M. Samimy, J. Kastner, J.-H. Kim, Y. Utkin, and I. Adamovich

The Ohio State University, Columbus, Ohio

C.A. Brown

Glenn Research Center, Cleveland, Ohio

Flow and Noise Control in High Speed and High

Reynolds Number Jets Using Plasma Actuators

NASA/TM—2006-214367

December 2006

AIAA–2006–2846
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Prepared for the

Third Flow Control Conference

sponsored by the American Institute of Aeronautics and Astronautics

San Francisco, California, June 5–8, 2006

M. Samimy, J. Kastner, J.-H. Kim, Y. Utkin, and I. Adamovich

The Ohio State University, Columbus, Ohio

C.A. Brown

Glenn Research Center, Cleveland, Ohio

Flow and Noise Control in High Speed and High

Reynolds Number Jets Using Plasma Actuators

NASA/TM—2006-214367

December 2006

AIAA–2006–2846
Acknowledgments

The support of the research at The Ohio State University by the NASA Glenn Research Center and the Ohio Department of

Development is greatly appreciated. We greatly appreciate the help from James Bridges in many facets of this research.

Available from

NASA Center for Aerospace Information

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Hanover, MD 21076–1320

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only. Their usage does not constitute an official endorsement,

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at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management.
NASA/TM—2006-214367 1 

Flow and Noise Control in High Speed and High Reynolds Number 

Jets Using Plasma Actuators 

M. Samimy, J. Kastner, J.-H. Kim, Y. Utkin, and I. Adamovich 

The Ohio State University 

Columbus, Ohio 43235 

C.A. Brown 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The idea of manipulating flow to change its characteristics is over a century old. Manipulating 

instabilities of a jet to increase its mixing and to reduce its radiated noise started in the 1970s. While the 

effort has been successful in low-speed and low Reynolds number jets, available actuators’ capabilities in 

terms of their amplitude, bandwidth, and phasing have fallen short in control of high-speed and high 

Reynolds number jets of practical interest. Localized arc filament plasma actuators have recently been 

developed and extensively used at Gas Dynamics and Turbulence Laboratory (GDTL) for control of highspeed and high Reynolds number jets. While the technique has been quite successful and is very 

promising, all the work up to this point had been carried out using small high subsonic and low 

supersonic jets from a 2.54 cm diameter nozzle exit with a Reynolds number of about a million. The 

preliminary work reported in this paper is a first attempt to evaluate the scalability of the technique. The 

power supply/plasma generator was designed and built in-house at GDTL to operate 8 actuators 

simultaneously over a large frequency range (0 to 200 kHz) with independent control over phase and duty 

cycle of each actuator. This allowed forcing the small jet at GDTL with azimuthal modes m = 0, 1, 2, 3, 

±1, ±2, and ±4 over a large range of frequencies. This power supply was taken to and used, with minor 

modifications, at the NASA Nozzle Acoustic Test Rig (NATR). At NATR, 32 actuators were distributed 

around the 7.5 in. nozzle (a linear increase with nozzle exit diameter would require 60 actuators). With 

this arrangement only 8 actuators could operate simultaneously, thus limiting the forcing of the jet at 

NATR to only three azimuthal modes m = ±1, 4, and 8. Very preliminary results at NATR indicate that 

the trends observed in the larger NASA facility in terms of the effects of actuation frequency and 

azimuthal modes are similar in both small GDTL and larger NASA jets. However, the actuation authority 

seems to fall short in the larger jet at higher Mach numbers, resulting in decreased amplitude response 

compared to the small jet, which is attributed at this point to the lack of sufficient number of actuators. 

The preliminary results seem also to suggest that amplitude of actuation tones is similar in both the small 

and larger jets. 

 Introduction

Excitation of Jets for Aeroacoustics 

The idea of manipulating the turbulence of jets for engineering benefit has been around for many 

decades. The promises for reducing jet noise, increasing mixing efficiency, and other applications have 

lead many to pursue actuator technologies which would bring these promises to fruition. Along the way, 

many problems have been found, often associated with the actuators being used. 
NASA/TM—2006-214367 2 

The Promise 

For jet noise, control of jet turbulence has had two potential benefits: enhanced understanding of 

noise generation mechanisms and direct manipulation of these mechanisms for noise reduction. It is 

widely observed that jet noise scales as a high power of jet velocity Uj. The acknowledged Uj

8

 law can be 

viewed as a conversion law for transforming energy from turbulence to acoustics (Lighthill 1952, 1954). 

The coefficient in front of this exponential factor is also known to be very small, somewhere around 10–7. 

The happy fact that most turbulent kinetic energy does not convert to sound is a result of the conservation 

laws for angular momentum, which at low Mach numbers (M~0) has each noise-producing impulse 

countered by another, oppositely directed impulse. This cancellation becomes less complete as Mach 

number increases, but still remains nearly perfect as one compares the energy available with that radiated. 

It also explains why the study of jet noise remains active today as researchers strive to understand and 

model this one in a few million portion of the flow which does radiate sound. It has also been shown 

theoretically that, for small Mach numbers, high order azimuthal undulations of the jet shear layer do not 

couple effectively with the acoustic field, making them particularly inefficient noise generators (Michalke 

and Fuchs 1975). The other aspect of jet noise research that has precluded understanding is that the 

turbulence has only been studied in statistical fashion, the motions which produce the noise being nondeterministic. 

To date, almost all jet noise reduction has been accomplished by reducing the jet exit velocity. This 

has been done by increasing bypass ratios of the engines, through improvements in materials and engine 

core architecture. But what if the efficiency of the conversion of turbulence to noise, while admittedly 

small, could be reduced even further? One possibility is to remove the statistically rare, but acoustically 

significant events that have been found to produce most of the noise in the jet. Perhaps if the turbulent 

structures could be controlled to further maximize the cancellation between impulse-carrying structures, 

the efficiency would be reduced even further. Another possibility is to pump more of the jet’s turbulent 

kinetic energy into high-order azimuthal modes, putting more of the kinetic energy into dynamic motions 

that are inefficient at making noise. Both of these possibilities require actuators with control authority 

greater than that imposed by the non-deterministic turbulence found in the jet. 

Even if ultimately we cannot find a way to reduce noise by controlling flow structures in a jet, it is 

still of interest to study how these structures relate to the generation of sound in jets. The promise of jet 

excitation for jet noise theory is that if the researcher could somehow remove the non-deterministic 

portion of the problem, then the mechanisms of noise production could be studied directly and with much 

greater precision. This is a much more realistic application of jet control technology, as requiring control 

of laboratory jets is somewhat more reasonable than implementing control on jet aircraft with their 

multiple constraints. 

