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/REV1

July 2011

U.S. ARMY

RESEARCH LABORATORY

AIAA–2011–3730
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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/REV1

July 2011

U.S. ARMY

RESEARCH LABORATORY

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

42nd Plasmadynamics and Lasers Conference

sponsored by the American Institute of Aeronautics and Astronautics

Honolulu, Hawaii, June 27–30, 2011

AIAA–2011–3730
Acknowledgments

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

thank Scott Jones, Michael Tong, and Chris Snyder from NASA Glenn Research Center for their help in developing the engine 

models. Thanks also to Alex Likhanskii (Tech-X Corporation), Andrey Starikovskiy (Princeton University), and Mikhail Shneider 

(Princeton University) 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

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7115 Standard Drive

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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.

Document Change History

NASA/TM—2011-217006/REV1, July 2011

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 and Douglas R. Thurman

This printing replaces NASA/TM—2011-217006, March 2011, because extensive changes were made to the document.
NASA/TM—2011-217006/REV1 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 

We address requirements for laboratory testing of Dielectric Barrier Discharge (DBD) plasma actuators for 

active flow control in air vehicles and jet engines. The performance of these actuators depends on the gas discharge 

properties, which in turn, depend on the pressure and temperature. When performing a characterization of actuators 

in a laboratory chamber without external flow, the test-chamber pressure and temperature must be set to the flight 

pressure and temperature at the actuator location. It is desirable to perform the tests at room temperature because it 

is technically challenging to simultaneously set pressure and high temperature in a chamber. We make the 

assumption that the plasma discharge depends only on the gas density, and ignore all other temperature effects. The 

tests can then be performed at room temperature with chamber pressure set to match the density in flight conditions. 

We calculated the needed test-chamber pressures for altitude flight of an air vehicle, and for jet engines at sea-level 

takeoff and altitude cruise conditions. For air vehicles, we show several examples of the chamber pressure 

dependence on flight at altitudes up to 65,000 ft. For jet engines, we first generated the needed data by constructing 

generic models of four engine thrust classes; 300-, 150-, and 50-passenger (PAX) aircraft engines, and a military jetfighter engine. We calculated the static and total pressure, temperature, and density distributions along the engine for 

sea-level takeoff and for altitude cruise conditions, and we present the corresponding test-chamber pressures. The 

range is from 12.4 to 0.03 atm, depending on the altitude, the engine class, and the placement of the actuator. For 

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

12.4 atm for takeoff, and at 6 atm for cruise. If it is placed at the low-pressure turbine exit, the test pressures are 0.5 

and 0.2 atm, respectively. The engine models are non-proprietary and the data can be used for evaluation 

requirements of other types of actuators and for other purposes. We also included the distributions of unit Reynolds 

number, Mach number and velocity along the engines. 

Nomenclature 

H, 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

∞ Freestream conditions

atm Atmospheric conditions
NASA/TM—2011-217006/REV1 2 

Acronyms: 

DBD Dielectric Barrier Discharge 

HPC High Pressure Compressor 

HPT High Pressure Turbine 

LPC Low Pressure Compressor 

LPT Low Pressure Turbine 

PAX Passengers 

I. Introduction 

There is a 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 have been proposed for active flow control of various flows, and the technology 

has been an active research area in the last decade. 

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 operated in steady 

mode, or in unsteady mode (by pulsing or modulating), and it 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 application 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 a 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 the DBD actuators area has been focused on applications in external flows, 

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

turbomachinery applications. There have been several successful experimental demonstrations of active flow control 

with DBD plasma actuators to eliminate low Reynolds number separation in Low-Pressure 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. Only a few 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. Before performing flight tests, the actuators must be tested in the laboratory. The aerodynamic performance 

of the actuators must 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, must be quantified, as it is needed for design of power supplies and for costbenefit analysis of the flow control technology. This paper addresses the test conditions needed to characterize the 

actuators in the laboratory. 
NASA/TM—2011-217006/REV1 3 

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 must 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 (Refs. 25 through 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 primarily motivated by their applicability to external aerodynamics. Tests at above-atmospheric 

pressures at room temperature were reported in References 34 and 35, and were mainly motivated by their 

application to internal aerodynamics. 

