Volume 51, Number 3 & 4/1999

and more . . .

Air Breathing ChemicalPropulsion: Meeting theChallenges

Timing is Everything p. 2

Laser Solved Mysteries p. 14

Going with the Coflow p. 24
Professor Ronald K. Hanson has been affiliated with the Mechanical Engineering Department of Stanford University since 1971, where he is the Woodward Professor and Department Chairman. Professor Hanson’s research has been in the fields of laser diagnostics, chemical kinetics, combustion, and propulsion. He is an internationally recognized

leader in the development of laser-based diagnostic methods for combustion and propulsion.

In the 1970s he pioneered the use of tunable IR diode laser absorption spectroscopy for non-intrusive measurements of

species and temperature in combustion environments. In the early 1980’s his group established a method, known as PlanarLaser-Induced Fluorescence (PLIF) and now used worldwide, for obtaining instantaneous 2-D images of combustion propertiesby means of laser excitation and digital camera recording. PLIF provides unique capability to observe complex flow structureswhich are currently beyond computation or measurement by other methods. Measurable quantities include species,temperature, pressure, and velocity. In the 1990s, Prof. Hanson and his students began to utilize room temperature tunablediode lasers (TDLs) as light sources in absorption diagnostics. These lasers offer high potential for rugged, compact, andeconomical diagnostics suitable for use in research and development. The potential of these diode lasers for sensing andcontrol of practical combustion and propulsion systems was recently validated through a collaboration with Naval AirWarfare Center (NAWC), China Lake, CA. Prof. Hanson’s group designed and implemented a multiple-wavelength diodelaser system which provided rapid sensing of temperature and key species in a forced combustor at NAWC. The temperaturedata were used in a fast, active control scheme to optimize combustor performance, with a particular objective of achievinghigh combustion efficiencies and minimum emissions of pollutants such as CO and unburned hydrocarbons. This pioneeringimplementation was highly successful, and has led to current research on sensors for other pollutant species.

A key advance in Prof. Hanson’s diode laser research was the development of measurement concepts based on the

shapes and positions of absorption lines. This measurement strategy enables simultaneous determination of multipleparameters, including quantities such as mass and momentum flux, which have not been measured previously by opticalmeans. During the past few years, under support from ONR and AFOSR, these TDL diagnostics have been applied assensors in research on combustion control, yielding near-real-time sensing ability which has aided development of newtypes of forced-combustion devices of interest to the Navy. At present, Prof. Hanson’s group is working to extend TDLdiagnostics to control gas turbine combustors, and to new applications in advanced propulsion systems such as pulsedetonation engines.

Professor Ronald K. Hanson

Mechanical Engineering Department

Stanford University

Stanford, CA

 could be understood, the

s can be validated, and the

ility can be improved. To this

various diagnostic technique

sition of state-of-the-art

ombustion. Prof. Ron Hanson

in his article, addresses the

ach he has taken to make

with the temporal and spatial

ol involves control of plume

ure and emission, as well as

n control of the jet plume with

lted in the invention of a novel

tor control. Here the exhaust

al fins, is performing the thrust

minating the drag induced by

s fins or bend-bodies. Prof. A.

e University and Prof. Paul

of Minnesota address this

on generated by the ONR basic

ssfully applied in various

gns such as ducted rockets,

ors, and stator-rotor designs.

Mongia of General Electric

he challenges in the design of

 combustion systems. The

onal combustion dynamics are

 probability density function

on and validation of various

described.

ogram is well coordinated with

and industry so that research

into applications. A balanced

 experimental approach is

 excellent contributions in the

eas with innovative concepts,

d experimentation.

Gabriel D. Roy

Office of Naval Research
Volume 51, Number 3 & 4 1

Timing is Everything

A Study on Active Combustion Control

for Advanced Combustors

by K. H. Yu and K. C. Schadow

Laser Solved Mysteries

Diode-Laser Absorption Sensors for

Combustion Measurements and Controls

by Ronald K. Hanson and Douglas S. Baer

Going with the Coflow

Vectoring Thrust Using Confined Shear

Layers

by P. J. Strykowski and A. Krothapalli

Runway Models

Aero-Thermal Design Methodology for

Gas Turbine Combustion Systems:

Future Direction

by Hukam C. Mongia

Profiles in Science

Dr. Ronald K. Hanson

2

14

24

35

45
2 Naval Research Reviews
Volume 51, Number 3 & 4 3

T

he Navy wants to build a better combustor—a dump combustor

to be more specific; the kind of passive intake, liquid-fueled device found on ramjet missiles and other military technology. In the

past, scientists have improved performance by changing the shape

of the combustor, a trial-and-error-method that is both expensive and

time consuming. Researchers at the Naval Air Warfare Center, China

Lake, believe the key lies in timing, not geometry.

In the following article, Yu and Schadow study the physical

mechanisms and critical processes that affect combustion control, and

experiment with fuel-injection and fuel-distribution techniques. They

were able to improve performance by actively controlling the injection of fuel directly into the combustor.

The switch in emphasis from space-based to time-based

control eliminates the need for costly design changes and

makes it possible to construct new combustors from

readymade components using an open-architecture system. Ultimately, these results may

aid the researchers in developing a practical combustor control system for future propulsion systems. - J.P.
4 Naval Research Reviews

K. H. Yu* and K. C. Schadow

Naval Air Warfare Center

China Lake, CA

A Study On Active Combustion

Control For Advanced Combustors

Abstract

Active combustion control (ACC) is an emerging art of

regulating combustion performance using a dynamic

hardware component that rapidly modifies combustion input.

With the recent advances in electronics, it is becoming a very

popular technology in propulsion and power industries. Yet,

adapting such technology into practical combustors has

proved difficult, due to lack of sufficient physical understanding. In this paper, we describe an experimental effort at

the Naval Air Warfare Center to study basic mechanisms of

ACC involving a dump combustor and present some of the

results which will shed more light on various issues and

processes of practical significance. The focus of our study

was to provide scientific basis for implementing a fastresponse in-situ closed-loop ACC system for liquid-fueled

ramjet dump combustors. This technology employs an openarchitecture-type design, and a successful transition of the

results will revolutionize the way that combustion control is

performed in advanced propulsion systems.

1. Introduction

As the requirements for future military combustors become

increasingly demanding, an advanced combustion control

system that can effectively shorten the combustor development time and improve the combustor performance will be

an important technological asset to our military. Active

combustion control (ACC) is an attractive idea because it

relies on proper timing of fuel injection rather than spatial

changes of flowfield as required in passive approaches.

Since timing adjustment is simpler than the potential

geometry modifications associated with passive control, ACC

provides flexibility in performance and eliminates costly

design changes.

In the past, studies on ACC have been motivated by undesirable combustor behaviors, that include combustion instabilities [1-10], poor burning efficiency [10-13], limited

operational range[10,14], and excessive production of

pollutants. [10,12,13,15] These studies have contributed

greatly to the present understanding of fast-response ACC,

but most of the previous ACC studies utilized gaseous fuel

instead of more practical liquid fuel. Even in the studies that

involve liquid fuel, fuel was either added upstream in a prevaporized state [6,9] or injected in a steady fashion [13],

compromising the practicality of the ACC system and

limiting the flexibility associated with temporal responsiveness.

Several technological challenges still remain before the ACC

technique can be implemented to practical propulsion

systems. One such challenge is to use liquid fuel for control

and minimize the fuel amount by directly pulsing it into the

combustion chamber. Because of the combustion delays

associated with liquid-fuel atomization, droplet heating,

vaporization and burning processes, it has been difficult to

perform fast-response in-situ control using liquid fuel. Also,

it was difficult to actuate liquid-fuel injection at high

frequencies. As a result, such a control has not only been

difficult to obtain, but the physical processes were also

poorly understood.

The present study was motivated by our desire to extend

ACC to liquid-fueled combustors and explore the use of ACC

approach in practical propulsion systems. Recent progress in

liquid-fuel actuator technology and enhanced understanding

of the combustor dynamics provided an ideal background to

further advance ACC technology to liquid-fueled processes.

Thus, novel ACC experiments have been performed in a

generic dump combustor using pulsed liquid fuel injection.

