David E. Ashpis

Glenn Research Center, Cleveland, Ohio

Douglas R. Thurman

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

DBD Plasma Actuators for Flow Control in Air

Vehicles and Jet Engines—Simulation of Flight 

Conditions in Test Chambers by Density Matching

NASA/TM—2011-217006

March 2011

U.S. ARMY

RESEARCH LABORATORY
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David E. Ashpis

Glenn Research Center, Cleveland, Ohio

Douglas R. Thurman

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

DBD Plasma Actuators for Flow Control in Air

Vehicles and Jet Engines—Simulation of Flight 

Conditions in Test Chambers by Density Matching

NASA/TM—2011-217006

March 2011

U.S. ARMY

RESEARCH LABORATORY

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

This project was supported by the NASA Subsonic Fixed Wing Project of the Fundamental Aeronautics Program. We wish 

to thank Scott Jones and Michael Tong from NASA Glenn Research Center for their help in developing the engine models. 

Thanks also to Mikhail Shneider (Princeton University), Andrey Starikovskiy (Princeton University), and Alex Likhanskii 

(Tech-X Corporation) for useful discussions. Thanks for NASA Glenn Publishing department—Nancy Mieczkowski for the 

graphics, Lisa Greeney for layout, and Mary Eitel-Kim for coordinating production of this report.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

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

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.
NASA/TM—2011-217006 1 

DBD Plasma Actuators for Flow Control in Air Vehicles and 

Jet Engines—Simulation of Flight Conditions in 

Test Chambers by Density Matching 

David E. Ashpis 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Douglas R. Thurman 

U.S. Army Research Laboratory 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Dielectric Barrier Discharge (DBD) Plasma actuators for active flow control in aircraft and jet 

engines need to be tested in the laboratory to characterize their performance at flight operating conditions. 

DBD plasma actuators generate a wall-jet electronically by creating weakly ionized plasma, therefore 

their performance is affected by gas discharge properties, which, in turn, depend on the pressure and 

temperature at the actuator placement location. Characterization of actuators is initially performed in a 

laboratory chamber without external flow. The pressure and temperature at the actuator flight operation 

conditions need to be simultaneously set in the chamber. A simplified approach is desired. It is assumed 

that the plasma discharge depends only on the gas density, while other temperature effects are assumed to 

be negligible. Therefore, tests can be performed at room temperature with chamber pressure set to yield 

the same density as in operating flight conditions. 

The needed chamber pressures are shown for altitude flight of an air vehicle and for jet engines at 

sea-level takeoff and altitude cruise conditions. Atmospheric flight conditions are calculated from 

standard atmosphere with and without shock waves. The engine data was obtained from four generic 

engine models; 300-, 150-, and 50-passenger (PAX) aircraft engines, and a military jet-fighter engine. 

The static and total pressure, temperature, and density distributions along the engine were calculated for 

sea-level takeoff and for altitude cruise conditions. The corresponding chamber pressures needed to test 

the actuators were calculated. The results show that, to simulate engine component flows at in-flight 

conditions, plasma actuator should be tested over a wide range of pressures. For the four model engines 

the range is from 12.4 to 0.03 atm, depending on the placement of the actuator in the engine. For example, 

if a DBD plasma actuator is to be placed at the compressor exit of a 300 PAX engine, it has to be tested at 

12.4 atm for takeoff, and 6 atm for cruise conditions. If it is to be placed at the low-pressure turbine, it has 

to be tested at 0.5 and 0.2 atm, respectively. These results have implications for the feasibility and design 

of DBD plasma actuators for jet engine flow control applications. In addition, the distributions of unit 

Reynolds number, Mach number, and velocity along the engine are provided. The engine models are nonproprietary and this information can be used for evaluation of other types of actuators and for other 

purposes. 

Introduction 

There is strong interest in active flow control techniques for applications in air vehicles as well as in 

jet engines, for example, to eliminate flow separation, improve efficiency or reduce noise (Ref. 1). 

