pressure gradients, the 

following correlation is used: 

( )

( ( ))









>

<

=

−

×−

−

0,10

, 0

ˆ

ˆ

5985.0

7.06

3227

100.75exp1

t

K

t

TuK

ZPG K

M K

n

n

t

t

σ

σ (8) 

with M defined as: 

(850 100 120)

3 5.0

= − +

− − M Tu Tu

where, ( )ZPG nˆσ is the value for flow at zero pressure 

gradient and can be obtained from equation (7) and 

( )( ) t t t K / / dxdUU

2

= ν is the flow acceleration 

parameter. The favorable pressure gradient part of the 

above correlation (for Kt > 0 ) is from Steelant and Dick 

(1996). The portion of the correlation for adverse 

pressure gradient flows for Kt < 0 , is formulated using 

the transition data of Gostelow et al. (1994) and Simon et 

al. (2000) (Suzen et al., 2000). 

The current approach uses the intermittency transport 

model to obtain the intermittency distribution for the 

transitional flows, while the onset of transition is 

defined by correlations. 

The onset of attached flow transition is determined by the 

following correlation in terms of turbulence intensity, Tu, 

and the acceleration parameter, Kt, 

( ) [ ( )]

3/2 5 Re += 150120 04coth . ×− 103

−

θt Tu Kt

 (9) 

where Kt was chosen as the maximum absolute value of 

that parameter in the downstream deceleration region 

(Suzen et al., 2000). This correlation maintains the good 

features of Abu-Ghannam and Shaw (1980) correlation 

in the adverse pressure gradient region and in addition it 

reflects the fact that the flow becomes less likely to have 

transition when subject to favorable pressure gradients by 

rapidly rising as Kt becomes positive. 

NASA/CR—2012-217656 150
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American Institute of Aeronautics and Astronautics

In order to determine the onset of separated flow 

transition Rest is expressed in terms of the turbulence 

intensity (Tu) and the momentum thickness Reynolds 

number at the point of separation (Reθs

) in the form 

(Suzen et al., 2001); 

[ Tu] st s Re874Re p 4.0ex

71.0

= θ − (10) 

This correlation provides a better representation of the 

experimental data than Davis et al. (1987) correlation and 

is used to predict onset of separated flow transition in the 

present computations. 

4 Results and Discussion

4.1 Simulations of Experiments of Lake et al.

(1999,2000)

The intermittency model is applied to predict the PAK-B 

blade experiments of Lake et al. (1999, 2000). In the 

computations, flows at Reynolds numbers of 43,000, 

86,000, and 172,000 based on inlet velocity and axial 

chord with freestream intensities of 1% and 4% were 

investigated. 

The computations are performed using the grid 

consisting of five zones shown in Figure 6. The four 

zones on which the blade grid is superposed each have 

125x225 grid points and the O-type grid around the blade 

has 401x101 points with first y+ less then 0.5. 

The comparisons of computed and experimental pressure 

coefficient distributions are shown in Figures 7(a) 

through (f). In these figures, the experimental 

distributions correspond to the measurements made on 

test blades 4 and 6. 

The computed results compare well with the experiments 

for high turbulence intensity, FSTI=4%, cases shown in 

Figures 7(a), (c) and (e). However, for FSTI=1% cases 

shown in Figures 7(b), (d) and (f) the extent of the 

separation bubbles are under predicted in the 

computations. For example, for Re=86,000, FSTI=1%, 

shown in Figure 7(d), the flow reattaches earlier in 

computations than it does in the experiment as can be 

observed from the difference in the pressure coefficient 

distributions between = 8.0/ Lx x

to 0.85. This 

discrepancy is more pronounced for the lowest Reynolds 

number case, Re=43,000 and FSTI=1% case. 

The onset of separation locations, reattachment locations, 

and onset of transition locations on the suction surface 

are summarized in Table 2 for these cases along with the 

corresponding values from experiments. In the 

experiments, the onset of transition locations and the 

reattachment locations are not reported. 

The experimental onset of separation and reattachment 

points are extracted from the experimental pressure 

coefficient data. The onset of separation is taken to be the 

axial location where the plateau in the pressure 

coefficient distribution of the suction side begins and the 

reattachment point is taken to be the axial location after 

the sharp change in Cp following the plateau. 

The onset of separation, reattachment, and onset of 

transition locations are plotted against Reynolds number 

in Figures 8(a) and (b) for FSTI=4% and 1% 

respectively. For the high turbulence intensity case, 

computation predicts onset of separation and 

reattachment slightly upstream of the experiment. For the 

low FSTI case shown in Figure 8(b) separation zone is 

predicted smaller than the experiments. The onset of 

transition is predicted over the separated flow region in 

the shear layer. From comparison of these figures it is 

evident that with decreasing freestream turbulence 

intensity, the separation zone becomes larger, and for a 

given FSTI condition separated flow region gets smaller 

with increasing Reynolds number. 