The objectives are to establish closed-loop ACC using direct

liquid-fuel injection, better understand the physical mechanisms as well as limitations, and provide the scientific basis

 * Presently Associate Professor, Aerospace Engineering

Department, University of Maryland, College Park, MD.
Volume 51, Number 3 & 4 5

for applying fast-response in-situ combustion control in

practical propulsion systems. Ultimately, this will lead to the

development of practical ACC technology for advanced

liquid-fueled ramjets.

2. Theoretical Consideration

Understanding heat release dynamics is an important

prerequisite for designing an efficient ACC system. Heat

release process is critical not only for thermodynamic

considerations but it is also important due to interactions

with fluid dynamic and acoustic processes, that affect

combustor performance. To enhance combustor performance

actively requires an ability to modify the dynamics of hear

release process both spatially and temporally. While

actuators modulate fuel mass flux providing some means of

temporal control, droplet interaction with flow features could

be used to disperse fuel droplets affecting spatial heat

release. Also, as will be shown in this section, fuel droplet

size is an important factor which affects both timing and

location of heat release.

2.1 Combustion-Acoustic Interaction (timing of heat

release)

The interaction between pressure fluctuation and heat release

fluctuation is an important consideration in ACC, because,

depending on the relative phase lag between the fluctuations,

the interaction may result in either a sourcing or a damping

term for acoustic energy. According to well-known Rayleigh

criterion [16,17], the change in acoustic energy density after

one oscillation cycle can be written in terms of pressure

oscillation, p’, and heat release oscillation, q’.

∆E dV p q

p

dt dV

V V

t

t F ≈∆ = − ∫ ∫∫ + ε γ

γ

1 1 ' ' /

 (1)

where higher order terms are neglected. Equation (1) can be

derived from the governing equations of continuity, momentum, energy, and state by applying perturbation technique

and linearizing the results.

According to Equation (1), if p’ and q’ are out of phase, then

their product will be negative and the interaction results in

acoustic energy loss, damping the oscillations. Therefore,

one key strategy in active instability suppression is to

actively create additional q’ at out of phase with respect to p’

so that the oscillation will be damped by acoustic energy

loss. In general, since q’ is very sensitive to the instantaneous fuel flux, a small amount of fuel flux can be modulated to achieve this task. The relative amplitude of properly

phased q’ necessary to suppress instability depends on how

promptly the instability must be suppressed as well as the

amplitude of instability p’. For instance, the acoustic energy

contained in an oscillation cycle is

E dV

V

= ∫ ε where

ε ρ

ρ

= + 



 















F ut ∫  p t

c

dt

F 2

2

2

1

'( ) '( )

/

(2)

whereas the maximum amount of suppression possible when

p’ and q’ are perfectly sinusoidal and out of phase everywhere can be deduced from Equation (1) as

∆ε γ

γ ≈ − () ' ' −1 max max

2

 

 

p q

p F (3)

To completely suppress the total amount of acoustic energy

in Equation (2) after only one cycle, the amplitude of

controlled heat release needs to be at least

q f c u

p ' p ( )

'

' ' max

max

≈ max −

( ) +















2 1 

2

γ

ρ 

(4)

Smaller q’ will take more than one cycle to suppress the

oscillations.

2.2 Flow Structure-Droplet Interaction (location of heat

release)

In order to supply the fuel droplets into a desired location,

one can point the injector to the location and directly inject

the droplets. While this is intuitively straight forward, it will

work only for large droplets that are relatively unaffected by

fluid motion – those droplets that are unsuitable for combustion control. The better method for combustion application

is to use the fluid motion and the interaction of fuel droplets

with the fluid motion to guide the droplets.

The extent of the interaction between liquid fuel droplets and

large-scale turbulent flow features such as vortices is

expressed in terms of the particle Stokes number, Sto, which

is the ratio of the characteristic time scale for particles

responding to the fluid motion and that for the flow dynamics. The particle response time is estimated by assuming the

drag on the particle (or fluid droplets [18]) follows Stokes’s

law (Re<1). For particle density ρf

 >> ρg, the response time

is taken to be

τ

ρ

μ p

f

g

D =

2

18 (5)

which equals the time needed for the velocity difference to

be reduced to e-1 of the original velocity difference. To

assess the interaction with large vortices, the flow time scale
6 Naval Research Reviews

can be defined as the large eddy turn-over time or simply the

period of vortex shedding. Then the resulting Stokes number

can be written as

Sto f D F

g

= ρ

μ

 2

18 (6)

and Sto determines the degree of interaction between the

droplets and vortices.

The droplets with very small Sto follow fluid motion very

closely with their physical location being determined entirely

by fluid motion, while the droplets with very large Sto are

nearly unaffected by fluid motion and follow ballistic

trajectories. Recent experiments [19-21] and computations

[22,23] suggested that only those droplets with Sto on the

order of unity are affected by fluid motion and yet have

enough momentum to be dispersed further outside the carrier

fluid path. Because small droplets burn efficiently with short

time delay, they are desirable for combustion control. For

controlling spatial fuel distribution, those droplets with

Stokes number close to unity would also be desirable,

because they disperse even further outside via partial

interaction with the structures yet sufficiently small for

effective combustion. If droplet size distribution was such

that those droplets with the latter characteristics constituted

the upper end of the size distribution, more control over

spatial distribution of fuel droplets could be obtained without

compromising efficient combustion.

2.3 Effect of Fuel Droplet Size (factors affecting the heat

release)

As shown in the previous section, it is important to control

the timing and location of q’ for instability suppression.

While one can influence the relative phase of q’ by controlling the timing of additional fuel flux, there is certain time

lag between the fuel flux modulation and the resulting effect

on heat release. For liquid-fueled control, the time lag is a

sensitive function of the fuel type as well as the actuator and

the flowfield characteristics. In particular, fuel-drop size

distribution will have a significant impact because of the

delays associated with droplet heating and vaporization

processes.

In general, liquid fuel droplets will attenuate the effect of

fuel flux modulation by delaying the mixing process and thus

adversely affecting combustion response. A wide distribution of fuel droplet sizes will further attenuate the response.

To assess the effect of droplet size, one may consider droplet

heating and vaporization stages separately. As a first order

approximation, one may assume that no vaporization takes

place while the droplets are being heated and the heat flux

into the droplets during this period is used to uniformly raise

the droplet temperature. While these assumptions are not

rigorous, they are nevertheless adequate approximations in

the present case where the vapor pressure is low and the Biot

number for droplet heating is small.

Under these assumptions, the fuel temperature during the

droplet heating period can be written as

Tt T T h

c D f f g t T

f p

g ( ) ( ) exp = − ( )  −















0 +

6

ρ 0 (7)

which is applicable until the droplet temperature reaches the

boiling point of fuel. Then, the characteristic time scale

associated with the induction period can be related to the fuel

droplet size as

τ

ρ

h

f pc D

h ∝ 0

6 (8)

Furthermore, the average heat transfer coefficient, h, which

depends on the droplet Reynolds number, Reƒ

, would vary as

D0

-0.4 [24] in the range of conditions likely to be encountered

here. Therefore, the induction time associated with the

droplet heating depends on the initial droplet size and would

vary according to

τ h ~ D .

0

1 4 (9)

which implies that the heat release process will not be

synchronized unless the initial droplet size is uniform.

In this simple model, the droplet starts vaporizing when it is

heated to the boiling temperature. For the vaporization

process, one may further assume that all of the heat flux

transferred into the droplet is used to vaporize the surface

layer so that

hDT T t hV π ∂ ρ∂ g f f fg f

2 ( )( ) − * = − (10)

which yields the rate of vaporization

− =

dV −

dt

hD T T

h

f g fb

f fg *

π ( )

ρ

2

(11)

Equation (11) has a maximum value when t*=0. The heat

release rate follows the rate of vaporization, provided that the

mixing and combustion time delays are either much smaller

or independent of the vaporization time scale. Then, the

maximum amplitude of controlled heat release associated

with single droplet burning is

Q

T T hh D

h

g fb RP

fg max

( ) = π − 0

2

4 (12)
Volume 51, Number 3 & 4 7

Since the number of droplets is inversely related to the

volume of each droplet, the maximum magnitude of heat

release oscillations for a given heat release potential is

Q N Q const

hT T

h D

T

h D D g fb

fg fg

( ) ∑ = =⋅ − ∆ −

max max . ( ) ~

0

0

1 4

(13)

This implies that pulsed fuel sprays with smaller droplet size

will be more effective in generating controlled heat flux

oscillations.