The actuator performance must 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 challenging to simultaneously set 

pressure and high temperature in a test chamber, and therefore 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 depends 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 done by 

NASA. These models include cycle, flow path and sizing. Then, by setting the test-chamber pressure at room 

temperature, the density in the chamber is matched to the in-flight operating conditions. The range of the needed 

chamber pressures was calculated by matching the densities and using the ideal gas law. 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 publicly available sources, is useful for 

evaluation of other types of actuators as well as for other purposes. 

II. Jet Engine Data Source—Engine Models 

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

from 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 

were acquired for two engine operating conditions: sea-level takeoff and altitude cruise at 35,000 ft. Data for the 

50 PAX engine are 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 through 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 unit Reynolds number for cruise Mach numbers of M = 0.5, 

0.6, and 0.8 at altitude of 65,000 ft. 
NASA/TM—2011-217006/REV1 4 

There are more data presented than are strictly needed for development of the subject test conditions. The reason 

for including the extra data is to make them available to the research community because it is hard to find nonproprietary actual engine data. The engine data presented here are unrestricted and can be used for other purposes. 

III. Test Conditions for DBD Plasma Actuators 

A. Assumptions 

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

(no flow) in a test-chamber: 

1) The effects of temperature and pressure on plasma kinetics and chemistry are ignored. This assumption is 

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

2) 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. 

3) 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. 

4) Gas thermodynamic properties are constant (except in the engine model data calculations). 

5) 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). Effects due to the presence of 

combustion products in the areas of the engine downstream of the combustors, and effects of humidity are 

assumed to be insignificant. 

6) Gas viscosity is assumed not to affect the actuator performance. In principle, viscosity dependence on 

temperature can affect the development of the wall-jet generated by the actuator. (Sutherland's law was used 

in calculating the unit Reynolds number in the engine models.) 

The main assumption is that the gas density is the only parameter that governs the physical process of the walljet generation by the plasma discharge. A process of collisions between ions and neutral molecules creates the forces 

that result in the wall-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 are a subject for future 

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

is captured by considering only the density. 

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 whether it is significant at higher temperature levels. If it turns out to have a large 

effect, 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 than 1800 R, as seen from Figures 4 

through 7. 

B. Test-chamber pressure 

The chamber pressure and temperature to yield the same density as in flight are 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 of a fixed species, and therefore with constant R, 
NASA/TM—2011-217006/REV1 5 

    / RTP 

the following relationship is obtained 

   c

c TT

P P

/  

where 

Pc , Tc , c

 – Laboratory chamber pressure, temperature, and density. 

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

C. Test-chamber pressures for air vehicles 

The chamber pressures Pc were calculated first for the case of an air vehicle (Figure 8). An air vehicle is flying 

at altitude h at Mach number M . The atmospheric pressure at altitude h is hP  atm and the temperature is hT  atm , 

and are given by Standard Atmosphere tables. The local Mach number at the actuator location is M, which can be 

larger or smaller than M , depending on the flow development on the body, as determined by its geometry and 

flight conditions. We assume isentropic flow without shock waves. Using isentropic relationships at the freestream 

and at the actuator placement location with common stagnation conditions, and taking the atmospheric conditions at 

altitude h as the static conditions, we obtain: 

 

 

 

  

 M

M

f

f

hT

hP

TP

atm

atm cc 

where 

  1

1

2 M

2

1 1M 







 



     f 

The chamber pressures for the following special cases are: 

1) Actuator placed on a stationary vehicle at altitude h:    0M ,0M

Therefore f     f  ,1M ,1M and we obtain from Eq. (4), 

 

  hT

hP TP

atm

atm  cc 

This result is expected, as the static density on a stationary vehicle is the same at the atmosphere density. 