Dielectric Barrier Discharge (DBD) Plasma actuators were proposed for active flow control of various 

flows, and the technology has been an active research area in the last decade. 
NASA/TM—2011-217006 2 

The main active flow control technique in aerodynamics is based on injection of small jets in a steady 

or unsteady manner into the flow. The small input creates a large global effect that provides the desired 

flow improvement. DBD plasma actuators create a wall-jet by purely electronic means. The jet can be 

steady, or in unsteady mode by pulsing or modulating, and can be used for active flow control, similar to 

any pneumatically- or mechanically-generated jet. 

A DBD actuator is shown in Figure 1. The actuator consists of a pair of thin conducting electrodes 

separated by a dielectric. Usually there is one exposed electrode and one covered electrode and the 

electrodes are offset. Typically, at atmospheric conditions, a high voltage (1 to 40 kV RMS), high 

frequency (1 to 20 kHz), signal is applied to the electrodes, creating localized weakly ionized gas plasma 

discharge on the surface near the edge of the exposed electrode. A typical discharge is shown in Figure 2. 

The jet is generated in the plasma region via the electrohydrodynamic effect, a process of collisions 

between ions and neutral molecules in the plasma. The actuator construction is very simple, but the 

physical mechanisms involved are quite complex and include interactions among electrical fields, 

electrons, positively and negatively charged species, the electrodes, and the dielectric surface (Refs. 2 to 

4). Gas is drawn from the surroundings to form a thin wall jet that is roughly parallel to the surface and 

directed away from the exposed electrode edge in the direction of the covered electrode, as visualized by 

experiments and computation (Refs. 5 to 7). There is slight heating involved, but its effect on the jet is 

negligible. Other types of plasma-based flow control devices that generate localized intense heating 

(Ref. 8) are not included in the scope of devices addressed in this paper. More detailed information and 

references on DBD actuators and their applications for aerodynamic flow control can be found in several 

review articles (Refs. 9 to 12). 

The advantages of DBD actuators are that they are surface-mounted, fully electronic, low power, high 

frequency-band devices. There are no moving parts, tubes, ducts or surface holes. Flexible operation is 

possible by controlling the input voltage and waveforms. DBD plasma actuators are particularly attractive 

for gas turbine and turbomachinery applications: They are thin, surface mounted, and do not require 

internal volumes or passages. Their construction can be made suitable for high temperature environment 

by choosing high-temperature alloys for the electrodes and temperature-resistant ceramic materials for the 

dielectric. They can easily be integrated with futuristic engine components to be made of ceramics and 

composites. 

The majority of the research in DBD actuators area was focused on applications in external flows, 

particularly for wings and airframes, rather than on propulsion. But there have been important efforts 

directed to turbomachinery applications. There have been several successful experimental demonstrations 

of active flow control with DBD plasma actuators to eliminate low Reynolds number separation in LowPressure Turbine flows (Refs. 13 to 16), and to reduce effects of turbine tip leakage (Refs. 17 to 20). 

Those experiments were performed in wind tunnels or linear cascades at room temperature. 

DBD actuators need to be tested in actual flight conditions in order to be used as flow control devices 

in practical applications. Some flight tests were performed on air vehicles; on a small remotely controlled 

airship (Refs. 21 and 22) a full-size piloted glider (Ref. 23), and a small UAV (Ref. 24), but no tests have 

been performed in jet engines. The aerodynamic performance of the actuators needs to be characterized to 

prove that they have sufficient authority at the flow conditions at the location of their placement in the air 

vehicle or jet engine. In addition, the electrical performance, particularly the power consumption of the 

actuators, needs to be quantified, as it is needed for design of power supplies and for cost-benefit analysis 

of the flow control technology. This paper addresses the test conditions needed to characterize the 

actuators in the laboratory. 

The basic characterization in the laboratory is performed without external flow. The aerodynamic 

performance is characterized by measuring the velocity profile of the wall jet and/or the thrust generated 

by the actuator. The electrical performance is characterized by measuring the current, voltage, and power. 

Most of the tests to date have been performed at room temperature and atmospheric conditions, but in 

order to simulate flight conditions the actuator has to be placed in a chamber with controlled temperature 

and pressure representative of the flow conditions in flight. A small number of tests are reported in the 

literature in chambers at room temperature and sub-atmospheric conditions, with some conflicting results 
NASA/TM—2011-217006 3 

(Refs. 25 to 32). Tests at atmospheric altitude conditions have been performed in an environmental 

chamber (Ref. 33), where temperature and pressure were varied simultaneously. These tests were mostly 

motivated by external aerodynamics applications. Tests at above-atmospheric pressures at room 

temperature were reported in References 34 and 35, and provide information relevant to internal flow 

applications. 