3. Experiments on Active Instability

Suppression

For active instability suppression, the approach we have

taken is to pulse a small amount of liquid fuel at the instability frequency and adjust the timing using a closed-loop

circuit. Figure 1a shows a schematic of a typical closed-loop

ACC system applied to a ramjet dump combustor. The

closed-loop ACC system consists of three separate components – sensor, controller, and actuator. The sensor is used to

detect a change in combustion process, and the controller is

used to make the proper actuation decision based on the

input from the sensor. The actuator then modifies the

combustor input accordingly. The ACC system can be

implemented in an open-architecture design by having each

component in a modular form thus making an upgrade

simple.

During unstable combustion process in dump combustors,

large vortical structures are shed from the dump plane. [25]

These vortices, that are typically formed at the instability

frequency, play an important role in sustaining the oscillations. [25,26] The actuation may be best served by modifying such processes that affect the instability mechanism. In

the present study, because our emphasis was on extending

active control to liquid-fueled combustors, a simple phasedelay circuit was utilized instead of a more sophisticated

controller using adaptive technique. Figure 1b shows the

actual configuration which was used to actively regulate the

fuel injection scheduling at the dump plane. A pressure

transducer was used as a sensor. In the controller, the

pressure transducer signal was filtered and phase-shifted to

produce square wave that regulated the fuel flux through the

actuators. Two sets of actuators with different atomization

characteristics were used in the experiment.

Table 1 shows the dimensions as well as the configuration of

the model ramjet dump combustor, in which the ACC

experiments were performed. Air was supplied from highpressure storage tanks and was metered using a choked

orifice nozzle which also defined the upstream acoustic

boundary of the inlet. Fuel was injected both in a steady

fashion and in a temporally controlled way at the dump

plane. A long-radius type exhaust nozzle was used at the

downstream end of the combustor. In an actual system, the

actuator fuel would be only a small fraction of the total fuel.

For the steady portion of the fuel flux, ethylene was supplied

Fig. 1 – Schematics of the closed-loop control system

applied to the dump combustor.

Table 1 – Dump combustor setup for the corresponding

cases.
8 Naval Research Reviews

at a constant rate into the inlet duct. For the controlled flux,

ethanol, heptane, or JP-10, which amounted to about 2 to 30

percent of the total fuel flux, was injected through the

actuators. Instabilities were observed under certain conditions and a closed-loop fuel injection process was used to

suppress the instabilities. Table 2 shows the specific operating conditions at which the unstable conditions were

encountered. Some characteristic flow parameters are also

shown for reference.

3.1 Establishing Control Authority With Liquid Fuel

In the present experiment, a novel approach which relied on

timing-dependent fuel-droplet dispersion behavior was

utilized. As explained previously, large vortical structures

are shed during the unstable operation. The controller fuel

was injected directly into these vortices at different timing

with respect to their shedding process.

Thus, the amount of slip velocity the

droplets were subjected to was dependent on the injection timing. As a

result, the droplet dispersion behavior

was affected by the relative phase

between the fuel injection and vortex

shedding. Figure 2 shows the fuel

droplet dispersion behavior with respect

to the large vortex at various fuel

injection timings. The red contour lines

represent the phase-averaged Miescattering intensity data from fuel

droplets, which are shown superimposed with the underlying vortex

structures in blue. The droplets

clustered in the core of the jet flow

when fuel was injected after the vortex

shedding, while those fuel droplets injected ahead of the

vortex shedding were subjected to the vigorous interaction

and dispersed well into the recirculation zone.

A detailed ACC experiment was performed under Case 1

conditions. Based on average fuel flux, the thermal power

output was between 66 and 82 kW. The instability in this

case was related to the acoustic quarter-wave mode of the

inlet. [27] When the timing of the liquid fuel injection was

closed-loop controlled, the amplitude of pressure oscillations

was very responsive to the controller parameter. Figure 3

shows the change in average spectral amplitude at the

instability frequency as a function of the electronic-controlunit (ECU) phase delay. Under an improper phase setting,

20 degrees for instance, the oscillation amplitude remained

high. The maximum amplitude in this case was close to the

natural oscillation amplitude, that was obtained without

closed-loop control. At the phase setting of 110 degrees,

which corresponded to pulsed fuel injection synchronized

with the vortex shedding process, the suppression of oscillation amplitude was most effective. Sound pressure level

reduction of up to 15 dB was achieved using this method.

When the flow conditions were changed slightly in Case 1B,

the dependency on the injection phase remained unchanged

from Case 1A.

Figure 4a shows the onset of active instability suppression, as

the fuel injection timing was switched from the ECU setting

of 20 degrees to 110 degrees at time t=0. With the timing

change, the oscillation amplitude was quickly brought under

control to about one-fifth of the uncontrolled level only after

a few cycles. The injection frequency was unaffected by the

injection timing change. Figure 4b shows a comparison of

pressure spectra between a closed-loop injection case with

the properly controlled injection timing and the baseline case

without closed-loop control. The spectral amplitude was

reduced significantly at all of the higher harmonic frequencies as well as at the fundamental. The results clearly

Table 2 – Average flow conditions and characteristic

parameters for test cases.

Fig. 2 – Mie-scattering images of the flowfield showing vortex-droplet

interaction as a function of the fuel injection timing.
Volume 51, Number 3 & 4 9

demonstrates the potential of active instability suppression

technique that is based on properly-timed pulsed fuel

injection directly into the vortices.

3.2 Determining the Minimum Fuel Flux Needed

To explore more practical issues that would affect the

transition of the above results, similar experiments were

extended to higher flow conditions using the same actuators.

However, instead of ethanol, liquid heptane was used as the

controller fuel as heptane is about 66 percent more energetic

per volume than ethanol. Furthermore, in this set of experiments, the combustor dimensions were varied as well as the

flow conditions to find naturally unstable conditions at which

ACC experiments could be performed. In general, the

controller fuel flow rate (heptane) was held constant while

the main air and main fuel (ethylene) flow rates were

systematically increased. As a result, the controller fuel flux

constituted increasingly smaller portion of the total fuel flux.

The controller performance became marginal at 360 kW

output operation, suggesting a critical fuel flux was reached

for maintaining the control authority.

In Figure 5a, the amount of relative reduction in peak

spectral amplitude was plotted as a function of the inlet flow

Reynolds number. Most of the data followed a general trend

that indicated the diminishing control effectiveness with flow

scale. However, this behavior was very gradual and was

more related to the fact that the general amplitude of

instability was greater at higher output conditions. On the

other hand, there was a sharp increase in the number of

cycles needed to reach the minimum amplitude at the on-set

of ACC. Figure 5b shows the number of cycles it took to

reach the controlled state as a function of the relative heat

release potential of the controller fuel to the overall fuel.

The results indicate that the control authority was marginal

when the controller fuel flux was dropped to 8 percent of the

total fuel flux. Using controller fuel flux below 8 percent,

active instability suppression was unsuccessful in the present

case.

4. Scaling Experiments With Different

Actuators

Further scale-up tests were continued with different actuators

(No. 2) and with JP-10 fuel, which was considered the likely

fuel that would be used in practical systems. The No. 2

actuators were prototype fuel injectors that used air-assisted

mechanism to obtain much finer sprays than the previous

actuators. Air-assisted atomization required a small amount

of secondary air flow, which was less than one percent of the

air needed for stoichiometry. Figure 6 shows the transfer

functions comparing the combustor response using No. 1 and

No. 2 actuators. For the conditions at which the actuators

were operated, the Sauter-mean diameter, D32, was about four

times smaller with No. 2 actuators than with No. 1 actuators.

This significant change in

actuator characteristics also

made it possible to study the

effect of fuel droplet size on

control efficiency.