The Problems 

The insight of Brown and Roshko (1974) that shear layer turbulence is composed of large, wellorganized structures even in high Reynolds numbers is usually credited with starting the study of largescale, ‘coherent’ structures and their control. During the 1970’s and 1980’s many studies were done 

where free shear flows, particularly jets, were perturbed using actuators, and their response studied (Crow 

and Champagne 1971, Moore 1977, Kibens 1980, Zaman and Hussain 1981, Ho and Huerre 1984, Ahuja 

and Blakney 1985). Most of these studies were done in low Reynolds number jets using acoustic drivers 

as actuators. However, it was found that the acoustic driver actuators did not retain control authority as 

either the jet diameter or speed (thus Reynolds number) increased. Jet noise could not be readily studied 

because scale limitations did not allow high enough Mach number flows without Reynolds numbers 

above where the acoustic drivers lost effectiveness. In short, the actuators did not scale up adequately to 

be useful in either research or reduction concepts. 
NASA/TM—2006-214367 3 

Actuator Coupling 

To discuss further the difficulties faced by any actuator to be used on a jet and the motivation for the 

current work, details of how actuators couple with the flow to manipulate it must be discussed. The way 

to control the turbulent structures in a jet is through the flow instabilities which govern their formation 

and interactions. Two dominant instabilities have been identified to date, the initial shear layer instability 

and jet column instability. 

The initial shear layer instability.—The first type of instability is of an intrinsically two-dimensional 

shear layer and is applicable in regions of the jet near the nozzle where curvature is negligible compared 

with the shear layer momentum thickness the main scaling dimension. Although there is some sensitivity 

to the shape factor of the velocity profile, the maximum amplification of the instability has a Strouhal 

number, f00/Uj  0.013. This often corresponds to the first formation of vortices in the jet, which gives 

rise to subsequent vortex interactions. The decimation of vortices by interactions, while not as clean a 

process as simple vortex pairing, does result in spectral peaks which continue to scale with momentum 

thickness as the jet shear layer evolves down the potential core. Excitation of the shear layer mode has 

been successful in the past either by modulation of the mean jet flow or more efficiently by deflection of 

the initial separation point on the nozzle. Given that many small laboratory jets have momentum thickness 

that scales as DRe–1/2, St  0.013 corresponds to 40 kHz for a 1 in. jet at Mach 0.5. 

Jet column instability.—The second type of instability is that of the jet column, where the applicable 

scaling dimension is the diameter of the column. This instability scaling is clearly found at and beyond 

the end of the jet potential core and accounts for the dominant spectral peak near Strouhal number 

fD/UJ  0.3 found in this region. This instability supports several azimuthal modes, which after the first 

few become successively weaker in their amplification. The jet column instability has been successfully 

controlled in the past by directly modulating the axial flow in the potential core using acoustic methods. 

Perhaps it is obvious that the shear layer modes and jet column modes must compete near the end of 

the potential core where their regions of influence intersect. This interaction can be quite dramatic if they 

occur such that the shear layer mode has the right shape and frequency to directly reinforce the jet column 

mode. This is often the case in small, low speed jet facilities where much of the previous work on excited 

jets has taken place. It is less likely to occur in large jets, and partly explains why actuators for jets have 

not been as successful at higher Reynolds numbers. 

Energy and size considerations.—Because the shear layer instability scales with local momentum 

thickness, the scale-up of this type of excitation should be straightforward. Techniques that perturb the 

flow by oscillation of the initial shear layer separation need only be scaled by the perimeter of the nozzle 

and the exit boundary layer thickness of the nozzle. If we assume that the initial boundary layer thickness 

scales by the square root of scale factor and the perimeter by the scale factor, then the energy required 

goes up with size as scale factor raised to the power 1.5. Increasing the number of actuators to keep the 

energy density on the nozzle lip constant nearly accounts for the required scale-up of energy, with the rest 

perhaps accounted for by an increase in actuator strength, depending upon the actual boundary layer 

properties. Likewise, the frequencies required will be determined by the momentum thickness, not the 

scale factor directly. 

The jet column, being primarily controlled by global modes, is much harder to excite as the jet 

becomes larger. Here the impact of the actuator on the jet must not only scale with the perimeter, but 

perhaps with the area of the cross-section. To effect a modulation of the potential core requires energy 

scaling perhaps with the jet area, not the diameter. It is also much less likely that shear layer instability 

modes can couple with the jet column modes as they scale on boundary layer thickness, not nozzle 

diameter. 

Another consideration in applying control actuators to jets for noise reduction purposes is the noise of 

the actuators themselves. Many potential actuators produce more noise than they remove at laboratory 

scale. However, if one scales up the actuators linearly with the perimeter of the jet, then the actuator selfnoise increases linearly while the noise of the jet and the noise removed scales with area of the jet. 

Depending upon how the actuators scale up, actuator self-noise may or may not be a problem. 
NASA/TM—2006-214367 4 

Motivation of Present Work 

To summarize, it seems apparent from simple scaling arguments that, for instability excitation 

techniques to be applicable to jets of practical size, energetic actuators are critical. Thus, while previous 

application of the current technique, plasma arc excitation, has proven useful for research into jet 

dynamics in small scale laboratory environment, it is important to understand whether this technique can 

be brought to larger scale research facilities for higher fidelity tests, and eventually to full-scale engine 

applications. 

The presentation of this paper will focus on application of the plasma arc jet excitation to the 

classically understood jet instabilities, first in the small-scale jet rig at the Gas Dynamics and Turbulence 

Laboratory (GDTL) at The Ohio State University (OSU), and then where possible in the much bigger jet 

rig at NASA Glenn Research Center. The objectives are to demonstrate that the actuator concept works in 

the small-scale rig, delivering results at much greater Reynolds numbers than have been possible with 

previous techniques, and to demonstrate the degree to which the actuators work when taken to a much 

larger scale jet. Certain limitations, not the least of which was test time, precluded extensive comparisons 

in the big facility, but the initial results are encouraging that the plasma arc jet excitation technique may 

be able to control jets of larger Reynolds numbers, including that of full-scale aircraft. 

The major limitation was the number of actuators that could be operated simultaneously. The power 

supply/plasma generator was designed and built in-house at GDTL to operate 8 actuators simultaneously 

over a large frequency range (0 to 200 kHz) with independent control over phase and duty cycle of each 

actuator (Samimy et al. 2006a and b). This allowed forcing the small 1 in. diameter jet at GDTL with 

azimuthal modes m = 0, 1, 2, 3, ±1, ±2, and ±4. A decision was made to take this power supply to NASA 

and distribute 32 actuators around the 7.5 in. jet at NASA (a linear increase would require 60 actuators). 

With this arrangement only 8 actuators could operate simultaneously, thus limiting the forcing of the jet at 

NASA to only three azimuthal modes m = ±1, 4, and 8. Obviously this arrangement along with relatively 

short period of time allowed for the experiment imposed significant limitation on what could be achieved. 

The results presented below must be viewed with these limitations in mind. 

Nomenclature 

cj Speed of sound in jet

co Ambient speed of sound 

D Nozzle exit diameter 

F Frequency 

f0 Most amplified initial shear layer frequency 

fF Forcing (excitation) frequency 

fp Jet preferred mode frequency 

m Azimuthal mode number 

M Gas dynamics Mach number, Uj/cj

Ma Acoustic Mach number, Uj/co 

ReD Reynolds number, Uj D/

StD Strouhal number based on the nozzle exit diameter, fD/Uj

StDF Strouhal number based on the excitation frequency, fFD/Uj

StStrouhal number based on boundary layer momentum thickness at the nozzle exit, f00/Uj

Uj Nozzle exit velocity (jet centerline velocity within the potential core) 

x Distance from the nozzle exit along the jet centerline (fig. 1) 

y Vertical distance from the jet centerline (fig. 1) 

Angle with respect to downstream jet axis (x-axis) 

0 Momentum thickness at the nozzle exit or at the trailing edge of the splitter plate 
NASA/TM—2006-214367 5 

x

z

Nozzle Jet plume 

CCD camera 

Streamwise laser sheet 

Figure 1.—Schematic of the jet and the PIV set up at GDTL. Y-coordinate is normal to the plane. 