4.2 Simulations of Experiments of Corke et 

al. (2002)

In this set of experiments, first the cases with no grid in 

tunnel corresponding to FSTI=0.08% are computed. In 

these computations, the same grid system used for the 

computations of experiments of Lake et al. (1999, 2000) 

shown in Figure 6 is used. 

The comparisons of computed and the experimental 

pressure coefficients are shown in Figures 9(a) through 

(e) for Re=100,000, 75,000, 50,000, 25,000, and 10,000 

based on inlet velocity and axial chord. The agreement 

between the experiments and computations is very good 

for all cases. 

The onset of separation, transition, and reattachment 

locations are tabulated in Table 3 for all cases and plotted 

against Reynolds number in Figures 10(a), (b), and (c) 

for FSTI=0.08%, 1.6%, and 2.85% respectively. 

Computed velocity profiles at seven axial stations along 

the suction surface of the blade are compared to the 

experiments for Re=100,000, 75,000, and 50,000 in 

Figures 11, 12, and 13, respectively. 

For Re=100,000 case, the computed velocity profiles 

compare very well with the experiment as shown in 

NASA/CR—2012-217656 151
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American Institute of Aeronautics and Astronautics

Figures 11(a) through (g). At the first three measurement 

stations, flow is laminar and attached as shown in 

Figures 11(a), (b) and (c). Flow separation takes place at 

= 725.0/Cx x

 and the separated flow region is visible in 

Figures 11(d) and (e) corresponding to axial locations of 

= 75.0/Cx x

 and 0.80. The flow transition and 

reattachment takes place around = 84.0/Cx x in the 

computation. Reattachment location is earlier than the 

experiment which takes place at = .875.0/Cx x In 

Figure 11(f) corresponding to axial station of 

= 85.0/Cx x

 the computed flow field has already 

attached, although the experimental profile indicates a 

very small separation zone close to wall. At = 9.0/Cx x

shown the flow is completely attached as shown in 

Figure 11(g). 

When Reynolds number is reduced to 75,000, the size of 

the separation bubble increases as can be observed from 

the comparison of the velocity profiles shown in Figures 

12(a) through (g). At this Reynolds number the flow 

separates around ≈ 72.0/Cx x

 and reattaches around 

≈ .87.0/Cx x

 The transition onset location is predicted 

at = .854.0/Cx x The size of the separation bubble is 

larger than Re=100,000 case from comparison of Figures 

12(d), (e), and (f) and 11(d), (e), and (f). 

Next, the Reynolds number is reduced to 50,000 and the 

comparison of computed and experimental velocity 

profiles are shown in Figures 13(a) through (g). For this 

case the separation bubble is much larger from the 

previous cases and extends until ≈ 975.0/Cx x

 in the 

experiment and ≈ 93.0/Cx x

in the computations as can 

be seen in Figures 13(d) through (g). Computations 

predicted the transition onset location at = .89.0/Cx x In 

the computations, the onset of separation is predicted 

well in agreement with experiment, however, the 

reattachment point is earlier making the size of the 

separation bubble smaller when compared to experiment. 

This is evident from the comparison of velocity profiles 

at the last two stations shown in Figures 13(f) and (g). 

The onset of separation and reattachment points for 

FSTI=0.08% cases are predicted upstream of the 

experiments as shown in Figure 10(a). 

Next, the high FSTI cases are computed using the six 

zone multiblock grid system shown in Figure 14. The 

computational domain is extended upstream of the blade 

in order to specify the correct turbulence intensity at the 

inlet and to match the decay of turbulence that reaches 

the blade. The matched computed and experimental 

turbulence decays are shown in Figures 3 and 4 for Grid 

0 and Grid 3 respectively. The cases considered have the 

grids placed 30 inches upstream of the blade, 

corresponding to turbulence intensities of 2.85% and 

1.6% for Grid 0 and Grid 3, respectively. 

The comparison of the computed and the experimental 

pressure coefficient distributions for Re=50,000, 75,000, 

and 100,000 for FSTI=2.85% cases are shown in Figure 

15. The agreement is very good between computations 

and experiments. 

Comparisons of computed velocity profiles with the 

experiments for Re=100,000 are given in Figures 16(a) 

through (g). In this case, the flow separates around 

≈ 74.0/Cx x

 and reattaches at ≈ .85.0/Cx x

 The onset 

of transition is predicted at = .806.0/Cx x The computed 

size and extent of the separation bubble is in well 

agreement with the experiment as tabulated in Table 3 

and as can be seen in Figures 10(c), and 16(d), (e), and 

(f). 

For the lower Reynolds number of 75,000, computed 

velocity profiles are compared with the experiments in 

Figures 17(a) through (g). The agreement between 

experiment and computation is well prior to the 

reattachment as shown in Figures 17(a) through (e). 

There is a discrepancy in the reattachment region. The 

flow separation takes place around ≈ 73.0/Cx x

 and

reattaches at ≈ 87.0/Cx x according to the experiment 

whereas computation predicts reattachment earlier at 

around ≈ 84.0/Cx x

 with the onset of transition is 

predicted at = .816.0/Cx x The difference in 

reattachment points is evident in the comparison of the 

computed and experimental velocity profiles shown in 

Figure 17(f). At this station the experimental profile 

indicates separated flow and the computed profile shows 

an already attached flow. 