4.1 Critical Fuel Flux

In the scaling experiments, up to

eight actuators were employed

at a time. By increasing the

combustor thermal output while

systematically reducing the

number of actuators, the

minimum controller fuel flux

was again investigated experimentally. The results showed

that the minimum fuel flux was

dependent on the actuator

characteristics. Using No. 2

actuators, the minimum fuel flux

Fig. 3 – Pressure oscillation amplitude as a function of

fuel injection timing.

Fig. 4 – Comparison of controlled and uncontrolled oscillations. (a) Onset of active

instability suppression, (b) Comparison of combustor pressure spectra.
10 Naval Research Reviews

was found to be slightly below 2 percent of the total fuel

flux. This can be compared with the limit with No. 1

actuators, which was about 8 percent. Figure 7 shows the

plot of relative amount of controller fuel flux with respect to

the relative amount of oscillation amplitude in the controlled

cases. While there was slight decline in control effectiveness

with decreasing amount of controller fuel flux, the controller

became no longer effective below these limits.

Experimentally obtained values of the critical fuel flux was

substantially different in the two sets of the scaling experiments. While there are many possible reasons for this

behavior, the most significant effect will be due to the

actuator and the fuel characteristics. Also, the effect of scale

would have played a significant role. Figure 8 shows the

normalized magnitude of rms pressure oscillations as a

function of the combustor thermal output. In general, the

instability became much more severe at higher output

conditions. This could be related to the fact that the amplitude of heat release oscillations per given volume increased

with the scale-up process, as the higher output conditions

were achieved by increasing the volumetric heat release rate.

4.2 Effect of Fuel Droplet Size

During the scaling experiments, since the amplitude of

instabilities was different in each case, it follows that the

amount of fuel flux needed for control should also be

different. To eliminate the effect due to the scale, each case

was normalized by the magnitude of naturally occurring

instabilities. Thus, using Equation (4), theoretical amount of

controlled heat flux that was needed to suppress the given

amplitude of instability was calculated, and this value was

used in normalizing the controller fuel flux. Figure 9 shows

the normalized relationship between heat flux and fuel flux.

The heat flux was normalized by the theoretical amount of

required heat flux considering the uncontrolled instability

amplitude, while the fuel flux was taken as a simple ratio

with the total fuel. The results showed a good overlap

between the various cases that used three different fuels and

two actuators.

It can be inferred from Figure 9 that actuator No. 2 was more

effective than actuator No. 1 in suppressing the instabilities.

Since No. 2 actuators produced much smaller drop size than

No. 1 actuators, this trend was consistent with the expectation. The control process using No.1 actuators was not very

efficient. The actual amount of fuel needed was more than

2.4 times that amount which is theoretically sufficient for

suppressing the instability in just one cycle. However, with

No. 2 actuators, the minimum fuel amount needed for control

authority can be as low as 0.31 times the theoretical amount.

As expected, many more cycles than one were required to

suppress the instability as the controller fuel flux was

relatively low.

Fig. 5 – Control effectiveness during scale-up

experiments. (a) peak amplitude reduction .vs. inlet

flow Reynolds number, and (b) required number of

cycles to reach the controlled state .vs. pulsed fuel

ratio.

Fig. 6 – Amplitude of transfer function between the

combustor pressure oscillations and the driving signal

using No. 1 injectors with Case 1 condition and No. 2

injectors with Case 2 condition.

Fig. 7 – Suppression of instability amplitude as a

function of the relative amount of controller fuel.
Volume 51, Number 3 & 4 11

The actual amount of improvement in going from No. 1 to

No. 2 actuators was about 7.7 fold reduction in the critical

fuel flux. This can be compared with the expected dependency on drop size, which was discussed in section 2. From

Equation (13), one would expect the improvement in heat

release due to fuel drop size and fuel type to be

( ) ∆

( ) ∆

×











 ≈ −

− T h

T h

D

D

fg

fg

No

No

/

/

, .

, .

.

JP 10

heptane

0 2

0 1

1 4

7

Considering that the analysis was for a simplified model

case, the quantitative agreement in the present experiments is

very good. Needless to say, such information can be used to

estimate the critical fuel flux required for maintaining control

authority in new combustors.

4.3 Limitation of Simple Phase-Delay Controller

The closed-loop approach using simple phase-delay was very

effective in controlling instability when the oscillation

frequency remained constant. However, for other cases, in

which the instability frequency drifted with the onset of

closed-loop fuel injection, this approach was not very

practical because the oscillation amplitude periodically

returned to the uncontrolled level after initial suppression.

The main cause of this undesirable behavior was found to be

frequency-dependent phase-shift, which is imposed by the

band-pass filter in the simple controller.

Figure 10 shows a mapping of instantaneous amplitude and

fuel injection phase in a typical case with the simple closedloop control established. The shaded region represents the

baseline behavior when the fuel injection was not closedloop controlled. First, when the closed-loop control circuit

was turned on, the fuel injection caused the heat release at

the proper phase so that the oscillation amplitude started to

be suppressed (process 1 in the figure). Then, as the amplitude was lowered, the instability frequency started to drift

causing shift in the injection phase (process 2). Finally, if

the injection phase shift became large enough, the heat

release started occurring at the driving phase increasing the

amplitude back to the original level (process 3). Then, the

suppression processes repeated again as large-amplitude

oscillations returned to the original frequency.

Such undesirable behavior may be avoided by employing a

higher order control algorithm based on adaptive strategy or

by adding other simple measures to ensure that the frequency

remains constant. In adaptive control strategy, a secondary

control loop can be employed that monitors the change in

frequency and shift the phase settings accordingly. Also, in

an alternative strategy, it may be possible to combine an

open-loop injection at a constant frequency, that will prevent

shift in the oscillation frequency. These ideas are being

investigated in on-going experiments to demonstrate more

robust control strategy suitable for practical systems.

5. Concluding Remarks

The opportunity exists to advance ramjet technology by

incorporating our understanding of two-phase turbulent

reacting flow in dump combustors with fast-response active

Fig. 8 – Observed instability amplitude at various scaleup conditions.

Fig. 9 – Theoretical ratio showing the potential amount

of controlled heat release oscillations normalized by

the acoustic energy of the instability.

Fig. 10 – Observed ECU phase with respect to the

oscillation amplitude of each cycle.
12 Naval Research Reviews

combustion control (ACC), which will replace current

controls that are based on passive strategies. Typically,

passive techniques for combustion control are limited in

operability range and expensive to develop, since they

require individually tailored spatial modifications of the

flowfield and material. In contrast, active control techniques

are based on precise scheduling of temporal events, thus

minimizing unnecessary geometry changes. This switch in

emphasis from spatial to temporal precision will allow more

flexibility into the combustor design and will make it

possible to implement open-architecture-type system design,

which in the long run will revolutionize combustion control

process.

Several milestones resulted from the present investigation.

The first was to identify vortex-droplet interaction as a key

process in controlling temporal as well as spatial distribution

of controller fuel. Fuel distribution affects the heat release

dynamics which is of critical importance in combustion

control. Second, fast-response in-situ-type ACC was

demonstrated using pulsed liquid-fuel injection directly into

combustor vortices. Direct injection of fuel into the combustor is not only a practical design, but it also can be used to

minimize the amount of controller fuel needed. Third, in

scaling experiments, it was revealed that the amount of

relative fuel flux played an important role in maintaining

control authority. In the ensuing experiments, the critical

amount of relative fuel flux was quantified as a function of

the actuator and instability characteristics. Lastly, the

dependence on fuel droplet size was derived and compared

with the experimental data from the scale-up experiments.

These results could be useful in developing more practical

ACC technology for liquid-fueled dump combustors.

The present study was conducted not just to better understand the physical mechanisms and critical processes that

affect the combustion control, but we were also motivated by

our desire to develop ACC technology that will be practical

for air-breathing propulsion systems. Thus, our earlier

emphasis was on extending ACC to liquid-fueled combustors

using a simple control strategy. Upon identifying the

important physical mechanisms, we were able to inject liquid

fuel directly into the dump combustor, and optimize the

combustor performance by modifying the fuel injection

timing that affected transient heat release. Detailed measurements were made to understand and explore various physical

mechanisms and scaling issues, that may affect potential

implementation of the research results. Presently, the

investigation is continuing with an effort to explore other

performance benefits of ACC and to make the control more

robust and practical. It is our belief that such detailed

mechanistic understanding combined with practical research

goals will produce much needed scientific basis for properly

designing and implementing ACC technology into future

propulsion systems.