The actuator performance has to be tested in conditions that simulate the operating conditions in 

flight, at the location where the actuator is placed. The applications in the jet engine are particularly 

challenging because of the high temperatures and high pressure ratios in several engine components. It is 

desirable to find a simpler approach to eliminate this technical complication. The question that arises is 

how to simulate the conditions in an operating jet engine in order to test the actuators properly. Because 

the principle of operation of DBD actuators depend on electrical discharges and on the associated force 

generation mechanisms, the performance of the actuator will be affected by the pressure, temperature and 

properties of the gas. It is assumed that for the range of temperatures in atmospheric flight and in the jet 

engine, the gas density alone is the significant gas property influencing the performance of the actuator. 

Therefore the flight conditions can be simulated by matching the density in the laboratory. It is a simple 

idea that has not been proposed before. 

The outline of this paper is as follows. After addressing a relatively simple case of an air vehicle 

flying at altitude, attention is focused on the jet engine applications. First, information is provided on 

pressure, temperature, density, unit Reynolds number, Mach number, and velocity distribution along the 

flow-path of four thrust classes of jet engines. This information is derived from non-proprietary engine 

models used in system analysis studies by NASA. These models include cycle, flow path and sizing. 

Then, by setting the chamber pressure at room temperature, the density is matched to the in-flight 

operating conditions. The range of the needed chamber pressures was calculated using ideal gas law, by 

matching the flow densities. It depends on the placement of the actuator, and the results are presented for 

the four engine classes at takeoff and cruise conditions. This information is useful as a guideline for 

testing requirements of DBD plasma actuators at engine flight conditions. The engine information 

documented here, which is often hard to find in publically available sources, is useful for evaluation of 

other types of actuators as well as for other purposes. 

Nomenclature 

H Altitude 

M Mach number 

P Static pressure 

R Gas constant 

Rey Reynolds number 

T Static temperature

V Velocity 

X Axial distance along the engine 

 Density 

Subscripts: 

c Conditions in chamber 

Acronyms: 

DBD Dielectric Barrier Discharge 

HPC High pressure Compressor 

HPT High pressure Turbine 
NASA/TM—2011-217006 4 

LPC Low Pressure Compressor 

LPT Low Pressure Turbine 

PAX Passengers 

Jet Engine Data Source—Engine Models 

The engine models used in this study were developed based on information available in open 

literature and empirical estimates. Cycle analysis was performed with the Numerical Propulsion 

Simulation System (NPSS) code (Refs. 36 and 37), providing performance parameters such as thrust, 

component pressure ratios, and velocities, temperatures, and pressures at each engine station. 

Aeromechanical analysis and estimates of engine and component weights were calculated using the 

Weight Analysis of Turbine Engines (WATE) code (Ref. 38), which also provides a flow path schematic 

of the engine. Generic engine models representing four different thrust classes were developed: 300-, 150- 

and 50-passenger (PAX) aircraft engines, and a military jet-fighter engine. These models are a good 

representation of actual engines. Note that the 150 PAX model is a conceptual design of a generic geared 

high bypass turbofan. The primary engine parameters of thrust, weight, overall pressure ratio, and bypass 

ratio are listed in Table 1 and the schematic of each engine is shown in Figure 3. The axial coordinates of 

the components outflow stations are listed in Table 2. 

Engine conditions were calculated at the inlet and exit flow stations of the various engine 

components. Data was acquired for two engine operating conditions: sea-level takeoff and altitude cruise 

at 35,000 ft. Data for the 50 PAX engine is also shown for an additional cruise altitude of 65,000 ft, as 

this type of engine is also used for high-flying air vehicles. Figures 4 to 7 show the following parameters 

for each engine: static and total (stagnation) pressure, temperature, and density, as well as unit Reynolds 

number, Mach number and velocity. Ideal gas conditions were assumed. For all data shown, the flights 

Mach numbers are, M = 0.8 at cruise, and M = 0 at takeoff. Figure 6(d2) for the 50 PAX engine also 

shows cruise Mach numbers of M = 0.5, 0.6, and 0.8 at altitude of 65,000 ft. 