2) Actuator placed on a flat plate at altitude h at zero angle of attack:   MM

We obtain from Eq. (4) the same relationship, 

 

  hT

hP TP

atm

atm  cc 

This result is also expected, as the static conditions are the same as the atmospheric conditions. 

 

3) Actuator placed at the stagnation point (as is common for flow control of airfoil leading edge stall): 

 0M , and we obtain from Eq. (4) 
NASA/TM—2011-217006/REV1 6 

 

    

 M

1

T h f

P h P T

atm

atm c c 

The chamber pressure Pc is shown in Figure 8 as a function of altitude for the first two cases (Eq. (6)), and for 

flight Mach numbers M  0.5 , with local Mach numbers of M  0 (stagnation point) and M  0.8 ((Eq. (4)). A 

curve for M 1 with local Mach number of M  0 (stagnation point), representing an extreme case, is also 

shown. An example of a flow with shocks is also presented in Figure 8, for an actuator placed behind the oblique 

shock wave on a 20 wedge at M = 2. Isentropic conditions were assumed upstream and downstream of the shock, 

and oblique shock relations were used to calculate the flow conditions on the wedge surface downstream of the 

shock. 

The calculations are relatively straight forward for the shown simplified air vehicle cases. Obtaining the density 

at the placement location of actuators in actual configuration and flow conditions will involve obtaining data from 

CFD or measurement, taking into account viscosity and non-isentropic flows. 

D. Test-chamber pressures for jet engines 

The situation for gas turbine engine application is more complex than for air vehicles because of the 

turbomachinery components and the combustion processes. The test-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 14.7 psi, and 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. 

IV. Discussion 

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

with altitude and 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 across the different engine classes, from sub-atmospheric 

to above atmospheric pressures, depending on the operating conditions and location 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 must 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 the high-pressure compressor duct (burner 

inlet), it must 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 must 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 

must 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 conditions are shown is that the 

total conditions are equal to static conditions 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, a low unit 

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

on the low-pressure turbine (LPT) suction surface at altitude. Low Reynolds number 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 density is decreased (note that there are insufficient and conflicting results in the published 

literature on this matter). Therefore, laboratory testing is critical for establishing that the DBD actuators can perform 

adequately under low density conditions. 

 
NASA/TM—2011-217006/REV1 7 

It is important to note that for research in the field of weakly ionized plasma, 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. 

V. Conclusions 

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

and, because they are non-proprietary, the data are useful for various applications related to formulating test 

conditions of flow control devices placed in various engine components. The data were used to develop test 

conditions for characterization of DBD plasma actuators in a chamber at room temperature. The underlying 

assumption is that the performance of DBD actuators depends only on the density and that all other temperaturerelated effects are negligible over the temperature range existing in the jet engine. Based on this assumption, and the 

engine models’ data, the required test-chamber pressure for simulating 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 classes’ models and the flight conditions studied. Chamber 

pressures needed to test actuators for flight vehicles 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. Modifications can be added to account for local flow conditions at actuator 

placement. The test pressures presented rely heavily on the density matching assumption. Deviation from this 

assumption is not expected to be large, and it is believed that the results are correct to at least first order. 

References 

1. Lord, W.K., MacMartin, D.G., and Tillman G., “Flow Control Opportunities in Gas Turbine Engines,” AIAA–

2000–2234, 2000. 

2. Likhanskii, A.V., Shneider, M.N., Macheret, M.O., and Miles, R.B., “Modeling of Interaction Between Weakly 

Ionized Near-Surface Plasmas and Gas Flow,” AIAA–2006–1204, 2006. 

3. Likhanskii, A.V., Shneider, M.N., Opaits, D.F., Miles, R.B., and Macheret, S.O., “Numerical Modeling of DBD 

Plasma Actuators and the Induced Air Flow,” AIAA–2007–4533, 2007. 