Acknowledgment

This study was sponsored by the Office of Naval Research,

with Dr. Gabriel D. Roy as the scientific officer.

Biographies

Ken H. Yu received his B.S., M.S., and Ph.D. degrees in

Mechanical Engineering from the University of California at

Berkeley, in 1985, 1988, and 1989 respectively. He worked at

NASA-Ames Research Center in 1985 as an aerospace

engineer. He was a postdoctoral fellow/visiting professor at

Ecole Central Paris, France, in 1989-1990. From 1990 to

1999, he worked as a physical scientist at the Propulsion

Research Lab, Naval Air Warfare Center, China Lake, CA.

He is currently an Associate Professor in the Department of

Aerospace Engineering at the University of Maryland,

College Park. His research interests are in the broad field of

fluid dynamics, combustion, and optical and laser diagnostics. His research has resulted in more than 100 scientific

publications in these subjects as well as 4 patents, some of

which are still pending. His research has been acknowledged

by three conference best paper awards from AIAA and

JANNAF.

Klaus C. Schadow is presently a Senior Scientist at the

Naval Air Warfare Center Weapons Division (NAWCWD)

after serving as Head, Propulsion Research Branch, from

1977 to 1998. He received his Ph.D. in Mechanical Engineering in 1975 from the Technical University Munich, Germany.

Dr. Schadow has been involved in a number of programs on

passive and active mixing/combustion control with application to airbreathing missile propulsion, plume signature

control, and compact incineration. Dr. Schadow has

authored/co-authored over 300 publications, which include

over 55 refereed journal publications and 12 patents.

Nomenclature

c = speed of sound

cp = specific heat at constant pressure

D = fuel droplet diameter

D32 = Sauter mean diameter of fuel droplets

D0 = initial diameter of fuel droplet

E = average acoustic energy in a cycle

F = frequency of instability

h = average heat transfer coefficient

hfg = latent heat of vaporization

hRP = enthalpy of combustion

ND = number of fuel droplets with size D

p = average pressure

p’ = fluctuating pressure

q’ = fluctuating heat release

Q = rate of heat release from burning fuel droplet

Re = Reynolds number

t, t* = time

T = temperature

u’ = fluctuating velocity
Volume 51, Number 3 & 4 13

V = combustor volume

Vf = fuel droplet volume

∆E = change in acoustic energy after one cycle

∆ε = change in acoustic energy density

ε = instantaneous acoustic energy density

ø = equivalence ratio

γ = ratio of specific heats

μ = dynamic viscosity

ρ = average density of combustor gas

ρ f = fuel droplet density

τ h = characteristic heating time

τ v = characteristic vaporization time

Subscripts:

avg = average quantity

f = characteristics of fluid drop

g = characteristics of gas

max = maximum quantity

n = related to n-th cycle

rms = root-mean-squared quantity

References

1. Candel, S.M., “Combustion Instabilities Coupled by Pressure

Waves and Their Active Control,” Twenty-Fourth Sympo.

(Inter.) on Comb., The Combustion Institute, Pittsburgh, PA,

1992, pp 1277-1296.

2. McManus, K.R., Poinsot, T., and Candel, S., “A Review of

Active Control of Combustion Instabilities,” Prog. in Energy &

Comb. Sci., 19:1-29 (1993).

3. Lang, W., Poinsot, T., and Candel, S., “Active Control of

Combustion Instability,” Comb. & Flame, 70:281-289 (1987).

4. Bloxsidge, G.J., Dowling, A.P., Hooper, N., and Langhorne,

P.J., “Active Control of Reheat Buzz,” AIAA J., 26:783-790

(1988).

5. Poinsot, T., Bourienne, F., Candel, S., and Esposito, E.,

“Suppression of Combustion Instabilities by Active Control,”

J. Propul. & Power, 5:14-20 (1989).

6. Langhorne, P.J., Dowling, A.P., and Hooper, N., “A Practical

Active Control System for Combustion Oscillations,” J.

Propul. & Power, 6:324-333 (1990).

7. Schadow, K.C., Gutmark, E., and Wilson, K.J., “Active

Combustion Control in a Coaxial Dump Combustor,” Comb.

Sci. & Tech., 81:285-300 (1992).

8. Gulati, A. and Mani, R., “Active Control of Unsteady

Combustion-Induced Oscillations,” J. Propul. & Power,

8(5):1109-1115 (1992).

9. Sivasegaram, S., Tsai, R-F., and Whitelaw, J.H., “Control of

Combustion Oscillations by Forced Oscillation of Part of the

Fuel Supply,” Comb. Sci. and Tech., 105:67 (1995).

10. McManus, K.R., Vandsburger, U., and Bowman, “Combustor

Performance Enhancement Through Direct Shear Layer

Excitation,” C.T., Comb. & Flame, 82:75-92 (1990).

11. Yu, K., Trouve, A., and Candel, S., “Combustion Enhancement

of a Premixed Flame by Acoustic Forcing with Emphasis on

Role of Large-Scale Structures,” 29th Aerospace Sciences

Meeting, AIAA-91-0367, 1991.

12. Gutmark, E., Parr, T.P., Hanson-Parr, D.M., and Schadow,

K.C., “Use of Chemiluminescence and Neural Networks in

Active Combustion Control,” Twenty-Third Sympo. (Inter.) on

Comb., The Combustion Institute, Pittsburgh, PA, 1990, pp

1101-1106.

13. Brouwer, J., Ault, B.A., Bobrow, J.E., and Samuelsen, G.S.,

“Active Control for Gas Turbine Combustors,” Twenty-Third

Symp. (Inter.) on Comb., The Combustion Institute, Pittsburgh,

PA, 1990, pp 1087-1092.

14. Gutmark, E., Parr, T.P., Hanson-Parr, D.M., and Schadow,

K.C., “Closed-Loop Amplitude Modulation Control of

Reacting Premixed Turbulent Jet,” AIAA J., 29 (12), 2155-

2162, 1991.

15. Parr, T.P., Gutmark, E.J., Wilson, K.J., Yu, K., Smith, R.A.,

Hanson-Parr, D.M., and Schadow, K.C., “Compact Incinerator

Afterburner Concept Based on Vortex Combustion,” TwentySixth Sympo. (Inter.) on Combustion, The Combustion

Institute, Pittsburgh, PA, 1996.

16. Lord Rayleigh, The Theory of Sound, Dover, p.227, 1945.

17. Putnam, A., Combustion Driven Oscillations in Industry,

Elsevier, New York, 1971.

18. Batchelor, G.K., An Intro. to Fluid Mechanics, Cambridge

Univ. Press, 1967, pp.230-238.

19. Lazaro, B.J. and Lasheras, J.C., “Paricle dispersion in a

turbulent, plane, free shear layer,” Phys. Fluids A 1: 1035-

1044 (1989).

20. Longmire, E.K. and Eaton, J.K., “Structure and control of a

partice-laden round jet,” JFM 236: 217-257 (1992).

21. Glawe, D.D. and Samimy, M., “Dispersion of Solid Particles in

Compressible Mixing Layers,” JPP 9 (1): 83-89 (1993).

22. Chung, J.N. and Troutt, T.R., “Simulation of particle

dispersion in an axisymmetric jet,”JFM 186: 199-222 (1988).

23. Chang, E. and Kaliasanath, K., “Simulations of Particle

Dynamics in a Confined Shear Flow,” AIAA J. 34(6): 1160-

1166 (1996).

24. Holman, J.P., Heat Transfer, McGraw-Hill, New York, 1981,

pp.247.