There is more data presented than is strictly needed for development of the subject test conditions. 

The motivation for including the extra data is to make it available to the research community because it is 

hard to find non-proprietary actual engine data. The engine data presented here is unrestricted and can be 

used for other purposes. 

Test Conditions for DBD Plasma Actutors 

Several assumptions are used to develop the test conditions in laboratory experiments in quiescent 

environment (no flow) in a chamber: 

(a) The effect of temperature and pressure on plasma kinetics and chemistry is ignored. This 

assumption is reasonable for the range of temperatures in the jet engines, and is further discussed below. 

(b) The effect of temperature on the electrical properties of the actuator, particularly on the 

capacitance of the dielectric, is negligible. The capacitance variation was calculated in Reference 33, and 

was shown to be small. 

(c) Actuator heat generation is negligible. This assumption is based on experimental observations for 

the range of power and voltages applied to conventional DBD plasma actuators. 

(d) Gas composition effects are ignored. There is a small effect of the composition of the atmosphere 

variation with altitude mainly due to variation of the oxygen/nitrogen ratio (Ref. 32). Potentially there 

may be an effect due to the presence of combustion products in the areas of the engine downstream of the 

combustors. For simplicity these effects are assumed to be insignificant. 

(e) Gas thermodynamic properties are constant (except in the engine model data calculations). 
NASA/TM—2011-217006 5 

The main assumption is that the gas density number is the only parameter that governs the physical 

process of the jet generation by the plasma discharge. A process of collisions between ions and neutral 

molecules create the forces that result in the jet. The collisions are governed by the mean-free-path and 

the number of molecules in a unit volume. Therefore, with the assumptions listed above, the gas density 

in laboratory tests should be set to be equal to the density at the application flight conditions. 

Assessment of the validity of these assumptions and possible subsequent refinements is a subject of 

future work. With these assumptions, the main factor affecting the jet generation dependence on pressure 

and temperature is captured by considering only the density, therefore, the conclusions provide at least a 

very good approximation. 

Validation of the assumption that temperature affects only the density is not trivial. It is known that 

some of the reactions between different charged molecules and electrons in the plasma are temperature 

dependent. Usually the dependence is weak except for the temperature dependence of electron attachment 

processes, which can be significant above 1800 R (1000 K). It seems that the only practical approach to 

assess the full effect of temperature and pressure on the momentum transferred to the fluid is to use 

numerical simulation of the DBD plasma actuator at different temperatures and pressures. This simulation 

is challenging and is planned to be performed in the future. The authors are not aware of any reported 

work on this topic. 

The assumption that the temperature affects only the density is very reasonable for temperatures 

under 1800 R (1000 K). The open question is if it is significant at higher temperature levels. If it turns 

out that there is divergence from this assumption at all, it is expected to affect the calculated test 

conditions only for jet engine applications in the combustion chamber and the high-pressure turbine, 

which are at temperatures higher the 1800 R, as seen from Figures 4 to 7. 

The chamber pressure and temperature to yield the same density as in flight is calculated as follows. 

The gas density in the laboratory test chamber should be the same as the density at the application 

flight conditions, 

c  

where subscript c indicates conditions in the chamber. 

Assuming an ideal gas with constant R, 

  P/ RT

the following relationship is obtained 

  c

c T T

P P

/ 

where 

Pc , Tc, c – Laboratory chamber pressure, temperature, and density. 

P, T,  – Static pressure, temperature, and density at flight conditions. 

The chamber pressures were calculated first for the case of an air vehicle. Figure 8 shows chamber 

pressures for simulating flight at altitude for a vehicle with local mach numbers of 0, 0.5, and 1, without 

shock waves, and for a 20 wedge at M = 2 with a shock wave. The calculations were performed by 

taking the density at altitude from standard atmosphere data. The atmosphere density is the stagnation 

density for the air vehicle. Then the isentropic relations were used to calculate the local density at the 

indicated Mach numbers. For the wedge, shock conditions were used to calculate the density on the 

wedge surface downstream of shock. 
NASA/TM—2011-217006 6 

The chamber pressures needed for jet engine applications were calculated from the engine model data 

for the four generic engine classes. In these calculations sea-level pressure was 17.7 psi. Sea-level 

temperature was 545.7 R (29.8 C). The latter was also taken as the value of the chamber room 

temperature. The results are displayed in Figure 9. Some notes on the calculations are provided in the next 

section. 