4. Likhanskii, A.V., “Study of Plasma Phenomena at High Electric Fields in Applications for Active Flow Control 

and Ultra-Short Pulse Laser Drilling,” Ph.D. Dissertation, The Pennsylvania State University, College Station, 

PA, 2009. 

5. Corke, T.C., Jumper, E.J., Post, M.L., Orlov, D., and McLaughlin, T.E., “Application of Weakly-Ionized 

Plasmas as Wing Flow-Control Devices,” AIAA–2002–350, 2002. 

6. Suzen, Y.B., Huang, P.G., Jacob, J.D., and Ashpis, D.E., “Numerical Simulations of Plasma Based Flow 

Control Applications,” AIAA–2005–4633, 2005. 

7. Suzen, Y.B., Huang, P.G., and Ashpis, D.E., “Numerical Simulations of Flow Separation Control in LowPressure Turbines Using Plasma Actuators,” AIAA–2007–937, 2007. 

8. Samimy, M., Adamovich, I., Webb, B., Kastner, J., Hileman, J., Keshav, S., and Palm, P., “Development and 

Characterization of Plasma Actuators for High Speed and Reynolds Number Jet Control,” Experiments in 

Fluids, Vol. 37 (4), 2004, pp. 577–588. 

9. Moreau, E., “Airflow Control by Non-Thermal Plasma actuators,” J. Phys. D: Appl. Phys. Vol. 40, 2007, pp. 

605–636. 

10. Corke, T.C., Post, M.L., and Orlov, D.M., “SDBD Plasma enhanced Aerodynamics: Concepts, Optimization 

and Applications,” Progress in Aerospace Sciences Vol. 43, 2007, pp. 193–217. 

11. Corke, T.C., Post, M.L., and Orlov, D.M., “Single Dielectric Barrier Discharge Plasma Enhanced 

Aerodynamics: Physics, Modeling and Applications,” Exp. Fluids, Vol. 46, 2009, pp. 1–26. 

12. Corke, T.C., Enloe, C.L., and Wilkinson, S.P., “Dielectric Barrier Discharge Plasma Actuators for Flow 

Control,” Annu. Rev. Fluid Mech. Vol. 42, 2010, pp. 505–29. 

13. Hultgren, L.S., and Ashpis, D.E., “Demonstration of Separation Delay with Glow-Discharge Plasma 

Actuators,” AIAA–2003–1025, 2003. 

14. List, J., Byerley, A.R., McLaughlin, T.E., and VanDyken, R.D., “Using a Plasma Actuator to Control Laminar 

Separation on a Linear Cascade Turbine Blade,” AIAA–2003–1026, 2003. 

15. Huang J., Corke T.C., and Thomas F.O., “Plasma Actuators for Separation Control of Low Pressure Turbine 

Blades,” AIAA J., Vol. 44, 2006, pp. 51–57. 
NASA/TM—2011-217006/REV1 8 

16. Huang J., Corke T.C., and Thomas F.O., “Unsteady Plasma Actuators for Separation Control of Low-Pressure 

Turbine Blades,” AIAA J., Vol. 44, 2006, pp. 1477–87. 

17. Morris, S.C., Corke, T.C., VanNess, D., Stephens, J., and Douvillev, T., “Tip Clearance Control Using Plasma 

Actuators,” AIAA–2005–782, 2005. 

18. VanNess, D.K., Corke, T.C., and Morris S.C., “Turbine Tip Clearance Flow Control Using Plasma Actuators,” 

AIAA–2006–21, 2006. 

19. Douville, T., Stephens J., Corke T., Morris S., “Turbine Blade Tip Leakage Flow Control by Partial Squealer 

Tip and Plasma Actuators,” AIAA–2006–20, 2006. 

20. VanNess, D.K., Corke, T.C., and Morris S.C., “Tip Clearance Flow Visualization of a Turbine Blade Cascade 

with Active and Passive Flow Control,” ASME Paper No. GT2008-5070, 2008. 