25. Yu, K.H., Trouve, A., and Daily, J.W., “Low-frequency

pressure oscillations in a model ramjet combustor,” JFM 232,

47-72 (1991)

26. Poinsot, T., Trouve, A., Veynante, D., Candel, S.M., and

Esposito, E., “Vortex driven acoustically coupled combustion

instability,” JFM 177, 265-292, (1987)

27. Yu, K.H., Wilson, K.J., and Schadow, K.C., “Liquid-Fueled

Active Instability Suppression,” Twenty-Seventh Sympo. (Inter.)

on Combustion, The Combustion Institute, Pittsburgh, PA,

1998.
14 Naval Research Reviews
Volume 51, Number 3 & 4 15

T

o learn more about combustion at the

molecular level, scientists at Stanford University

are using advanced laser technology to peek

inside an engine’s chambers. Until recently, the flows

generated during the burning process have been

difficult to quantify due to the harsh environment

presented by an operating engine.

Semiconductor diode lasers that operate in

the near infrared range are ideally suited for sensing

a range of parameters, such as temperature,

velocity, and gas concentrations, because they are

rapidly tunable to wavelengths absorbed by the

combustion gases. Another important advantage

is their compatibility with high-quality fiber optics.

The researchers can use off-the-shelf beam splitters

to send multiple sensor beams to multiple locations

from a single laser source. Information gleaned from

the rapidly evolving technology of laser-based

sensors may ultimately lead to improved

engineering designs for combustion systems. The

same technology may also result in better control

of combustion and propulsion systems, rendering

them more efficient and less polluting. - D.B.
16 Naval Research Reviews

Ronald K. Hanson and Douglas S. Baer

Stanford University

Stanford, CA

Diode-Laser Absorption Sensors for

Combustion Measurements and

Control

Summary

The increasing availability and reliability of lasers over the

last 30 years has revolutionized measurement capabilities in

research and development laboratories worldwide. In

particular, advances over the last decade in semiconductor

diode lasers which were developed primarily for fiber-optic

telecommunications, computer networks and optical data

storage applications, have enabled the development of a new

class of gas sensors that may be applied for measurements

and control of a wide range of engineering systems. These

sensors, comprised of economical and rugged diode lasers

and fiber-optic components and based on absorption spectroscopy techniques, are especially promising for nonintrusive, time-resolved, species-selective measurements in

hostile environments where the use of conventional probes

may be difficult or undesirable. Here we describe the

development and application of diode-laser sensors for realtime measurements and active control of combustion and

propulsion systems. These same sensors may also be applied

to industrial facilities for process control, and for compliance

and emissions monitoring.

Introduction

Over the last twenty-five years, researchers in Stanford

University’s High Temperature Gasdynamics Laboratory

have developed and applied novel absorption diagnostic

techniques for measurements of important quantities,

including gas concentrations, temperature, velocity, and mass

and momentum flux, in laboratory and industrial combustion, propulsion, incineration and hypersonic flows. Quantitative data for these parameters are difficult to obtain by

other methods and yet are of extreme importance in characterizing the flows produced in high-temperature or highspeed test facilities. These diagnostic strategies have

generally employed continuous-wave (cw) laser sources,

ranging from lead-salt diode lasers (in the 1970’s), ring-dye

lasers (1980’s) and most recently (1990’s), near-infrared

semiconductor diode lasers. The output wavelength of these

sources may be precisely and rapidly controlled (i.e., tuned)

to coincide with absorption transitions of specific gases of

interest.1-3 The key limitation of these line-of-sight absorption techniques, the inability to provide full spatial resolution, is generally not significant when probing flows with

reasonable or high uniformity (along the measured path).

Furthermore, the ability to measure integrated quantities

across a flow is often specifically required or highly desirable.

Semiconductor diode lasers operating in the near infrared

(IR) are attractive light sources for sensing applications

because they are rapidly tunable, small and lightweight, lowcost, efficient, robust, operate at near-ambient temperatures,

and produce narrow-bandwidth radiation over a broad

wavelength range. Figure 1 presents calculated absorption

spectra of important combustion-related gases (including

CO, CO2, H2O, NO, CH4, O2) in the spectral region currently

accessible using room-temperature semiconductor diode

lasers (0.6-3.0 microns) to illustrate the range of important

gases that may be probed at present. In general, careful

selection of the probe wavelength will allow the measurement of a single species (gas) and avoid cross-sensitivity

effects which are difficult to exclude with conventional

sensors (electrochemical cells) or low resolution absorptionbased systems (non-dispersive infrared detection systems).

Perhaps the most significant advantage in the use of near-IR

diode lasers is their compatibility with rugged, communications-grade fiber-optic hardware. Fiber optics may be used to

combine (i.e., multiplex) the outputs of multiple lasers at

different wavelengths into a single probe beam, and thus

allow simultaneous measurements of temperature and the

concentrations of multiple species along a single optical
Volume 51, Number 3 & 4 17

path. Moreover, the resultant multi-wavelength probe beam

may be subsequently divided into several paths using

inexpensive off-the-shelf fiber-optic splitters, and thus

allowing measurements of temperature and species concentrations in multiple locations simultaneously, with only a

single set of laser sources. Since the lasers presently represent the most expensive component of this measurement

system, the use of fiber optics can thus dramatically reduce

the cost per measurement location. The combination of fiberoptic and diode-laser technology with laser-absorption

spectroscopy techniques can provide a compact, rugged and

economical sensor system for species-specific measurements

that are insensitive to interferences, particulates, and background luminosity. These qualities make diode-laser sensors

ideal for process control in a variety of hostile environments,

including combustion and propulsion flows, as well as

industrial process streams.

In this work, we describe recent work by Stanford researchers in the development and application of diode-laser sensors

for measurements in and control of pulsed combustion

facilities at the Naval Air Warfare Center at China Lake

(NAWC) for shipboard incineration and at the Naval

Postgraduate School (NPS) at Monterey for an advanced

propulsion system (pulse detonation engine). These measurements performed in high-temperature, high-pressure, timevarying flows, underscore the potential of diode-laser sensors

for process control of hostile environments where conventional probes and measurement techniques are inappropriate

or ineffective.

Measurement Principles

The measurement principle for diode laser sensors is based

on the absorption of laser light at a particular wavelength

(i.e., color) by a specific gas of interest. In brief, the laser

radiation induces the molecule to undergo a transition from

an initial quantum state to a final state, which corresponds to

a particular vibration-rotation transition. The laser wavelength corresponds to the energy difference between the

initial and final quantum states. The magnitude of the

absorption is related to the population of molecules in the

absorbing state, as described by statistical mechanics and is a

function of temperature. Furthermore, the shape and relative

position of an absorption feature is related to local environmental effects such as temperature and pressure through

Doppler and pressure broadening, respectively, and is

affected by the gas velocity through the Doppler shift.

Spectrally resolved absorption measurements can exploit

these dependencies to allow simultaneous determinations of

multiple gasdynamic parameters.

The transmission of light directed through a flowfield may be

described using the Beer-Lambert relation

T I

I v k L v ≡= − v

0

exp( )

where Tν is the fractional transmittance of the medium at

frequency ν, I0 is the incident intensity of the laser beam, Iν is

the intensity after propagation through a length L [cm] of the

absorbing medium, and kν [cm-1] is the absorption coefficient

at frequency ν. The absorption coefficient is a function of the

total pressure, P [atm], the mole fraction of absorbing

species, χ, species-specific spectroscopic parameters, and the

gas temperature. At a fixed wavelength and a constant

temperature and pressure, the absorption coefficient is

proportional to the probed species mole fraction. The gas

temperature may be determined from measurements of the

transmission intensity at two selected wavelengths, or from

the shape of an absorption line dominated by Dopplerbroadening effects. The appropriate line pairs for thermometry may be determined a priori from known spectroscopic

parameters for a particular gas. The partial pressure (and

concentration if the pressure is known) of a particular gas

can be inferred from the measured absorption and temperature. For measurements of weakly absorbing gases, i.e.,

where kν

 is small, measurement sensitivity may be improved

by increasing the path length using folded-path optical

arrangements if necessary. Furthermore, measurements of the

magnitude, shape and shift of an absorbance feature may

enable determination of velocity, as we have demonstrated in

arcjet thrusters4 (where velocities reach V ≈ 15 km/s) and

hypersonic test facilities5

 (Mach numbers of 16 have been

monitored), and of electron number density and temperature6

(up to 15,000 K), as demonstrated in atmospheric pressure

plasmas and shock tubes.7

Measurements in Forced Combustion Systems

Multiplexed diode-laser absorption sensors were applied to

measure and control gas temperature and H2

O concentration

Fig. 1 – Calculated absorption spectra of important

combustion gases in the 0.6-3.0 micron spectral region

that is currently accessible using room-temperature

semiconductor diode lasers. The calculations use a

10% mole fraction for each gas, 296-K temperature,

760-torr total pressure, and 1-cm path length.
18 Naval Research Reviews

in the combustion region of an experimental (50-kW) pulsed

annular dump combustor.8 This combustor serves as a model

of an afterburner in a compact, multi-stage, waste incineration system under development at the Naval Air Warfare

Center (NAWC) at China Lake (California) for use aboard

Navy ships. The concentrations of important pollutants and

unburned fuel constituents (e.g., CO, C2H2, C2H4) were also

monitored on line in the exhaust stream.