Next case considered has the same FSTI= 2.85% but 

with Reynolds number being reduced to 50,000. The 

comparison of velocity profiles are shown in Figures 

18(a) through (g). The computations agree well with the 

experiment and the size and extent of the separation 

bubble are well predicted as can be seen from Figure 

10(c). The onset of separation is around ≈ 72.0/Cx x

and 

the flow reattaches around ≈ 9.0/Cx x with transition 

onset at = .837.0/Cx x

In Figure 19 computed and experimental pressure 

coefficient distributions for Grid 3 case are compared for 

Re=50,000, 75,000, and 100,000. Again, very good 

agreement between computations and experiments are 

obtained. The onset of separation and reattachment 

NASA/CR—2012-217656 152
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American Institute of Aeronautics and Astronautics

locations shown in Figure 10(b) compare well with the 

experiments. 

Overall, Figures 10(a), (b), and (c) indicate that as FSTI 

increases, the separated flow region decreases, and at a 

given FSTI, increasing Reynolds number has the same 

effect on the separated flow region. 

4.3 Simulations of PAK B Experiments of 

Volino (2002)

In computation of experiments of Volino (2002) the flow 

field is modeled with the 31 zone multiblock grid shown 

in Figure 20. The bleed flaps below the lower blade and 

above the upper blade are defined by fitting third order 

polynomials through the available points obtained from 

experimental setup and these curves are used as the flap 

shapes in generating the computational grid. Initial 

computations indicated that shape of the bleed flaps and 

the orientation of the tailboard behind the upper blade 

greatly affect the computed results especially the onset of 

separation and reattachment points on the lower blade’s 

suction surface. In order to select the most accurate 

orientation for the tailboard and the shape of the bleed 

flaps, several test computations were performed for the 

case with Re=41,162 and FSTI=0.5% using different 

tailboard orientations and bleed flap shapes. In these 

computations the main goal was to match the 

experimental velocity profiles in the laminar flow part 

and to capture the correct onset point of separation. Once 

an acceptable geometry is obtained, the final bleed flap 

shapes and tailboard orientation is used for computation 

of all other Reynolds number cases. 

Computed pressure coefficient distributions are 

compared to experiments in Figures 21(a) through (d) for 

Re=82,324, 41,162, 20,581, and 10,291 and the 

separation onset, reattachment and transition onset 

information is summarized in Table 4. The Cp

comparison for Re=82,324 shown in Figure 21(a) 

indicates that the computation predicts early reattachment 

of the flow and in the recovery region following 

reattachment pressure coefficient distribution is 

overpredicted. 

The computed pressure coefficient distributions for the 

lower Reynolds number cases shown in Figures 21(c) 

and (d) compare well with experiments. For Re=41,162 

case shown in Figure 21(b) the onset of separation and 

reattachment locations match the experiment as given in 

Table 4, however, in the recovery region pressure 

coefficient distribution is overpredicted. 

Computed velocity profiles are compared to experiment 

at 11 stations along the suction surface of the blade in 

Figures 22(a) through (k) for Re=82,324 and 

FSTI=0.5%. The results compare well with the 

experiment up to = 732.0/Cx x

 shown in Figures 22(a) 

through (g). After this station flow separation takes place. 

Separation onset and reattachment are slightly earlier in 

the computations compared to experiment as given in 

Table 4. This can be also observed from the velocity 

profiles at stations = 798.0/Cx x

 to 0.912 shown in 

Figures 22(h) through (j). Overall computations compare 

well with the experimental measurements. 

Next Reynolds number is reduced to 41,162 and the 

computed and experimental velocity profiles are 

compared in Figures 23(a) through (k). The computed 

profiles agree well with experiments except at 

= 912.0/Cx x

 shown in Figure 23(j). At this station the 

computation indicates a smaller separated flow region 

close to reattachment in contrast to the experiment. 

However, the flow reattaches around = 95.0/Cx x

 both

in computation and experiment and in the next 

measurement station the agreement is well. 

The next case considered has Reynolds number of 

20,581. Computed velocity profiles are shown along with 

the experimental data at 11 axial stations in Figures 24(a) 

through (k). 

In this case flow separates around ≈ 76.0/Cx x

 and does 

not reattach in experiment, however computations 

indicated reattachment at ≈ .98.0/Cx x This discrepancy 

is evident from the comparison of velocity profiles at the 

last two measurement stations shown in Figures 24(j) and 

(k). The computation indicates a smaller separated region 

in these stations and finally reattaches very close to the 

trailing edge. Onset of transition was predicted at 

= .978.0/Cx x

Final case in this set of experiments is the one with 

Re=10,291. The computed velocity profiles are compare 

very well with the experimental data as shown in Figures 

25(a) through (k). In this case the flow separates around 

≈ 76.0/Cx x

and does not reattach. The flow is 

completely laminar, transition was not predicted on the 

blade. 