Discussion 

For the air vehicle application the chamber pressures decrease with altitude according to the variation 

of the density with altitude and the local flow conditions at the actuator location. The situation is more 

complex for the jet engine. The results show that the test chamber pressure varies greatly, from subatmospheric to above atmospheric pressures, depending on the operating conditions and location of the 

placement of the actuator in the engine. For example, if a DBD plasma actuator is to be placed at the inlet 

to the high pressure turbine for the 300 PAX engine, it has to be tested at 6 atm if it is intended to operate 

at sea-level takeoff conditions, and at 2.9 atm if it is intended to operate at 35,000 ft cruise. If it is to be 

used in the exit of high-pressure compressor duct (burner inlet) it has to be tested at 12.4 atm at takeoff 

conditions, and at 6.2 atm at 35,000 ft cruise conditions. If it is to be used at the low-pressure turbine exit, 

it has to be tested at 0.5 atm for operation at takeoff, and at 0.2 atm at cruise. If the actuator is to be placed 

on the low-pressure turbine of the 50 PAX engine flying at 65,000 ft, its performance has to be tested at a 

very low chamber pressure of 0.03 atm. 

Note that the calculations are based on conditions at the inflow and outflow planes of the various 

engine components. The calculated points are connected with straight lines. Further modification is 

needed to account for local flow conditions inside the component. For example, in turbomachinery there 

are inter-row and inter-stage variations, and in inter-blade passages there is acceleration or diffusion or 

even shock waves that will modify the results. Those local modifications are not included in this study 

and are left for future work. 

Note also that the results shown in Figure 9 display the chamber pressures based on total (stagnation) 

as well as static conditions in the engine. The reason that the results corresponding to total condition are 

shown is that the total conditions are equal to static condition at locations where the velocity is zero, 

corresponding to placement of the actuator at locations such as the leading edge of a turbine or 

compressor airfoil. As can be seen in the figure, the differences are not large. 

Additional insights can be gained from the distribution of the unit Reynolds number. Usually low unit 

Reynolds number may indicate flow separation. For example, it is known that there is a flow separation 

problem on the low pressure turbine (LPT) at altitude. Those locations are good candidates for 

implementation of active flow control. However, those locations are also characterized by low density, 

requiring the plasma actuator to be tested at low chamber pressures. DBD plasma actuators may suffer 

from loss of performance as the pressure is decreased (note that there are insufficient and conflicting 

results in published literature). Therefore, laboratory testing is critical to establish that the DBD actuators 

can perform adequately under those conditions. 

It is important to note that in the field of weakly ionized plasma research, laboratory experiments 

were traditionally performed in a vacuum chamber at room temperature. It therefore became common in 

that field to specify the chamber pressure as an experimental parameter. This may have led to habitually 

considering the pressure, rather than the density, as the relevant parameter. 

Conclusions 

Data on flow conditions in four generic jet engines, representing four different thrust classes, was 

presented, and because it is non-proprietary, the data is useful for various applications related to 

formulating test conditions of flow control devices placed in various engine components. The data was 

used to develop test conditions for characterization of DBD plasma actuators in a chamber at room 
NASA/TM—2011-217006 7 

temperature. The underlying assumption is that the performance of DBD actuators depends only on the 

density and that all other temperature-related effects are negligible at the temperature range existing in the 

jet engine. Based on this assumption and the engine models data, the pressure needed to be set in the test 

chamber to simulate in-flight engine operating conditions was calculated. The pressures vary with the 

location of the actuator in the engine, the type of engine, and the flight operating conditions. There is a 

wide spread in the pressure range, depending on the specific application. The pressure varies from 12.4 to 

0.03 atm for the four engine models and flight conditions studied. Chamber pressures needed to test 

actuators for flight vehicle were also shown. Unlike the engine environment, the flow conditions for 

testing flight vehicles can simply be calculated with data readily available from standard atmospheric 

tables, for any flight speed and altitude. 