A schematic diagram of the NAWC multi-stage incineration

concept including diode laser sensors for real-time control is

illustrated schematically in Figure 2. The incinerator

converts solid waste to gaseous waste using a starved-air

pyrolysis chamber and then removes the hazardous components using a secondary oxidation chamber or afterburner

that utilizes the concepts of forced vortex combustion for a

compact and efficient design. Details of the afterburner,

including the design, application of advanced diagnostics,

and determination of the destruction and removal efficiency

(DRE) for both small- and large-scale systems may be found

in the references.9-11

The strategy for increasing the combustion efficiency in the

afterburner involves acoustically forcing (modulating) the

input fuel and air to generate stabilized vortices which allow

for rapid and controlled mixing of the air and pyrolysis

gases. The result of better mixing is more efficient and

complete combustion and less pollution, which allows the

construction of smaller incineration systems onboard ships

where space is a premium. However, full exploitation of this

forced combustor technology has required the development

of real-time sensing and fast active control. The real-time

measurements in the combustion and exhaust regions of the

afterburner combined with active control strategies demonstrate the utility of diode-laser sensors for realistic engineering applications including performance assurance and

compliance monitoring.

Figure 3 presents photographs of the (5-kW) laboratory-scale

combustor at Stanford and the (50-kW) combustor at NAWC

that served as experimental test beds for the development of

diode-laser sensors for combustion measurements and

control. The combustors were operated on C2

H4, which

served as a waste surrogate, and air for these projects.

For in situ combustion measurements, two independently

operated diode lasers were tuned repetitively over the desired

H2O transitions (vibration-rotation transitions

in the 2ν1 and ν1+ ν3 vibrational bands) to yield

absorption measurements at both 1343 nm and

1392 nm every 300 microseconds. The

individual laser outputs were combined into a

single path using fiber splitters and couplers.

An optical fiber delivered the multi-wavelength

beam to the combustor and a GRIN lens fused

to the end of the fiber collimated the light

through the flowfield. The transmitted multiwavelength light was de-multiplexed (spectrally separated) into the constituent laser

wavelengths by directing the beam onto a

diffraction grating. The beams were diffracted

at angles specific to each wavelength and were

subsequently monitored with photodiodes. The

voltage signals from the photodiodes were then

sent to a personal computer for analysis and to

determine the appropriate feedback control

signal. Temperature was determined from the

ratio of measured peak absorbances at each

wavelength. Water-vapor mole fraction was

determined from the measured absorbance of a single

transition and the gas temperature. The measurements were

recorded, analyzed and an updated feedback control signal

was sent to the actuator at a 3-kHz rate.

For these forced combustion systems, a strong correlation

was established between the magnitude of the measured

temperature oscillations (i.e., Trms) in the combustion region

and the measured concentrations of CO and hydrocarbons in

the exhaust. The expectation, confirmed by experiment, is

that higher temperature oscillations correspond to more

complete combustion (and hence higher energy release and

higher peak temperatures) in the burning vortices of fuel and

air which pass by the monitoring station. Consequently,

adaptive control strategies were developed to maximize the

temperature oscillations by continuously adjusting the

relative phase between the acoustically driven (modulated)

fuel and air flows.

As an example, figure 4a presents measurements and control

of gas temperature in the combustion region of the NAWC

incinerator. Large coherent temperature oscillations at the

forcing frequency (Trms) are prominent when control is

Fig. 2 – Schematic diagram of the setup for measurements in and

control of a multi-stage waste incineration system using diodelaser sensors.
Volume 51, Number 3 & 4 19

applied and the phase

between the fuel and

air flows is optimized. The time

response of the

control strategy is

presented in Figure

4b. The optimum

phase value was

reached in approximately 100 milliseconds, illustrating the

rapid response of the

measurement and

control strategies. It

is the uniquely fast

response of the diodelaser sensors, coupled

with clever algorithms and fast

actuators, which

enables this significant reduction in system control time.

For on-line measurements of combustion emissions, a fast

extractive-sampling method was developed for sensitive

detection of pollutants and unburned fuel (e.g., CO, CO2

,

CH4, C2H4, and C2H2) in the combustor exhaust. For these

measurements, post-combustion exhaust gases were sampled

continuously with a water-cooled stainless-steel probe (4-

mm internal diameter) with several inlet holes (0.5-mm

diameter). The gases were directed into a compact (0.3-liter

volume) multi-pass (folded optical path) cell which consists

of two mirrors with a 20-cm separation and a 33-meter

absorption pathlength. Accurate species concentrations in

sampled gases flowing through the cell were determined

from measured absorbances with response times on the order

of a few seconds. Faster response times may be obtained by

increasing the pump speed and decreasing the length of

tubing from the sampling probe to the measurement cell.

For these exhaust emissions measurements, two fibercoupled distributed feedback (DFB) diode lasers operating

near 1.64 μm and 1.66 μm, and an external cavity diode laser

(ECDL) with a wavelength tuning range of 1.49–1.58 μm

were used. The DFB laser was wavelength-tuned over the

selected CH4

 and C2H4 transitions for measurements of

unburned fuel. The tuning range of the ECDL spanned

Fig. 3 – Photographs of the acoustically forced

combustors at Stanford, left, and NAWC China

Lake, below, for development of diode-laser

sensors for combustion measurements and control.

[Facilities courtesy of Drs. T. Parr and K. Schadow

at NAWC]

Fig. 4 – Measurements and control of temperature in the combustion region of the forced incinerator at NAWC

China Lake. Large coherent temperature oscillations at the desired frequency (Trms) correspond to efficient mixing

and optimized combustion (a). Time response of the control strategy using the diode-laser-based temperature

measurements as input and the relative phase between the acoustic forcing of the fuel and air flows in the

combustor as a control variable (b).

(a) (b)
20 Naval Research Reviews

various CO, CO2, and C2H2 transitions for measurements of

unwanted pollutants. The laser outputs were combined into

one optical fiber using a fiber combiner. The beams were

directed through the multi-pass absorption cell and focused

at the exit onto a photodetector. The transmission signals

were transferred to a personal computer for real-time

determination of species concentrations. The system may be

easily expanded to other important species (e.g., NO) as

lasers become available at the required wavelengths.

Figure 5 illustrates measurements of CO and C2H4 (as

unburned fuel) at various distances from the combustion

region with and without active combustion control. The

decrease in the CO and C2H4 concentrations and the increase

in the rate of destruction of CO and C2H4 with control

validates the forced-combustion scheme for compact

incinerators and illustrates the utility of the laser-sensor

measurement and control strategy.

Measurements in Advanced Propulsion

Systems

Research and development of advanced propulsion systems

is a particularly attractive target for diode-laser sensing

owing to the need for accurate species and temperature data

with high specificity and rapid response in a very hostile

environment. A system of current interest to the Navy is the

pulse detonation engine (PDE) which is under study at the

Naval Postgraduate School (NPS) and elsewhere. Here we

report early work by Stanford researchers to extend and

apply diode-laser sensors for measurements of gas temperature and species concentrations in a PDE facility at NPS.

An air-breathing pulse-detonation engine offers the potential

for providing long range, high cruise speed and subsonic

loiter, and in principle, can provide high performance with

fewer moving parts and lower cost. The rapidly varying

flows generated by pulse detonation engines create challenging environments for sensors due to the high pressures (can

exceed 100 atm) and temperatures (approaching 4000 K) and

the presence of soot and liquid-fuel droplets, and yet detailed

measurements of the type possible with tunable diode lasers

are critically needed to facilitate development of reliable,

efficient propulsion systems. Under these conditions, the

duration of a single pulse is on the order of a millisecond or

less and absorption features are broadened considerably.