Experimental Facility and Techniques

Flow Facilities 

Two sets of experiments were conducted: one set in the smaller facility of GDTL to obtain detailed 

results on the application of plasma actuators for flow and acoustic control and the second set in the larger 

NASA Glenn Research Center facility to explore the scalability of the technique. A brief description of 

each facility will be given here. More detailed descriptions of GDTL and NASA facilities can be found in 

Samimy et al. (2006b) and in Bridges and Wernet (2004). Also, additional and more detailed results on 

flow and noise control using plasma actuators at GDTL can be found in Samimy et al. (2006a and b) and 

Kastner et al. (2006). 

At GDTL, the ambient air is compressed, dried, and stored in two cylindrical tanks at a pressure of up 

to 16 MPa with a capacity of 36 m3

. The compressed air is supplied to the stagnation chamber of the jet 

facility and conditioned before entering into a nozzle. Either a converging or a converging-diverging 

nozzle could be used and the jet could be operated continuously from low subsonic to Mach 2.5. An 

axisymmetric converging nozzle was used for the experiments discussed in this paper and the acoustic 

Mach number was varied in the high subsonic range (Ma from 0.5 to 0.84). The air is discharged 

horizontally through the nozzle into an anechoic chamber (fig. 1). The exit diameter of the nozzle is 

2.54 cm (1.0 in.). A nozzle extension, made of boron nitride, was attached to the exit of the nozzle to 

house the plasma actuators. 

Experiments were conducted on the dual flow jet rig in the Nozzle Acoustic Test Rig (NATR) at the 

NASA Glenn Research Center’s Aeroacoustic Propulsion Laboratory. This dual flow engine simulator 

provides exhaust gases at pressure and temperature conditions typical of modern turbofan engines with 

mass flows up to 65 kg/sec. The NATR is a 1.35 m free jet in which the jet rig sits, providing a moving 

free stream around the jet plume at speeds up to M = 0.35. For this set of experiments, a nozzle system 

typical of internally mixed engine exhaust systems with a confluent axisymmetric splitter was used and 

both streams were run at the same pressure ratio without heating the air. Also, no forward flight stream 

was employed, creating a basic single-stream static jet to compare with the GDTL facility. The final 

section of the nozzle was replaced with a boron nitride piece having the same internal flow lines as the 

high-fidelity nozzle, but with a much fatter lip to house the plasma actuator electrodes in a fashion similar 

to the GDTL hardware. A photograph of the plasma actuators mounted on at the NASA Nozzle Acoustic 

Test Rig is shown in figure 2. 
NASA/TM—2006-214367 6 

Figure 2.—Photograph of the plasma actuators mounted on at the NASA Nozzle

 Acoustic Test Rig. 

TABLE I.—NOZZLE EXIT DIAMETER AND JET FLOW 

VARIABLES FOR GDTL AND NASA FACILITIES 

D, 

cm 

Ma M ReD Uj, 

m/s 

fF, 

kHz 

(StDF = 0.3) 

fF, 

kHz 

(StDF = 1.8) 

GDTL 2.54 0.5 0.51 3.8105 166 2.0 11.8 

 .7 .74 5.7105 235 2.8 16.6 

 .84 .9 7.2105 285 3.4 20.2 

NASA 19.05 .5 .51 2.9106 166 0.3 1.6 

 .7 .74 4.3106 235 .4 2.2 

 .84 .9 5.5106 285 .4 2.7 

 .9 .99 6.4106 301 .5 2.8 

Table I depicts the nozzle exit diameter, and the jet acoustic and gas dynamic Mach numbers, 

velocity, and Reynolds number for the jets at GDTL and NASA used for the work presented in this paper. 

The boundary layer thickness at the exit of the nozzle at GDTL is very thin, making it almost impossible 

to obtain a boundary layer profile to determine its momentum thickness and its state. Kastner et al. (2004) 

used a similar converging nozzle and measured velocity at a few points within the boundary layer of an 

M = 0.9 jet. They estimated the boundary layer to be turbulent with a thickness of about 1 mm and a 

momentum thickness of approximately 0.1 mm. The characteristics of the boundary layer in the current 

experiments are expected to be quite similar at similar Mach numbers. At the lower Mach numbers, 

especially at Ma = 0.5, the boundary is probably laminar. Also shown in the table are the forcing 

frequencies at GDTL and NASA, which were adjusted to keep the same forcing Strouhal number (StDF). 

For example, for the Ma = 0.5 jet, the forcing frequency for forcing Strouhal number StDF = 0.3 was 2 kHz 

at GDTL and 300 Hz at NASA. 
NASA/TM—2006-214367 7 

An important point about scale up on actuator concepts becomes clear when considering the last 

column of Table I. Taking the same actuator design (and indeed actual hardware) to a much larger jet 

requires great flexibility as the frequencies change dramatically when Strouhal numbers are kept constant. 

The boundary layer of the NASA rig was not measured, but is sure to be fully turbulent given the 

relatively small contraction of the rig and nozzle. 

Flow and Acoustic Diagnostic Techniques 

At both GDTL and NASA, time-resolved pressure was measured using pressure transducers within 

the jet to explore the growth and decay of the imparted perturbations by the actuators and the ensuing 

instability waves/large-scale structures. Single-point static pressure and multi-point stagnation pressure, 

respectively, were measured at GDTL and NASA. Far field acoustic measurements were also carried out 

both at GDTL and NASA. In addition, PIV measurements were carried out at GDTL. Further details on 

these measurements are given below. 

At GDTL, the axis of the Kulite (Kulite Semiconductor Products, Inc.) pressure transducer probe was 

normal to the jet axis so that time-resolved static pressure—rather than total pressure—was measured. 

The probe tip was located at the lip-line of the nozzle radially, at the center of actuator 7 (fig. 3) 

azimuthally, and traversed in the streamwise direction manually from x/D = 0.5 to 7.5. The signal from 

the probe was amplified, band-pass filtered, and stored on a PC. An average spectrum was obtained from 

one hundred short-time spectra for each case and the perturbation level at the forcing frequency was 

calculated from the average spectrum. For the baseline case, the level was obtained at the frequencies 

matching the forcing frequencies. The perturbation level was normalized by 210–5 Pa in all cases. 

Multi-point time resolved pressure measurements were recorded at NASA using an array of Endevco 

pressure transducers. The array used 16 pressure transducers (8 50 psia and 8 50 psig transducers 

alternating in the azimuthal direction) mounted on a ring such that the each transducer was located at the 

nozzle lip-line and spaced 22.5° azimuthally from the previous transducer. The array was mounted on a 

traverse and moved in the streamwise direction from x/D = 0.4 to 6.0, where the flow threatened to 

Ceramic

capacitor

8 channels

HV switch

PCI

PC

8 channel DAC

Figure 3.—Schematic of the in-house fabricated plasma generator. 

8

7

6

5

4

3

2

1

DC power

supply

1A, 10kV

Ceramic nozzle extension

with 8 sets of pins

High power

resistor #2

High power

resistor #1
NASA/TM—2006-214367 8 

damage the array. Endevco amplifiers removed the DC signal and provided initial signal amplification. 