21. Göksel, B., Fischer, M., Rechenberg, I., and Thallemer, A., “Elektrostatischer Plasma-Wellantrieb für Bionische 

Luftschiffe,” Proceedings of the German Aerospace Congress 2005, Friedrichshafen, Germany, Deutsche 

Gesellschaft für Luft- und Raumfahrt Paper No. DGLR-2005-261, Vol. 3, 2005, pp. 1853–1856. 

22. Göksel, B., Private Communication, 2008, http://www.electrofluidsystems.com/airfish/b-ionic-airfish2008.wmv [cited 15 June 2011]. 

23. Sidorenko, A., Budovsky, A., Pushkarev, A., and Maslov, A., “Flight Testing of DBD Plasma Separation 

Control System,” AIAA–2008–373, 2008. 

24. Grundmann, S., Frey, M., and Tropea, C., “Unmanned Aerial Vehicle (UAV) with Plasma Actuators for 

Separation Control,” AIAA–2009–698, 2009. 

25. Abe, T., Takizawa, Y., Sato, S., and Kimura, N., “A Parametric Experimental Study for Momentum Transfer by 

Plasma Actuator,” AIAA–2007–187, 2007. 

26. Gregory, J.W., Enloe, C.L., Font, G.I., and McLaughlin, T., “Force Production Mechanisms of a DielectricBarrier Discharge Plasma Actuator,” AIAA–2007–185, 2007. 

27. Schuele, C.Y. and Corke, T., “Characteristics of Single Dielectric Barrier Discharge Plasma Actuators at Subatmospheric Pressures,” 61st Annual Meeting of the APS/DFD, San Antonio, Texas, Nov. 2008, 

http://meetings.aps.org/link/BAPS.2008.DFD.ET.9 [cited 15 June 2011]. 

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 [cited 15 June 2011]. 

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 [cited 15 June 2011]. 

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/REV1 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/REV1 31 
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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 

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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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John H. Glenn Research Center at Lewis Field 

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

13. SUPPLEMENTARY NOTES

This printing replaces NASA/TM--2011-217006, March 2011. 

14. ABSTRACT

We address requirements for laboratory testing of Dielectric Barrier Discharge (DBD) plasma actuators for active flow control in air vehicles and jet engines. The performance 

of these actuators depends on the gas discharge properties, which in turn, depend on the pressure and temperature. When performing a characterization of actuators in a 

laboratory chamber without external flow, the test-chamber pressure and temperature must be set to the flight pressure and temperature at the actuator location. It is desirable to 

perform the tests at room temperature because it is technically challenging to simultaneously set pressure and high temperature in a chamber. We make the assumption that the 

plasma discharge depends only on the gas density, and ignore all other temperature effects. The tests can then be performed at room temperature with chamber pressure set to 

match the density in flight conditions. We calculated the needed test-chamber pressures for altitude flight of an air vehicle, and for jet engines at sea-level takeoff and altitude 

cruise conditions. For air vehicles, we show several examples of the chamber pressure dependence on flight at altitudes up to 65,000 ft. For jet engines, we first generated the 

needed data by constructing generic models of four engine thrust classes; 300-, 150-, and 50-passenger (PAX) aircraft engines, and a military jet-fighter engine. We calculated 

the static and total pressure, temperature, and density distributions along the engine for sea-level takeoff and for altitude cruise conditions, and we present the corresponding 

test-chamber pressures. The range is from 12.4 to 0.03 atm, depending on the altitude, the engine class, and the placement of the actuator. For example, if a DBD plasma 

actuator is placed at the compressor exit of a 300 PAX engine, it must be tested at 12.4 atm for takeoff, and at 6 atm for cruise. If it is placed at the low-pressure turbine exit, 

the test pressures are 0.5 and 0.2 atm, respectively. The engine models are non-proprietary and the data can be used for evaluation requirements of other types of actuators and 

for other purposes. We also included the distributions of unit Reynolds number, Mach number and velocity along the engines.

15. SUBJECT TERMS

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

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