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Actuators at Sub-atmospheric Pressures,” 61st Annual Meeting of the APS/DFD, San Antonio, Texas, 

Nov. 2008, http://meetings.aps.org/link/BAPS.2008.DFD.ET.9 

28. Takagaki, M., Isono, S., Nagai, H., and Asai, K., “Evaluation of Plasma Actuator Performance in 

Martian Atmosphere for Applications to Mars airplanes,” AIAA–2008–3762, 2008. 

29. Benard, N., Balcon, N., and Moreau, E., “Electric Wind Produced by a Surface Dielectric Barrier 

Discharge Operating in Air at Different Pressures: Aeronautical Control Insights,” J. Phys D: Appl. 

Phys., Vol. 41, 2008, pp. 042002. 

30. Benard, N., Balcon, N., and Moreau, E., “Electric Wind Produced by a Single Dielectric Barrier 

Discharge Actuator Operating in Atmospheric Flight Conditions: Pressure Outcome,” AIAA–2008–

3792, 2008. 

31. Valerioti, J., and Corke, T., “Effect of Dielectric Properties on Functional Relationship between 

Plasma Initiation and Ambient Pressure,” 62nd Annual Meeting of the APS/DFD, Minneapolis, MN, 

Nov 2009. http://meetings.aps.org/link/BAPS.2009.DFD.GD.1 

32. Font, G.I., Enloe, C.L., Newcomb, J.Y., Teague, A.L., Vasso, A.R., and McLaughlin, T.E., “Effects 

of Oxygen Content on the Behavior of the Dielectric Barrier Discharge Aerodynamic Plasma 

Actuator,” AIAA–2010–545, 2010. 

33. Benard, N., and Moreau, E., “Effects of Altitude on the Electromechanical Characteristics of 

Dielectric Barrier Discharge Plasma Actuators,” AIAA–2010–4633, 2010. 

34. Valerioti, J.A., “Pressure Dependence of Plasma Actuated Flow Control,” MS Thesis, University of 

Notre Dame, Indiana, 2010. 

35. Valerioti, J., and Corke, T., “Pressure Dependence of Plasma Actuated Flow Control,” 63rd Annual 

Meeting of the APS Division of Fluid Dynamics, Long Beach, California, November 2010. 

http://meetings.aps.org/link/BAPS.2010.DFD.QJ.4 

Also: Submitted for AIAA Journal. 

36. Lytle, J.K., “The Numerical Propulsion Simulation: An Overview,” NASA/TM—2000-209915, 2000. 

37. Jones, S.M., “An Introduction to Thermodynamic Performance Analysis of Aircraft Gas Turbine 

Engine Cycles Using the Numerical Propulsion System Simulation Code,” NASA/TM—2007-

214690, 2007. 

38. Tong, M.T., Naylor, B.A., “An Object-Oriented Computer Code for Aircraft Engine Weight 

Estimation,” ASME Paper No. GT2008-50062, 2008. 
NASA/TM—2011-217006 9 

TABLE 1.—PARAMETERS OF FOUR ENGINE MODELS 

Engine Thrust Weight OPR BPR

At sea-level static Bare engine Overall Pressure Ratio Bypass Ratio 

(lbf) (lb) At sea-level At 35k ft At 35k ft 

300PAX 86,700 18,400 37.8 45.7 8.3 

150PAX (future) 23,400 5,100 33.5 42.0 14.3 

50PAX 7,600 1,300 23.5 28.4 5.3 

Military 18,500 3,800 33.4 33.6 0.4 

TABLE 2.—AXIAL COORDINATES OF ENGINE MODEL COMPONENTS 

300 PAX 150 PAX 50 PAX Military 

X, 

in. 

Component 

outflow plane 

X, 

in. 

Component 

outflow plane 

X, 

in. 

Component 

outflow plane 

X, 

in. 