5 Concluding Remarks

A transport equation for the intermittency factor is 

employed to predict three sets of recent low-pressure 

turbine experiments on PAK-B blade. The intermittent 

behavior of the transitional flows is taken into account by 

modifying the eddy viscosity with the intermittency 

factor. Comparisons of the computed and experimental 

data are made and overall good agreement with the 

NASA/CR—2012-217656 153
9 

American Institute of Aeronautics and Astronautics

experimental data is obtained. The predicting capabilities 

of the current intermittency approach and the 

intermittency transport model in prediction of transitional 

flows under a wide range of low-pressure turbine 

conditions is demonstrated. 

Acknowledgements

This work is supported by NASA Glenn Research Center 

under grant NCC3-590 and followed by grant NCC3-

1040. The project is part of the Low Pressure Turbine 

Flow Physics program of NASA-Glenn. We like to thank 

Dr. David Ashpis for his coordination of the LPT effort. 

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NASA/CR—2012-217656 154
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AIAA Aerospace Science Meeting & Exhibit, 6-9 Jan, 2003, 

Reno, NV. 

Suzen, Y.B., Xiong, G., Huang, P.G., 2000, “Predictions of 

Transitional Flows in Low-Pressure Turbines Using an 

Intermittency Transport Equation,” AIAA-2000-2654, Fluids 

2000, Denver, CO. 

Suzen, Y.B., and Huang, P.G., Hultgren, L.S., Ashpis, D.E., 

2001, “Predictions of Separated and Transitional Boundary 

Layers Under Low-Pressure Turbine Airfoil Conditions Using 

an Intermittency Transport Equation,” AIAA-2001-0446, 39th

AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV. 

Suzen, Y.B., and Huang, P.G., Hultgren, L.S., Ashpis, D.E., 

2003, “Predictions of Separated and Transitional Boundary 

Layers Under Low-Pressure Turbine Airfoil Conditions Using 

an Intermittency Transport Equation,” ASME Journal of 

Turbomachinery, accepted for publishing. 

Suzen, Y.B., Xiong, G., Huang, P.G., 2002, “Predictions of 

Transitional Flows in Low-Pressure Turbines Using an 

Intermittency Transport Equation,” AIAA Journal, Vol. 40, No 

2, pp. 254-266. 

Volino, R.J., 2002, "Separated Flow Transition Under 

Simulated Low-Pressure Turbine Airfoil Conditions: Part 1- 

Mean Flow and Turbulence Statistics," ASME-GT-30236, 

ASME TURBO EXPO 2002, Amsterdam, The Netherlands. 

Source Test Section 

Cx

(inch) 

Re 

(UinCx/ν) 

FSTI (%) 

Data used for 

Comparison 

Lake et.al (1999, 2000) P&W PAK-B cascade 7.0 43,000 1 & 4 Cp distribution 

 86,000 1 & 4 Cp distribution 

 172,000 1 & 4 Cp distribution 

Corke et al. (2002) P&W PAK-B cascade 6.28 10,000 0.08 Cp distribution 

 25,000 0.08 Cp distribution 

 50,000 0.08, 1.6, 2.85 

Cp distribution, 

velocity profiles* 

 75,000 0.08, 1.6, 2.85 

Cp distribution, 

velocity profiles* 

 100,000 0.08, 1.6, 2.85 

Cp distribution, 

velocity profiles* 

Volino (2002) 

P&W PAK-B single 

passage 

6.05 10,291 0.5 

Cp distribution, 

velocity profiles 

 20,581 0.5 

Cp distribution, 

velocity profiles 

 41,162 0.5 

Cp distribution, 

velocity profiles 

 82,324 0.5 

Cp distribution, 

velocity profiles 

* Velocity profiles are available for FSTI=0.08% and 2.85% from experiments. 

Table 1. Details of the experiments used for comparison with computations. 

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Re 

(UinCx/ν) 

FSTI 

(%) 

xs/Cx 

(Computation) 

xs/Cx 

(Experiment) 

xr/Cx 

(Computation) 

xr/Cx 

(Experiment) 

xtr/Cx 

(Computation) 

172,000 4 0.732 0.74 0.83 0.84 0.806 

86,000 4 0.725 0.74 0.86 0.88 0.832 

43,000 4 0.718 0.73 0.93 0.93 0.881 

172,000 1 0.728 0.72 0.82 0.84 0.808 

86,000 1 0.722 0.72 0.87 0.90 0.849 

43,000 1 0.711 0.62 0.95 0.99 0.896 

 

Table 2: Separation, reattachment, and transition locations for cases of Lake et al. (1999, 2000). 