Precision amplifiers were used for further signal amplification. A 16-bit DataMax recording system, built 

by R.C. Electronics, digitized the data at 200 kHz (using a 190 kHz anti-aliasing filter). Ten seconds of 

continuous data were recorded for each jet condition allowing approximately 485 spectral averages based 

on a sample size of 8192 points and 50 percent sample overlap. 

At GDTL, far field noise was measured by using two 0.25 in. Bruel & Kjaer microphones, located 83 

and 45 nozzle exit diameters from the nozzle exit at 30° and 90° relative to the downstream jet axis, 

respectively. The far field acoustic results were normalized to a radius of 80 D. The acoustic signal from 

each microphone was conditioned and amplified by a four-channel Bruel & Kjaer Nexus conditioning 

amplifier. The signal was sampled at 200 kHz per channel by an NI A/D board and recorded on the hard 

disk. As in the Kulite data processing, an average spectrum was obtained from one hundred short-time 

spectra. With a sample size of 8192 points, the frequency resolution is 24.4 Hz. 

 At NASA, far-field noise was measured using an overhead array of twenty-four 1/4-in. Bruel & 

Kjaer microphones, located on an arc 72 nozzle diameters away. The arc spanned polar angles from 135° 

to 20° relative to the downstream jet axis. The microphone signals were conditioned and amplified by six 

units of four-channel Bruel & Kjaer Nexus conditioning amplifiers before being digitized at 190 kHz by a 

dedicated 16-bit DataMax recording system built by RC Electronics. Due to electronic noise produced by 

the plasma actuator delivered to the data recorder on signal lines intended to record the actuator timing, 

only the first 10 ms of data was reliable in most cases, limiting the frequency resolution of the spectra 

available for comparison with GDTL. The uncertainty of the spectral data is also increased to ±0.7dB. 

At GDTL, particle imaging velocimetry (PIV) was used to measure the x and y components of 

velocity on the x-y plane (fig. 1). Images were acquired and processed using a LaVision PIV system with 

a 2000 by 2000 pixel Redlake CCD camera employing a 75 to 300 mm Vivitar zoom lens with a 532 nm 

narrow band optical filter. Other essential hardware and software are housed in a dedicated computer with 

dual Intel Xeon processors. The system triggers a dual-head Spectra Physics PIV-400 Nd:YAG laser 

operating at the 2nd harmonic (532 nm). Image pairs were acquired at a sampling rate of approximately 

5 Hz. The jet flow was seeded with DEHS fluid introduced upstream of the stagnation chamber by a fourjet atomizer. This location was chosen to provide homogenous dispersion of the particle seed throughout 

the jet. The ambient was seeded using a fog generator. The fog was injected into a 15-in.-diameter 

co-annular jet and a small amount of air flow was introduced to generate a very low speed co-flow. The 

camera views the streamwise laser sheet orthogonally over about 9 jet diameters. The time separation 

between laser pulses was 2.0 μs for images divided into 32 by 32 pixel interrogation windows. Subregions for each image pair were cross-correlated using multi-pass processing with a 50 percent overlap 

in order to improve spatial resolution and prevent the appearance of spurious vectors by adaptively 

improving the window size. Initial passes used 64 by 64 pixel interrogation windows which were then 

used as a reference for the 32 by 32 pixel windows in the final pass. This experimental setup produced a 

velocity vector grid of 115 by 65 over the measurement domain. This translated to each velocity vector 

being separated by approximately 2 mm. 

Plasma Actuators and Plasma Generator System 

Each actuator consists of a pair of pin electrodes. The electrodes are distributed around the nozzle 

perimeter, approximately 1 mm upstream of the nozzle exit plane. A ring groove of 0.5 mm deep and 

1 mm wide was used to house the electrodes and to shield and stabilize the plasma. In our earlier 

experiments, the plasma was swept downstream by the flow without such a groove. We have used various 

nozzle extensions attached to the nozzle exit to house the electrodes, as well as various types and sizes of 

electrodes (Samimy et al. 2004 and 2005). For the work at GDTL and NASA presented here, the nozzle 

extension was made of boron nitride and steel wires or tungsten wires of 1 mm diameter were used for 

electrodes. Tungsten wires have proven to be more resistant to the erosion caused by the arc discharge. 

Measured center-to-center, the spacing between a pair of electrodes for each actuator is 3 mm, and the 
NASA/TM—2006-214367 9 

8 actuators at GDTL and 32 actuators at NASA were uniformly distributed around the nozzle exit. 

Figure 3 shows a schematic of the actuator arrangement used at GDTL. 

Figure 3 also shows a schematic of the multi-channel high-voltage plasma generator, designed and 

built in-house, used at GDTL. The plasma generator enables simultaneous powering of up to eight 

localized arc filament plasma actuators distributed around the perimeter of the ceramic nozzle extension, 

with independent frequency, duty cycle, and phase control of individual actuators. Each actuator is 

connected in series with a fast response, high repetition rate, high-voltage MOSFET switch (Behlke 

Electronic GmbH), two approximately 15 k high power solid body ceramic ballast resistors, and a highvoltage, high-current (10 kV, 1A) DC power supply (Glassman High Voltage, Inc.). Two of these power 

supplies are used to energize eight actuators. If all eight actuators are powered at the same time, the single 

actuator current is limited to 0.25A. The switches are controlled by using an 8-channel digital-to-analog 

output PCI card and the LabView software, which allows their independent frequency, duty cycle, and 

phase control. The switches are capable of producing high voltage pulses (up to 10 kV) at repetition rates 

from 0 to 200 kHz, with a very short pulse rise/fall time (~0.1 s ) and a variable duty cycle (from 0 to 

100 percent). Every switch is liquid cooled to allow continuous operation at high frequency, voltage and 

current. 

By turning the electronic switch on and off, positive high voltage pulses can be applied to the 

corresponding actuator. The high initial voltage is needed to produce breakdown in the approximately 

atmospheric pressure air in the gap between the two electrodes of an actuator, which are 3 mm apart in the 

present work. After the breakdown, the arc is generated and the voltage across the gap rapidly falls to a 

few hundred volts. The plasma generator is compact, robust, and simple to operate. In the present work, 

continuous operation of all eight actuators in an M = 0.9 flow (up to several minutes) has been achieved.

To reduce the EMI interference with the computer and data acquisition board, an in-house built optical 

isolation circuit was implemented in the low voltage side. 

In a typical acoustic driver, which has been used extensively in the literature for flow control, the 

input signal is a sine wave of prescribed frequency, peak amplitude and phase. The input amplitude 

changes gradually between the given positive and negative peaks in a cycle. On the other hand, the input 

signal to a plasma actuator is a rectangular wave with a variable duty cycle (for more details, see Kastner 

et al., 2006). In a plasma actuator, the input or forcing amplitude cannot easily be altered, as in an 

acoustic driver. Although the imparted forcing energy to the flow in a cycle can be adjusted by adjusting 

duty cycle, the effective forcing amplitude is not directly related to the total energy in a cycle. However, it 

is straightforward to force the jet in any simple azimuthal mode (m = 0, 1, 2, 3, ..., N; N depending on the 

number of actuators used) since there is no forcing amplitude variation. For combined/mixed azimuthal 

modes (m = ±1, ±2, ...), the input amplitude is the modulated amplitude of the positive and negative 

waves. These modes were simulated or mimicked by grouping the actuators without any amplitude 

modulation, which is not possible with the plasma actuators. For example, the top three actuators (1, 2, 

and 8) and the bottom three actuators (4 to 6) in figure 3 were grouped together and operated 180° out of 

phase to simulate m = ±1 mode. For this mode, actuators 3 and 7 were not operated to make these 

positions nodal points of m = ±1 mode. For other mixed mode, a similar grouping was done. The focus of 

this paper is on forcing jet instabilities using these actuators. There are two additional secondary effects. 