Component 

outflow plane 

Core Core Core Core 

–60 Ambient –45 Ambient –20 Ambient –20 Ambient 

0 Inlet 0 Inlet 0 Inlet 0 Inlet 

24 Fan 15 Fan 6 Fan 22 Fan 

24 Splitter 15 Splitter 12 Duct 25 Duct 

26 Duct 15 Duct 46 HPC 56 HPC 

46 LPC 28 LPC 48 Duct 60 Duct 

59 Duct 39 Duct 55 Burner 69 Burner 

110 HPC 65 HPC 58 HPT 77 HPT 

111 Duct 66 Duct 59 Duct 78 TDuct 

120 Burner 73 Burner 71 LPT 86 LPT 

129 HPT 77 HPT 86 Mixer 95 Mixer 

138 Duct 84 Duct 103 Nozzle 106 Duct 

181 LPT 100 LPT Bypass 146 Augmentor 

188 Duct 103 Duct –20 Ambient 190 Nozzle 

212 Core Nozzle 115 Core Nozzle 0 Inlet Bypass 

Bypass Bypass 6 Fan –20 Ambient 

–60 Ambient –45 Ambient 71 Duct 0 Inlet 

0 Inlet 0 Inlet 86 Duct 22 Fan 

24 Fan 15 Fan 86 Duct 

24 Splitter 15 Splitter 95 Mixer 

59 Bypass EGV 30 Bypass EGV 

95 Duct 47 Duct 

149 Bypass Nozzle 86 Bypass Nozzle 

 
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NASA/TM—2011-217006 31 
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1. REPORT DATE (DD-MM-YYYY)

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4. TITLE AND SUBTITLE

DBD Plasma Actuators for Flow Control in Air Vehicles and Jet Engines--Simulation of 

Flight Conditions in Test Chambers by Density Matching 

5a. CONTRACT NUMBER

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6. AUTHOR(S)

Ashpis, David, E.; Thurman, Douglas, R. 

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WBS 561581.02.08.03.21.13.02 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

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E-17602-1 

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National Aeronautics and Space Administration 

Washington, DC 20546-0001 

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Adelphi, Maryland 20783-1145 

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Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Dielectric Barrier Discharge (DBD) Plasma actuators for active flow control in aircraft and jet engines need to be tested in the laboratory to characterize their performance at 

flight operating conditions. DBD plasma actuators generate a wall-jet electronically by creating weakly ionized plasma, therefore their performance is affected by gas discharge 

properties, which, in turn, depend on the pressure and temperature at the actuator placement location. Characterization of actuators is initially performed in a laboratory 

chamber without external flow. The pressure and temperature at the actuator flight operation conditions need to be simultaneously set in the chamber. A simplified approach is 

desired. It is assumed that the plasma discharge depends only on the gas density, while other temperature effects are assumed to be negligible. Therefore, tests can be performed 

at room temperature with chamber pressure set to yield the same density as in operating flight conditions. The needed chamber pressures are shown for altitude flight of an air 

vehicle and for jet engines at sea-level takeoff and altitude cruise conditions. Atmospheric flight conditions are calculated from standard atmosphere with and without shock 

waves. The engine data was obtained from four generic engine models; 300-, 150-, and 50-passenger (PAX) aircraft engines, and a military jet-fighter engine. The static and 

total pressure, temperature, and density distributions along the engine were calculated for sea-level takeoff and for altitude cruise conditions. The corresponding chamber 

pressures needed to test the actuators were calculated. The results show that, to simulate engine component flows at in-flight conditions, plasma actuator should be tested over a 

wide range of pressures. For the four model engines the range is from 12.4 to 0.03 atm, depending on the placement of the actuator in the engine. For example, if a DBD 

plasma actuator is to be placed at the compressor exit of a 300 PAX engine, it has to be tested at 12.4 atm for takeoff, and 6 atm for cruise conditions. If it is to be placed at the 

low-pressure turbine, it has to be tested at 0.5 and 0.2 atm, respectively. These results have implications for the feasibility and design of DBD plasma actuators for jet engine 

flow control applications. In addition, the distributions of unit Reynolds number, Mach number, and velocity along the engine are provided. The engine models are nonproprietary and this information can be used for evaluation of other types of actuators and for other purposes.

15. SUBJECT TERMS

Jet engines; Aeropropulsion; Turbomachinery; Compressor; Turbine; Gas turbine; Plasma; Dielectric barrier discharge; Flow control 

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