Re 

(UinCx/ν) 

FSTI 

(%) 

xs/Cx 

(Computation) 

xs/Cx 

(Experiment) 

xr/Cx 

(Computation) 

xr/Cx 

(Experiment) 

xtr/Cx 

(Computation) 

10,000 0.08 0.661 0.725 --- --- --- 

25,000 0.08 0.656 0.725 0.980 --- 0.936 

50,000 0.08 0.714 0.725 0.925 0.975 0.890 

75,000 0.08 0.718 0.725 0.860 0.870 0.854 

100,000 0.08 0.725 0.725 0.840 0.875 0.840 

50,000 1.6 0.722 0.728 0.900 0.900 0.854 

75,000 1.6 0.728 0.730 0.867 0.875 0.834 

100,000 1.6 0.732 0.730 0.860 0.877 0.821 

50,000 2.85 0.728 0.722 0.887 0.900 0.837 

75,000 2.85 0.732 0.729 0.840 0.870 0.816 

100,000 2.85 0.735 0.734 0.842 0.850 0.806 

Table 3: Separation, reattachment, and transition locations for cases of Corke et al. (2002). 

Re 

(UinCx/ν) 

FSTI 

(%) 

xs/Cx 

(Computation) 

xs/Cx 

(Experiment) 

xr/Cx 

(Computation) 

xr/Cx 

(Experiment) 

xtr/Cx 

(Computation) 

10,291 0.5 0.760 0.750 --- --- --- 

20,581 0.5 0.765 0.760 0.980 --- 0.978 

41,162 0.5 0.760 0.770 0.950 0.950 0.840 

82,324 0.5 0.757 0.767 0.890 0.900 0.857 

Table 4: Separation, reattachment, and transition locations for cases of Volino (2002). 

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Figure 1: P&W PAK-B blade cascade details. 

Figure 2: Experimental set up for used by Lake et al. (1999, 2000). 

Chord length, L 

Axial chord length, Lx

Axial chord to chord ratio, Lx/L=0.906

Pitch to chord ratio, P/L = 0.8 

Blade inlet angle, β1 = 35o 

Blade inlet angle, β2 = -60o

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Distance(inches)

Turbulence Intensity (%)

12 14 16 18 20 22 24 26 28 30 32 1

2

3

4

5

6

7

Computation

Experiment

Turbulence IntensityforGrid0 (Meshsize=1inch)

Figure 3: Comparison of computed and experimental decay of turbulence for experiments of Corke et al. (2002), 

with Grid 0. 

Distance(inches)

Turbulence Intensity (%)

12 14 16 18 20 22 24 26 28 30 32 1

1.5

2

2.5

3

3.5

4

Computation

Experiment

TurbulenceIntensityforGrid3 (Meshsize=0.80cm)

Figure 4: Comparison of computed and experimental decay of turbulence experiments of Corke et al. (2002), for 

Grid 3. 

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Figure 5: Schematic of the test section for experiments of Volino(2002). 

x / Cx

y / Cx

-0.5 0 0.5 1 1.5

-1

-0.5

0

0.5

1

Figure 6: Multiblock grid used for computations of experiments of Lake et al. (1999, 2000) and FSTI=0.08% 

experiments of Corke et al. (2002) . 

NASA/CR—2012-217656 159
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x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Blade 4

Blade 6

Computation

(a) Re=172,000, FSTI=4%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Blade 4

Blade 6

Computation

(b) Re=172,000, FSTI=1%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Blade 4

Blade 6

Computation

(c) Re=86,000, FSTI=4%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Blade 4

Blade 6

Computation

(d) Re=86,000, FSTI=1%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Blade 4

Blade 6

Computation

(e) Re=43,000, FSTI=4%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Blade 4

Blade 6

Computation

(f) Re=43,000, FSTI=1%

Figure 7: Comparison of computed pressure coefficient with experiments of Lake et al. (1999, 2000). 

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Re (Uin Cx/ ν )

x / Cx

0 50000 100000 150000 200000 0.6

0.7

0.8

0.9

1

Onset of separation (Comp.)

Onset of separation (Exp.)

Reattachment (Comp.)

Reattachment (Exp.)

Onset of transition (Comp.)

(a) FSTI = 4%

Re (Uin Cx/ ν )

x / Cx

0 50000 100000 150000 200000 0.6

0.7

0.8

0.9

1

Onset of separation (Comp.)

Onset of separation (Exp.)

Reattachment (Comp.)

Reattachment (Exp.)

Onset of transition (Comp.)

(b) FSTI = 1%

Figure 8: Comparison of separation, reattachment and transition locations for experiments of Lake et al. (1999, 

2000). 

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x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(a) Re =100,000, FSTI=0.08%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(b) Re =75,000, FSTI=0.08%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(c) Re =50,000, FSTI=0.08%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(d) Re =25,000, FSTI=0.08%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(e) Re =10,000, FSTI=0.08%

Figure 9: Comparison of computed pressure coefficients with experiments of Corke et al. (2002) for 

FSTI=0.08% cases. 

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Re (Uin Cx/ ν )

x / Cx

0 50000 100000 150000 200000 0.6

0.7

0.8

0.9

1

Onset of separation (Comp.)

Onset of separation (Exp.)

Reattachment (Comp.)

Reattachment (Exp.)

Onset of transition (Comp.)

(b) FSTI = 1.6%

Re (Uin Cx/ ν )

x / Cx

0 50000 100000 150000 200000 0.6

0.7

0.8

0.9

1

Onset of separation (Comp.)