First, the actuator input signal has a prescribed frequency, but since it is a rectangular wave, it contains 

higher frequencies as well. In addition, each actuator would perhaps generate a pair of very weak 

streamwise vortices. While we plan to explore the second issue in the near future, any potential effects of 

the first subject, perceived to be small, is difficult. 

At NASA, the nozzle exit diameter was 19.05 cm (7.5 in.), which is 7.5 times the jet diameter at 

GDTL. To partially account for the increased diameter, we used 32 actuators uniformly distributed around 

the nozzle. A simple linear scaling-up would have required 60 actuators. But the available power supply 

could only power 8 actuators at a time. Therefore a decision was made to use 32 actuators equally 

distributed around the nozzle, enabling us to force the jet only with azimuthal modes m = ±1, 4, and 8. 

For this arrangement, additional 48 HV resistors were added to the generator and the wiring was 
NASA/TM—2006-214367 10 

significantly modified to run larger number of actuators so that each transistor was able to switch 4 

actuators at the same time. 

Experimental Results

The focus of the experimental results will be on forcing over a large frequency range high subsonic 

jets (Ma = 0.5 to 0.84). The forcing will be mostly with azimuthal mode m = ±1 with limited results from 

m = 4 and 8 forcing. 

Plasma Actuator Input Characterization 

Figure 4 shows time-dependent voltage, current, and power in the arc discharge between two steel 

wire electrodes of a plasma actuator taken at GDTL. During these measurements, only one actuator was 

operated at 20 kHz frequency and a 20 percent duty cycle in the ideally expanded M = 1.3 jet. Results in 

M = 0.9 jet are similar. The static pressure at the actuators location (1 mm before the nozzle extension 

exit) was approximately 1 atm. The top pulse train is the input signal to the high-voltage switch. The 

second graph from the top shows time-dependent voltage between the actuator electrodes, measured at 

point 1 (see fig. 3) by a Tektronix high voltage probe P6015A. A very large voltage overshoot (about 

4.0 kV) at the onset of the cycle (t ~12 μs on the graph) corresponds to electrical breakdown in the air 

0 100 150 250 50 200 300

Power, W Pulse, V

0

2

4

6

0 100 50

Time, μs

150 250 200 300

0 100 150 250 50 200 300

Curremt, mA 200

0

400

50

0

100

150

Voltage, kV

0 100 150 250 50 200 300

0

2

4

Figure 4.—Time-dependent voltage, current, and power in a plasma actuator operated at 20 kHz frequency

 and 20 percent duty cycle (only one actuator was operated). The top pulse train is the input signal to the

 high-voltage switch. 
NASA/TM—2006-214367 11 

between the electrodes, followed by a sharp voltage drop to about 400 V, as a stationary arc discharge is 

formed. After the discharge current is interrupted by rapidly closing the switch (10 μs later), the plasma is 

quickly blown off by the flow. As a result, the high voltage electrode (anode) becomes floating and the 

voltage remains relatively low (at about 600 to 700 V) until the moment the switch is opened again 50 μs 

later. 

The actuator current was measured simultaneously with the voltage measurement using a LeCroy 

CP031 current probe and is shown in figure 4. The current trace (third from top) is nearly a step function, 

with the steady state current of about 250 mA when the switch is open and no current when it is closed. 

The time-dependent arc discharge power can be obtained simply by multiplying the current and voltage 

traces. The result is shown in the bottom graph of figure 3. The power during the current pulse reaches 

approximately 100 W, which corresponds to a time-averaged actuator power of only 20 W (i.e., 160 W 

net power for all eight actuators in operation). For comparison, the flow power (the total enthalpy flux) of 

the jet at these conditions is about 28 kW. These results demonstrate that high-speed flow control by 

localized arc plasma actuators can be highly energy efficient. 

In an effort to understand the mechanism behind such efficiency we have conducted a series of 

plasma temperature measurements. The temperature was measured by monitoring the emission of second 

positive band system of nitrogen from the arc region. The emitted light was directed through an optical 

fiber to a Princeton Instrument Optical Multi-channel Analyzer, which was equipped with an ICCD 

camera. Least squares fit was used to compare the experimental spectra of nitrogen second positive 

system and the synthetic spectra generated by a theoretical code with temperature being a variable 

parameter. The results show that the gas temperature rises extremely fast after the breakdown and reaches 

1700 °C in several microseconds. This rapid heating might be the key for the actuator’s dramatic 

influence on the flow that has been demonstrated in our experiments. 

The measurements also showed that the voltage, current, and power traces for all eight actuators are 

essentially the same. Depending on the frequency, mode, and duty cycle the arc discharges produced by 

different actuators may overlap in time. The same is true about the 32 actuators that we ran at NASA 

facility. The traces look very similar to those in figure 4, except that the duty cycles were kept very low 

(usually 1 to 3 percent) because of much of much lower forcing frequencies at NASA’s larger facility. 

Growth and Decay of Perturbations and the Ensuing Instability Waves 

The perturbation caused by the plasma actuators in the jet shear layer was measured both in the OSU 

and NASA facilities. The measured pressure is static at GDTL, while it is total at NASA. At GDTL, 

detailed analysis has been done in the Ma = 0.84 jet in order to explore the growth and decay of the 

perturbation seeded into the flow by plasma actuators and the ensuing instability waves. The pressure 

transducer was located at the same azimuthal location as actuator number 7 in figure 3, was traversed in 

the streamwise direction from x/D = 0.5 to 7.5. The tip of the transducer grazes the shear layer at the first 

measurement location (x/D=0.5), but is well into the shear layer in farther downstream locations, as the 

shear layer is growing. The amplitude of the local hydrodynamic pressure at the forcing frequency was 

obtained using power spectrum of the time-resolved pressure signal. 

The spatial development of the perturbation level for the GDTL jet is shown in figures 5 and 6 for the 

flapping and non-flapping plane, respectively. At all StDF, the maximum perturbation level is about the 

same on the flapping plane. The growth rate of the perturbation is slower at a lower forcing frequency and 

the saturation occurs at a farther downstream location when it is compared to a higher frequency forcing 

case (fig. 5). When the jet was forced at StDF = 0.36, which is in the jet preferred frequen 

cy range, saturation occurred at x/D = 2.5, but the fluctuation level remained relatively high to a 

farther downstream location. This growth and decay is consistent with the flow visualization in an ideally 

expanded Mach 1.3 jet (Samimy et al. 2005), where the large-scale structures grew in size until they reach 

the end of potential core. As the forcing frequency increases, the saturation position is getting closer to 

the jet exit and the perturbation level decays very rapidly after saturation. It seems that the maximum 

perturbation level was reached upstream of the first measuring location of x/D = 0.5 for StDF = 1.81. 
NASA/TM—2006-214367 12 

0 23 5 1 4 6

StD = 0.36

BL = (0.36)

StD = 0.91

BL = (0.91)

StD = 1.81

BL = (1.81)

7 8

Amplitude, dB

x/D

110

120

130

140

80

70

90

100

Figure 5.—Spatial development of seeded perturbations on the flapping plane at several

 Strouhal numbers (StDF) for azimuthal mode m = ±1.