Onset of separation (Exp.)

Reattachment (Comp.)

Reattachment (Exp.)

Onset of transition (Comp.)

(c) FSTI = 2.85%

Re (Uin Cx/ ν )

x / Cx

0 50000 100000 150000 200000 0.6

0.7

0.8

0.9

1

Onset of separation (Comp.)

Onset of separation (Exp.)

Reattachment (Comp.)

Reattachment (Exp.)

Onset of transition (Comp.)

(a) FSTI = 0.08%

Figure 10: Comparison of separation, reattachment and transition locations for experiments of Corke et al. 

(2002).

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U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.5

(a)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.6

(b)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.7

(c)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.75

(d)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.80

(e)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.85

(f)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.9

(g)

Figure 11: Comparison of computed velocity profiles with experiments of Corke et al. (2002), Re=100,000, 

FSTI=0.08% case. 

NASA/CR—2012-217656 164
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American Institute of Aeronautics and Astronautics

U / Uin

yn (mm)

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.5

(a)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.6

(b)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.7

(c)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.75

(d)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.80

(e)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.85

(f)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.9

(g)

Figure 12: Comparison of computed velocity profiles with experiments of Corke et al. (2002), Re=75,000, 

FSTI=0.08% case. 

NASA/CR—2012-217656 165
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American Institute of Aeronautics and Astronautics

U / Uin

yn (mm)

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.5

(a)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.6

(b)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.7

(c)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.75

(d)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.80

(e)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.85

(f)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.9

(g)

Figure 13: Comparison of computed velocity profiles with experiments of Corke et al. (2002), Re=50,000, 

FSTI=0.08% case. 

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x / Cx

y / C

x

-2 -1 0 1

-2

-1

0

1

Figure 14: Grid used for computation of experiments of Corke et al. (2002) with FSTI=1.6% and 2.85%. 

NASA/CR—2012-217656 167
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x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(a) Re = 50,000, FSTI = 2.85%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(b) Re = 75,000, FSTI = 2.85%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(c) Re = 100,000, FSTI = 2.85%

Figure 15: Comparison of computed pressure coefficients with experiments of Corke et al. (2002) for 

FSTI=2.85% cases.

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U / Uin

yn (mm)

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.5

(a)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.6

(b)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.7

(c)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.75

(d)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.80

(e)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.85

(f)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.9

(g)

Figure 16: Comparison of computed velocity profiles with experiments of Corke et al. (2002), Re=100,000, 

FSTI=2.85% case. 

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U / Uin

yn (mm)

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.5

(a)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.6

(b)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.7

(c)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.75

(d)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.80

(e)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.85

(f)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.9

(g)

Figure 17: Comparison of computed velocity profiles with experiments of Corke et al. (2002), Re=75,000, 

FSTI=2.85% case. 

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U / Uin

yn (mm)

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.5

(a)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.6

(b)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.7

(c)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.75

(d)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.80

(e)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.85

(f)

U / Uin

yn / Cx

0 1 2

0

0.02

0.04

0.06

0.08

0.1

x / cx= 0.9

(g)

Figure 18: Comparison of computed velocity profiles with experiments of Corke et al. (2002), Re=50,000, 

FSTI=2.85% case. 

 

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x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(a) Re = 50,000, FSTI = 1.6%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(b) Re = 75,000, FSTI = 1.6%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(c) Re = 100,000, FSTI = 1.6%

Figure 19: Comparison of computed pressure coefficients with experiments of Corke et al. (2002) for 

FSTI=1.6% cases. 

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x / Cx

y / C

x

-1 0 1 2 3 4

-1

0

1

2

3

Figure 20: 31 zone multi-block grid used for computation of experiments of Volino (2002). 

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x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(c) Re = 20,581, FSTI = 0.5%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(d) Re = 10,291, FSTI = 0.5%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(a) Re = 82,324, FSTI = 0.5%

x / Cx(%)

Cp

0 25 50 75 100

-4

-3

-2

-1

0

1

Computation

Experiment

(b) Re = 41,162, FSTI = 0.5%

Figure 21: Comparison of computed pressure coefficient distributions with experiments of Volino (2002), 

FSTI=0.5%. 

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U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.093

(a)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.197

(b)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.314

(c)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.435

(d)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.550

(e)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.651

(f)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.732

(g)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.798

(h)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.856

(i)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.912

(j)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.964

(k)

Figure 22: Comparison of computed velocity profiles with experiments of Volino (2002), Re=82,324, FSTI=0.5%. 

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U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.093

(a)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.197

(b)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.314

(c)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.435

(d)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.550

(e)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.651

(f)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.732

(g)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.798

(h)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.856

(i)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.912

(j)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.964

(k)

Figure 23: Comparison of computed velocity profiles with experiments of Volino (2002), Re=41,162, FSTI=0.5%. 

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American Institute of Aeronautics and Astronautics

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.964

(k)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.197

(b)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.093

(a)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.314

(c)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.435

(d)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.550

(e)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.651

(f)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.732

(g)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.798

(h)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.856

(i)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.912

(j)

Figure 24: Comparison of computed velocity profiles with experiments of Volino (2002), Re=20,581, FSTI=0.5%. 

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U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.964

(k)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.093

(a)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.197

(b)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.314

(c)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.435

(d)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.550

(e)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.651

(f)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.732

(g)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.798

(h)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.856

(i)

U / Uin

yn / Cx

0 2

0

0.02

0.04

0.06

0.08

0.1

x / Cx= 0.912

(j)

Figure 25: Comparison of computed velocity profiles with experiments of Volino (2002), Re=10,291, FSTI=0.5%. 

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Abstract of presentation at the 55th Annual Meeting of the Division of Fluid Dynamics of the American Physical 

Society, Dallas, TX, 2002 

Separation Control Using Synthetic Vortex Generator Jets 

Ralph Volino (U.S. Naval Academy) 

Boundary layer separation is a problem on some low-pressure turbine airfoils in aircraft engines. Airfoils are 

typically designed for optimal performance at takeoff, where maximum power is needed. When the aircraft 

climbs, however, the lower pressure at altitude (and lower Re) can result in separation from the suction side of 

the airfoils. Flow control would be desirable to minimize separation at low Re, without sacrificing performance at 

high Re. In the present experiments, synthetic vortex generator jets are used for flow control. This is the first 

known application of synthetic vortex generator jets. At low Reynolds numbers, without flow control, a large 

separation bubble is present. With the jets, the bubble is eliminated. The time averaged boundary layer appears 

laminar and attached. Ensemble averaged data shows that during each jet pulsing cycle, a turbulent patch 

moves down the airfoil, followed by a calmed period of attached laminar flow. Losses are lower than for a case in 

which reattachment was forced with a passive boundary layer trip. Smoke visualization as well as animations of 

ensemble averaged hot-wire data from boundary layer profiles will be presented. 

-------------------------------------------------------------------------------------------------------------------------------------------- 

Abstract of presentation at the 56th Annual meeting of the Division of Fluid Dynamics of the American Physical 

Society, East Rutherford, NJ, 2003 

Separated Flow Transition Mechanism with Passive Flow Control under Low Pressure Turbine 

Conditions 

Ralph J. Volino, Douglas G. Bohl (U. S. Naval Academy) 

Boundary layer separation and transition have been studied experimentally on the suction side of a simulated 

low pressure turbine passage. Rectangular ribs were used as passive flow control devices to induce transition 

and promote reattachment. Thin ribs promote reattachment after a small separation bubble, resulting in lower 

aerodynamic losses. The ribs appear to have no immediate effect on the mean or rms streamwise velocity 

profiles, but do impart a very small disturbance in the boundary layer. The boundary layer separates, as in the 

unmodified flow, but undergoes transition and reattachment upstream of the location in the unmodified flow. 

Measurements of the wall normal velocity show the regular growth of small disturbances in the pre-transitional 

flow in cases with and without ribs. The transition mechanism appears similar in all cases. Ribs and elevated 

freestream turbulence increase the magnitude of the initial disturbance, leading to an earlier attainment of a 

disturbance large enough to initiate transition. These observations lead to a new correlation for transition onset 

location which accounts for rib height, freestream turbulence level, and Reynolds number 

-------------------------------------------------------------------------------------------------------------------------------------------- 

Abstract of presentation at the 57th Annual Meeting of the Division of Fluid Dynamics of the American Physical 

Society, Seattle, WA, 2004 

Three-Dimensional Flow Structure Resulting from Oscillating Vortex Generator Jets 

Ralph Volino, Douglas Bohl (U.S. Naval Academy) 

A row of oscillating vortex generator jets is used in an experimental study to control boundary layer separation 

from an airfoil. Without flow control a large, burst separation bubble is present at low Reynolds numbers. With 

the jets active, the separation is eliminated. Ensemble averaged data (relative to the jet pulsing) show a turbulent 

patch moving down the surface after each outward jet pulse, followed by an extended calmed period 

characterized by a thin, attached laminar boundary layer. In the present study, profiles of streamwise and wall 

normal velocity have been measured at multiple spanwise locations between adjacent jets. The structure of the 

streamwise vortices is clear, with upwash and downwash regions. Ensemble averaged Reynolds shear and 

normal stresses, wall skin friction, mean velocity and integral quantities have been documented. Animations of 

the data show the generation of the vortices and their progression downstream, followed by periods of laminar, 

spanwise-uniform flow. The action of the vortices in bringing high speed fluid into the near wall region to achieve 

the separation control is illustrated 

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ACTIVE AND PASSIVE FLOW CONTROL ON LOW PRESSURE TURBINE AIRFOILS 

Modern low-pressure turbine airfoils are subject to increasingly stronger pressure gradients as designers 

impose higher loading to improve efficiency and lower cost by reducing the number of airfoils in an 

engine. If the adverse pressure gradient on the suction side becomes strong enough, the boundary layer 

will separate. Separation bubbles, particularly those which fail to reattach, can result in a significant 

degradation of engine efficiency. The problem is particularly relevant in aircraft engines. Airfoils 

optimized to produce maximum power under takeoff conditions may still experience separation at cruise 

conditions, due to the thinner air and lower Reynolds numbers at altitude. An efficiency drop of 2% may 

occur between takeoff and cruise in large commercial transport engines, and the difference could be larger 

in smaller engines operating at higher altitudes. Needed is a means of controlling separation at low Re, 

without sacrificing the gains achieved at high Re. 