0246

StD = 0.36

BL = (0.36)

StD = 0.91

BL = (0.91)

StD = 1.81

BL = (1.81)

8

Amplitude, dB

x/D

110

120

130

140

80

70

90

100

Figure 6.—Spatial development of seeded perturbations on the non-flapping plane at

 several Strouhal numbers (StDF) for azimuthal mode m = ±1. 
NASA/TM—2006-214367 13 

StDF = 0.3

BL = (0.3)

StDF = 1.8

BL = (1.8)

StDF = 0.3

BL = (0.3)

StDF = 1.8

BL = (1.8)

0 2

(a)

1 3456

Amplitude, dB

x/Dj

5 dB

Amplitude, dB

Figure 7.—Axial growth of input perturbations in the flapping (a) and non-flapping (b) plane at a Ma = 0.5 jet

 condition perturbed in the azimuthal ±1 mode. 

5 dB

0 2 1 34 56

x/Dj

(b)

When the jet is excited at a higher frequency, the flow structures are smaller and they decay quickly 

as seen in Samimy et al. (2005). This explains why the perturbation level decays rapidly at higher forcing 

frequencies. 

On the non-flapping plane, the maximum fluctuation level is 10 to 25 dB lower than that on the 

flapping plane. This lower fluctuation level was expected since the probe was on the nodal point of the 

flapping motion. Although the perturbation level was reduced, the amplification trends appear to be 

similar to those observed on the flapping plane. 

The axial growth of the input perturbations measured at NASA is shown in figure 7 for the flapping 

and non-flapping planes, at a Ma = 0.5 jet condition. Only perturbation frequencies of StDF = 0.3 and 

StDF = 1.8 were considered during this phase of NASA testing due to time constraints. In the flapping 

plane, at a perturbation frequency StDF = 0.3, near the jet preferred frequency, the perturbed jet shows a 

significant amplitude increase compared to the baseline jet. The saturation point for the perturbed jet 

appears between x/D = 2 and x/D = 3 and is followed by a slow decay. In the non-flapping plane, the 

perturbed jet again saturates at a higher amplitude than the baseline jet and at x/D = 40, slightly 

downstream of the jet observed in the flapping plane. 

The Ma = 0.5 jet perturbed at StDF = 1.8 shows limited response in the flapping plane close to the 

nozzle exit (less than x/D = 2). No significant response is observed further downstream in the flapping 

plane or at any axial location in the non-flapping plane. The NASA jet appears to be fairly insensitive to 

this forcing frequency. 

The spatial development of the input perturbations were also recorded at NASA using a Ma = 0.84 jet 

condition (fig. 8) to compare against the detailed data recorded at GDTL. Results in the flapping plane 

show the jet is responding to the perturbations based on comparison with the baseline. Saturation of the 

excited mode occurs around x/D = 5, where the excited jet is 3 dB above the baseline. Data in the nonflapping plane also shows increased amplitude in the jet excited at StDF = 0.3 versus the baseline, but, as 

expected, the difference is smaller. The saturation peak in the non-flapping plane appears much broader, 

occurring somewhere x/D = 3 and x/D = 5. In both the flapping and non-flapping planes, the amplitude 

difference between the excited and unexcited jet is smaller and the saturation point occurs further 

downstream in NASA data than was observed in the GDTL data. No significant response is observed 

when the jet is perturbed at StDF = 1.8, again showing that the NASA jet, unlike the GDTL jet, is not 

particularly sensitive to this perturbation frequency. 
NASA/TM—2006-214367 14 

Figure 8.—Axial growth of input perturbations in the flapping (a) and non-flapping (b) plane at a Ma = 0.9 jet

 condition perturbed in the azimuthal ±1 mode. 

StDF = 0.3

BL = (0.3)

StDF = 1.8

BL = (1.8)

StDF = 0.3

BL = (0.3)

StDF = 1.8

BL = (1.8)

0 2

(a)

1 3456

Amplitude, dB

x/Dj

5 dB

Amplitude, dB

5 dB

0 2 1 34 56

x/Dj

(b)

0246

Baseline

m = 0.5

Baseline

m = 0.7

Baseline

m = 0.9

8 10

Mach number

x/D

0.4

0.6

0.8

1.0

0.0

0.2

Figure 9.—Centerline Mach number for various jet Mach numbers with forcing Strouhal

 number of ~0.3 and azimuthal mode m = ±1. 

Flow Results 

 PIV measurements on a streamwise plane passing through the centerline of the jet (fig. 1) were 

carried out at GDTL at three Ma = 0.5, 0.7, and 0.84 (M = 0.51, 0.74, and 0.9). The measured centerline 

Mach number profiles are shown in figure 9 for the baseline cases as well as forced cases with a Strouhal 

number of approximately 0.3 (the actual forcing Strouhal numbers are 0.31, 0.33, and 0.27 for the 

acoustic Mach numbers 0.5, 0.7, and 0.84, respectively). These Strouhal numbers are close to the jet 
NASA/TM—2006-214367 15 

0246

Baseline

m = 0.5

Baseline

m = 0.7

Baseline

m = 0.9

8 10

Normalized TKE

x/D

0.04

0.05

0.06

0.07

0.00

0.02

0.01

0.03

Figure 10.—Centerline turbulent kinetic energy for various jet Mach numbers with forcing

 Strouhal number of ~0.3 and azimuthal mode m = ±1. 

preferred Strouhal numbers, based on the jet centerline Mach number measurements. In all cases, the 

forcing amplitude remained the same, as there is no control on this amplitude. In the baseline cases, as the 

jet Mach number increases, the potential core becomes longer: Lp/D = 5.0, 6.1, and 6.5, respectively, for 

the jet Ma = 0.5, 0.7, and 0.84. The potential core lengths were shortened to approximately Lp/D = 3.0, 

3.3, and 3.5 by the forcing for the jet Ma = 0.5, 0.7, and 0.84, respectively. While the centerline Mach 

number decay rate downstream of the jet potential core is significantly increased for the forced cases, in 

comparison with the baseline cases, they are about the same for all three forcing cases. The centerline 

Mach number for the forced cases is about 40 percent of the corresponding baseline case at x/D = 9.0.

The centerline turbulent kinetic energy variations for all three Mach numbers are shown in figure 10. 

While the baseline Mach 0.7 and 0.84 cases show similar centerline kinetic energy distribution, the values 

for the lowest Mach number are much higher. The cause for the difference is not clear at this time, but it 

could be Reynolds number effect—the boundary layer is perhaps laminar at the nozzle exit for Mach 0.5. 

For the controlled cases, although there are some differences for different Mach numbers, the overall 

effects of forcing at m = ±1 appears to be similar for all the jet Mach numbers. For all the cases, the 

centerline TKE increases by several folds over a significant portion of the jet. 