In the present study, passive and active flow control are applied to the suction surface boundary layer on 

an LP turbine airfoil. Experiments are conducted in a single passage cascade simulator. Reynolds 

numbers (based on exit velocity and suction surface length) from 25,000 to 300,000 are considered under 

both high (8% inlet) and low (0.5%) free-stream turbulence (FSTI) conditions. In the passive control 

experiments, thin rectangular bars are applied to the airfoil near the suction surface velocity peak. Bars 

that are sufficiently large immediately trip the boundary layer to turbulent and prevent separation. 

Smaller bars initially appear to have little or no effect, and the boundary layer separates. Some distance 

downstream, however, small disturbances induced by the bars induce transition in the shear layer over the 

separation bubble, causing reattachment to move upstream relative to its location in the unmodified flow. 

The cases with the shortened separation bubbles appear to have lower losses than those with the larger 

trips. Bars which produce optimal results at low Re, however, invariably cause higher losses at the 

highest Re, suggesting the possible benefit of active flow control. 

Active control is achieved using synthetic (oscillating, i.e. no net mass flow) vortex generator jets. An 

airfoil was constructed with a central cavity and a spanwise row of small holes extending from the cavity 

to the suction surface. The cavity is pulsed with a loudspeaker, causing jets to enter the boundary layer at 

a compound angle relative to the blade surface and the main flow. A single case has been documented to 

date with Re=25,000 and low FSTI. The separation bubble is completely eliminated, as shown through 

smoke visualization and animations of phase locked quantitative data. Ensemble averaged data (relative 

to the jet pulsing) show a turbulent patch moving down the blade after each outward jet pulse, followed 

by an extended "calmed" period characterized by a thin, attached laminar boundary layer. Losses appear 

substantially lower than with passive control. 

NOTE: The following pages were excerpted from Minnowbrook IV—2003 Workshop on Transition and 

Unsteady Aspects of Turbomachinery Flows, J.E. LaGraff and D.E. Ashpis, Editors, NASA/TM—2004-

212913, August 2004, pp. 78–97. 

Ralph J. Volino 

United States Naval Academy 

Annapolis, Maryland 21402 
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4. TITLE AND SUBTITLE

Experimental Studies of Low-Pressure Turbine Flows and Flow Control 

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This manuscript has been authored by an employee of the United States Naval Academy under a NASA Interagency Agreement No. C-31011-K with the National Aeronautics 

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13. SUPPLEMENTARY NOTES

This report contains a supplemental CD-ROM containing processed data in ASCII format of streamwise pressure profiles and velocity 

profiles, along with the full text PDF file of this report. The CD-ROM can be obtained from the Center for AeroSpace Information (CASI) 

Web site at http://www.sti.nasa.gov 

Grant technical monitor, David E. Ashpis, Aeropropulsion Division, Glenn Research Center, organization code RTTO, ashpis@nasa.gov 

14. ABSTRACT

This report summarizes research performed in support of the NASA Glenn Research Center (GRC) Low-Pressure Turbine (LPT) Flow 

Physics Program. The work was performed experimentally at the U.S. Naval Academy faculties. The geometry corresponded to “Pak B” 

LPT airfoil. The test section simulated LPT flow in a passage. Three experimental studies were performed: (a) Boundary layer 

measurements for ten baseline cases under high and low freestream turbulence conditions at five Reynolds numbers of 25,000, 50,000, 

100,000, 200,000, and 300,000, based on passage exit velocity and suction surface wetted length; (b) Passive flow control studies with three 

thicknesses of two-dimensional bars, and two heights of three-dimensional circular cylinders with different spanwise separations, at same 

flow conditions as the 10 baseline cases; (c) Active flow control with oscillating synthetic (zero net mass flow) vortex generator jets, for one 

case with low freestream turbulence and a low Reynolds number of 25,000. The Passive flow control was successful at controlling the 

separation problem at low Reynolds numbers, with varying degrees of success from case to case and varying levels of impact at higher 

Reynolds numbers. The active flow control successfully eliminated the large separation problem for the low Reynolds number case. Very 

detailed data was acquired using hot-wire anemometry, including single and two velocity components, integral boundary layer quantities, 

turbulence statistics and spectra, turbulent shear stresses and their spectra, and intermittency, documenting transition, separation and 

reattachment. Models were constructed to correlate the results. The report includes a summary of the work performed and reprints of the 

publications describing the various studies. 

15. SUBJECT TERMS

Gas turbines; Turbomachinery; Turbine; Low-pressure turbine; Turbulence; Flow control; Synthetic jets; Wakes; Separation; Vortex 

generated jets; Boundary layer; Tripps 

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