For the jet Mach number of 0.84 and the forcing azimuthal mode of m = ±1, the forcing Strouhal 

number was varied from 0.1 to 3.0 to investigate the effects of forcing frequency. As shown in figure 11, 

the jet core shortens and the centerline Mach number decays much faster as the forcing Strouhal number 

approaches StDF = 0.27 from either lower or higher side. This trend with forcing Strouhal number is 

similar to that for an ideally expanded Mach 1.3 jet (Samimy et al. 2006a). At this preferred (or close to 

preferred) Strouhal number, the centerline turbulent kinetic energy grows faster than at any other forcing 

Strouhal number. The centerline turbulent kinetic energy trend, shown in figure 12, is similar to that of 

the centerline Mach number. Forcing at the preferred mode, TKE level increased by an order of 

magnitude in comparison with the baseline case and with those forcing at much higher or lower Strouhal 

numbers. It was shown Samimy et al. (2006a), that large and specially and temporally coherent structures 

were generated by the forcing a Mach 1.3 ideally expanded jet with m = ±1 at the preferred mode. 
NASA/TM—2006-214367 16 

0246

Baseline

StD = 0.09

StD = 0.27

StD = 0.54

StD = 0.91

StD = 1.8

8 10

Mach number

x/D

0.6

0.7

0.8

0.9

0.4

0.5

Figure 11.—Centerline Mach number for the jet acoustic Mach number 0.84 with forcing

 azimuthal mode m = ±1 and various forcing Strouhal number. 

0246

Baseline

StD = 0.09

StD = 0.27

StD = 0.54

StD = 0.91

StD = 1.8

8 10

Mach number

x/D

0.03

0.04

0.05

0.06

0.00

0.02

0.01

Figure 12.—Centerline turbulent kinetic energy forthe jet acoustic Mach number 0.84 with

 forcing azimuthal mode m = ±1 and various forcing Strouhal number. 
NASA/TM—2006-214367 17 

Although it is not possible to visualize the large-scale structures at these subsonic jets by the same 

techniques, it is expected that similar large-scale structures exist in these jets as well. Obviously, the 

periodic large-scale-structures are responsible for the entrainment mixing and thus significant increase in 

turbulent kinetic energy. 

Far Field Acoustic Results 

For the preliminary experiments at NASA we had two major constraints. First, we had to reduce the 

actuator number density by a factor of two (8 at GDTL for 1 in. nozzle and 32 at NASA for 7.5 in. 

nozzle). Second, we could drive only 8 of the actuators at a time. These constraints limited the azimuthal 

mode forcing to only m = ±1, 4, and 8. On the other hand the highest azimuthal mode we could run at 

GDTL was m = 3. Therefore, we will present far field acoustic from both GDTL and NASA while forcing 

at various frequencies with m = ±1 and also some limited forcing cases at m = 4 and 8 from NASA. 

Figures 13 and 14 show power spectra at several forcing Strouhal numbers for far field acoustic at 90° 

and 30° with respect to the downstream jet axis for acoustic Mach number 0.5 jets at GDTL and NASA, 

respectively. The azimuthal mode is ±1 and the microphones are located on the flapping plane of the jet 

for both cases. It appears that the small jet at GDTL at this Mach number is overwhelmed by forcing, the 

actuation tone and its harmonics appear in the spectrum, and noise is increased across the broadband at 

both angles in most of the forcing cases. The amplification is strongest when forcing close to the jet 

preferred mode frequency. Only in the two highest Strouhal number forcing cases (StDF = 1.7 and 5.0) is 

the broadband noise slightly reduced over a portion of the spectrum. There are some similarities and some 

differences between the results at GDTL and NASA. The main differences are that the jet at NASA is not 

overwhelmed by the actuation, and the forcing tones and their harmonics are not very visible until much 

higher forcing frequencies. The similarities are noise amplification in lower forcing frequencies, 

maximum amplification at StD ~ 0.3 and slight noise reduction at much higher forcing frequencies. The 

results from the microphones on the non-flapping plane show similar trends. 

10 dB 10 dB

10–1 100

StD

10–1 100

StD

Ma = 0.5, m = ±1 FP, θ = 90° Ma = 0.5, m = ±1 FP, θ = 30°

Figure 13.—Far field noise spectra for Ma = 0.5 jet at GDTL for forcing at various Strouhal number with

 azimuthal mode m = ±1. 

Baseline

StDF = 0.3

StDF = 0.7

StDF = 1.1

StDF = 1.7

StDF = 5.0
NASA/TM—2006-214367 18 

10 dB 10 dB

10–1 100

StD

101 102 10–1 100 101 102

StD

Baseline

StDF = 0.3

StDF = 0.7

StDF = 1.1

StDF = 1.8

StDF = 5.0

Ma = 0.5, m = ±1 FP, θ = 90° Ma = 0.5, m = ±1 FP, θ = 30°

Figure 14.—Far field noise spectra for Ma = 0.5 jet at NASA for forcing at various Strouhal number with

 azimuthal mode m = ±1. 

10 dB 10 dB

10–1 100

StD

10–1 100

StD

Ma = 0.84, m = ±1 FP, θ = 90° Ma = 0.84, m = ±1 FP, θ = 30°

Figure 15.—Far field noise spectra for Ma = 0.84 jet at GDTL for forcing at various Strouhal number with

 azimuthal mode m = ±1. 

Baseline

StDF = 0.3

StDF = 0.7

StDF = 1.1

StDF = 1.8

StDF = 4.5

Figures 15 and 16 show far field noise power spectra at the same forcing Strouhal numbers, forcing 

azimuthal mode, observation angles and observation plane as those in figures 13 and 14, but for 

Ma = 0.84. For GDTL (fig. 15) the results are very similar to Ma = 0.5 in figure 13. For NASA, the noise 

is still increased when forcing at the lower StDF, but no longer is there any reduction when forcing the 

high StDF. As stated earlier, the number of actuators did not directly scale between the two facilities, thus 

the lack of actuator authority at higher Mach numbers probably points to the need for more actuators. 

Detailed far field acoustic results from the GDTL facility for jet Ma = 0.84 were presented in Samimy et 

al. (2006b). Figure 17 summarizes the results seen from that paper by presenting the change in the 

OASPL, over the StD range of 0.01 to 5.0, when exciting over a StDF range of 0.1 to 5 with various 

azimuthal modes. Several observations can be made which complement the results seen in figures 13 

and 15. 
NASA/TM—2006-214367 19 

0 2

StDF

1 12 3 3 4 4 5 5 0

StDF

Figure 17.—Changes in the far field sound pressure level between StDF = 0.01 and 5.0 at 90° and 30° over a

 large forcing Strouhal number (StDF) range for different azimuthal modes.

Ma = 0.84, θ = 90° Ma = 0.84, θ = 30°

5

4

3

2

1

0

–1

–2

ΔOASPL, dB

5

4

3

2

1

0

–1

–2

ΔOASPL, dB

m = 0

m = 1

m = ±1 NFP

m = ± FP

m = 2

m = 3

m = ±4

m = 0

m = 1

m = ±1 NFP

m = ± FP

m = 2

m = 3

m = ±4

(1) There is a significant noise increase when the jet is excited around the preferred frequency of the 

jet column instability (StDF = 0.2 to 0.5). This is in agreement with the results in the literature using 

acoustic drivers (e.g., Jubelin, 1980; Ahuja et al., 1982; Lu, 1983). 

(2) Noise reduction of 0.6 to over 1.0 dB is achieved over a large range of excitation frequencies. This 

reduction seems to peak between StDF = 1.5 to 2.0 at 30° and around StDF = 3.0 to 3.5 at 90°. 

(3) There seems to be a clear advantage of excitation with azimuthal modes m = 3 and ±4. These were 

the highest azimuthal modes that could be excited with the 8 actuators used at GDTL. However at NASA, 

32 actuators were used and azimuthal modes m = 4 and 8 could be excited. 
NASA/TM—2006-214367 20 

10 dB 10 dB

10–1 100

StD

101 102 10–1 100 101 102

StD

Baseline

StDF = 0.3

StDF = 0.7

StDF = 1.1

StDF = 1.8

StDF = 5.0

Ma = 0.5, m = 4, θ = 90° Ma = 0.5, m = 4, θ = 30°

Figure 18.—Far field noise spectra for Ma = 0.5 jet at NASA for forcing at various Strouhal number with

 azimuthal mode m = +4. 

10 dB 10 dB

10–1 100

StD

101 102 10–1 100 101 102

StD

Baseline

StDF = 0.3

StDF = 0.7

StDF = 1.1

StDF = 1.8

StDF = 5.0

Ma = 0.5, m = 8, θ = 90° Ma = 0.5, m = 8, θ = 30°

Figure 19.—Far field noise spectra for Ma = 0.5 jet at NASA for forcing at various Strouhal number with

 azimuthal mode m = +8. 

Figure 18 and 19 shows spectra for forcing at NASA the m = 4 and m = 8, respectively. For m = 4 and 

at 30°, the noise is reduced with all forcing frequencies. At 90° the broadband level is similar to the 

baseline jet. For m = 8, the suppression of the broadband at 30° is similar to that seen when forcing m = 4. 

The ability to force higher azimuthal modes appears to be advantageous when looking for noise 

mitigation. Many authors have reported on the first few azimuthal modes containing most of the far-field 

energy with the higher modes being less efficient radiators. The results shown at both facilities seem to 

support this conclusion. The higher Mach number jet was also forced with azimuthal modes, m, 4 and 8 at 

NASA (not shown here), but similar to the m = ±1 results (see fig. 16), the lack of sufficient actuators 

seemed to inhibit the actuator authority. 
NASA/TM—2006-214367 21 

Concluding Remarks 

The idea of manipulating flow to change its characteristics is over a century old. Manipulating 

instabilities of a jet to increase its mixing and to reduce its radiated noise started in 1970s. While the 

effort has been successful in low-speed and low Reynolds number jets, available actuators’ capabilities in 

terms of their amplitude, bandwidth, and phasing have fallen short in control of high-speed and high 

Reynolds number jets of practical interest. Over the past couple of years, a class of high amplitude and 

high bandwidth actuators called localized arc filament plasma actuators (LAFPA) have been developed at 

OSU and extensively used at GDTL for control of high-speed and high Reynolds number jets. The results 

of the work has been presented and published in the literature. While the technique has been quite 

successful and is very promising, all the work up to this point had been carried out using small high 

subsonic and low supersonic jets from a 2.54 cm diameter nozzle exit with a Reynolds number of about a 

million. 

The preliminary work reported in this paper is a first attempt to evaluate the scalability of the 

technique. The power supply/plasma generator was designed and built in-house at GDTL to operate 8 

actuators simultaneously over a large frequency range (0 to 200 kHz), covering over the jet instabilities 

frequencies, with independent control over phase and duty cycle of each actuator. This allowed forcing 

the small jet at GDTL with azimuthal modes m = 0, 1, 2, 3, ±1, ±2, and ±4 over a wide range of 

frequencies. This power supply was literally taken to and used at NASA with relatively minor 

modifications. At NASA, a total of 32 actuators were distributed around the 7.5 in. nozzle at NATR. A 

linear increase with nozzle exit diameter would require 60 actuators at NATR. LAFPA design does not 

require any change with nozzle scale. However, it is expected that the number of actuators will linearly 

increase with nozzle exit diameter. With this arrangement at NATR only 8 actuators could operate 

simultaneously, thus limiting the forcing of the jet at NASA to only three azimuthal modes m = ±1, 4, 

and 8. 

Very preliminary results at NASA indicate that the trends observed in the larger NASA facility in 

terms of the effects of actuation frequency and azimuthal modes are similar to the smaller GDTL facility. 

This was most prominent at the lower Mach number (Ma = 0.5). However, the actuation authority seems 

to fall short in the larger jet at higher Mach numbers, resulting in decreased amplitude response compared 

to the small jet, which is attributed at this point to the lack of sufficient number of actuators. The 

preliminary results seem also to suggest that amplitude of actuation tones is similar in both the small and 

larger jets. 

References 

Ahuja, K.K., Lepicovsky, J., and Burrin, R.H., “Noise and Flow Structure of a Tone-Excited Jet,” AIAA 

Journal, Vol. 20, No. 12, 1982, pp. 1700–1706. 

Ahuja, K.K. and Blakney, D.F., “Tone Excited Jets, Acoustic Measurements,” Journal of Sound and Vibration, 

Vol. 102, 1985, pp. 93–117. 

Bridges, J. and Wernet, M.P., “Cross-Stream PIV Measurements of Jets with Internal Lobed Mixers,” AIAA 

Paper 2004–2896 (2004). 

Brown, G.L. and Roshko, A., “On Density Effects and Large Structure in Turbulent Mixing Layers,” Journal 

of Fluid Mechanics, Vol. 64, 1974, pp. 715–816. 

Crow, S. and Champagne, F., “Orderly Structure in Jet Turbulence,” Journal of Fluid Mechanics, Vol. 48, 

1971, pp. 547–591. 

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28

Flow and Noise Control in High Speed and High Reynolds Number Jets Using

Plasma Actuators

M. Samimy, J. Kastner, J.-H. Kim, Y. Utkin, I. Adamovich, and C.A. Brown

Jet flow; Aeroacoustics; Shear layer; Actuator; Plasma (physics); Kelvin-Helmholtz

instability; Active control

Unclassified -Unlimited

Subject Categories: 07, 09, and 34

Prepared for the Third Flow Control Conference sponsored by the American Institute of Aeronautics and Astronautics, San

Francisco, California, June 5–8, 2006. Responsible person, Clifford A. Brown, organization code RTA, 216–433–2896.

Localized arc filament plasma actuators have recently been developed and extensively used at Gas Dynamics and

Turbulence Laboratory (GDTL) for control of high-speed and high Reynolds number jets. While the technique has

been quite successful, all the work up to this point had been carried out using small high subsonic and low supersonic jets from a 2.54 cm diameter nozzle exit with a Reynolds number of about a million. The preliminary work

reported in this paper is a first attempt to evaluate the scalability of the technique. The localized arc filament plasma

actuators were installed and tested at the NASA Nozzle Acoustic Test Rig (NATR) using a 7.5 in. diameter nozzle.

Very preliminary results at NATR indicate that the trends observed in the larger NASA facility in terms of the effects

of actuation frequency and azimuthal modes are similar in both small GDTL and larger NASA jets. However, the

actuation authority seems to fall short in the larger jet at higher Mach numbers, which is attributed at this point to the

lack of sufficient number of actuators.