Furthermore, at combustion temperatures soot generates

strong blackbody emission that can complicate analysis of

the laser transmission signals, and soot and liquid fuel

droplets can contribute to the extinction of the probe beam

that must be distinguished from molecular absorption.

Density gradients tend to refract beams propagating through

Fig. 5 – Measurements of CO (a) and C2H4 (b), i.e., unburned fuel, using diode-laser sensors in the forced incinerator

at NAWC China Lake with and without active combustion control. [Facility courtesy of Dr. T. Parr and Dr. K.

Schadow]

Fig. 6 – Schematic diagram of the setup for diode-laser

diagnostics (a). Photograph of the PDE facility at NPS

during exhaust (b). [Courtesy of Profs. David Netzer

and Chris Brophy]

(a)

(b)

(a) (b)
Volume 51, Number 3 & 4 21

the flow (often called beam steering) and stress-induced

birefringence at optical ports (windows) can affect (or

modulate) the transmitted laser intensities. Moreover, since

reliable spectroscopic data at elevated temperatures and

pressures are sparse, there is a strong need for basic spectroscopic studies of key species in order to allow quantitative

use of diode laser absorption. Finally, the rapidly changing

conditions in PDE flows demand fast measurements and

high-bandwidth detection to enable the desired temporal

resolution to accurately describe the flowfield.

A schematic diagram of the experimental setup for measurements in PDE flows is shown in Figure 6a. Fuel and oxygen

are injected and ignited at the head end of the tube. The

detonation wave builds and travels toward the tail end of the

tube at approximately the Chapman-Jouget velocity (Vwave ≈

2 km/s) for a detonation wave. Several optical diagnostics of

the type described here can be installed along the length of

the detonation tube; only one is shown for simplicity.

Light from each of four diode lasers operating in a remote

control room is combined (multiplexed) and carried to the

PDE via communications-grade optical fiber. The multiplexed beam is passed through windows in the facility,

demultiplexed by a lens-grating system, and detected by

appropriate photodetectors. In addition to demultiplexing the

four wavelengths, the grating serves to reduce blackbody

emission sensed by the detectors. The entire demultiplexing

and detection system is purged with dry nitrogen gas to

eliminate absorption due to the natural presence of room

water (humidity).

The wavelengths of the four diode lasers were tuned to the

peaks of three water-vapor absorption features (λ1 = 1.34

μm; λ2 = 1.39 μm, and λ3 = 1.80 μm) and to a non-resonant

wavelength (λnr = 1.65 μm) to account for laser extinction

due to soot and liquid fuel droplets. All beams follow the

same optical path through the flow and are subsequently

monitored using individual photodetectors. The collection

system was designed to be insensitive to deflections in the

beams due to refractive effects. H2O is the target species for

this diagnostic since it is a major combustion product and the

strongest absorber within the wavelength range of currently

available diode lasers.

Preliminary data were recorded in a 1.5-inch diameter pulse

detonation tube operating on liquid JP-10 / O2 at the Naval

Postgraduate School. This tube serves as the initiator for a

larger (6”) diameter PDE and has been extensively characterized. A photograph of the facility during detonation and

exhaust of unburned gases is shown in Figure 6b. The

luminosity is predominantly due to soot emission that results

from incomplete combustion of the liquid JP-10 fuel.

Data taken during a single pulse of the NPS detonation tube

are shown in Figure 7a. The pressure trace is from a transducer located at the same axial location as the diode-laser

sensor. The measured absorbance versus time was obtained

for each of the three water transitions (only one wavelength

shown for clarity). Values of gas temperature were determined from two-line ratio of H2

O absorption and from twocolor soot pyrometry (not shown in Fig. 6). The water mole

Fig. 7 – Measurements of pressure, gas temperature, soot temperature, and water vapor recorded in a single pulse

(a); and temperature and H2

O during repetitive pulses at different firing rates using diode-laser sensors and

emission diagnostics in the PDE facility at NPS (b).

(a) (b)
22 Naval Research Reviews

fraction was determined from the measured values of H2O

absorbance, temperature, and pressure.

The TDL absorption diagnostic provides many insights that

were not previously attainable with traditional sensors. For

instance, although the pressure trace suggests that the

flowfield becomes relatively quiet during the first millisecond following the detonation pulse, the diode laser sensor

shows a strong increase in water mole fraction near t=1.8

milliseconds. This increase corresponds to an expansion

wave traveling from the exit to the head end after the

detonation wave leaves the tube.

Figure 7b presents continuous values of temperature and

water mole fraction during pulse detonation rates of 1, 3, and

5 Hz. The advantages to fast, continuous monitoring are

apparent. For instance, the data show that at a 5-Hz firing

rate, a cycle’s post-detonation exhaust is not completed

before the fill and ignition of the following cycle. This effect

may explain why this particular tube was unable to detonate

reliably near or above a 5-Hz rate. Information of this kind

will likely play an important role in the development and

optimization of PDE systems with high repetition rates.

We anticipate that additional measurements with higher

accuracy and including other species (candidates include

OH, CO, and unburned hydrocarbons) will be obtained as the

diagnostic is applied to both the NPS facility and the

Stanford PDE facility currently under construction. In

addition, velocity measurements based on Doppler shifts of

absorption features are planned to enable direct, time

resolved measurements of exhaust velocity and thrust, the

most important parameter for propulsion systems. The

diagnostic promises to enhance detailed modeling of PDE

flows and, eventually, enable active closed-loop control of

the PDE.

Future

Continuing advances in semiconductor and communications

technologies over the next decade will almost certainly add

momentum to the further development of laser-based

sensors, well-suited for real-time monitoring and process

control of combustion and propulsion systems. These

sensors, comprised of semiconductor diode lasers and fiberoptic components, and based on absorption spectroscopy

techniques, already hold significant potential for measurements of key gas parameters inside realistic combustion

facilities and present the opportunity of novel system control

strategies not previously considered. Figure 8 illustrates our

vision of the potential of diode-laser sensors for measurements and control of multiple parameters (including temperature, velocity, concentrations, mass flux and thrust) in

gas turbine aero-propulsion systems. In addition, these same

sensors will serve as important diagnostic tools to allow

improved understanding of chemical and physical processes

that govern the operation of these systems, thus enabling

improved design and development of next-generation

facilities. As new and more broadly tunable diode-laser

sources become available, based on novel designs and

semiconductor materials currently under development,

access to the ultraviolet and infrared spectral regions will be

enabled and this will lead to significant new opportunities for

sensitive measurements of additional species and with

increased sensitivity. The growing use of lasers for sensing

and control of combustion and propulsion systems is thus

almost a certainty.

Fig. 8 – Schematic

diagram of the

setup for real-time

control of aeropropulsion gas

turbines using

diode-laser sensors

for in situ and

extractive sampling

measurements.

Gas parameters

that may be

measured at the

specified locations

are listed in red.
Volume 51, Number 3 & 4 23

Acknowledgments

This work was sponsored by the Office of Naval Research

with Drs. Gabriel Roy and Klaus Schadow as technical

monitors. Partial support from the U.S. Air Force Office of

Scientific Research, Aerospace Sciences and Materials

Directorate, with Dr. J. Tishkoff as technical monitor, is also

gratefully acknowledged.

Biographies

For the biography of Ronald Hanson, see Profiles in

Science on the inside back cover.

Douglas Baer received his undergraduate education in

Engineering Physics at the University of California, Berkeley

(B.S. 1982). Following 18 months as a Research Associate at

the Desert Research Institute, Ben Gurion University, Israel,

he earned graduate degrees in Mechanical Engineering at the

High Temperature Gasdynamics Laboratory (HTGL) at

Stanford University (M.S. 1985, Ph.D. 1993) under the

guidance of Professor Ron Hanson. From 1993 through

1999, Dr. Baer was affiliated with the Department of

Mechanical Engineering at Stanford University as a

postdoctoral Research Scientist in the High Temperature

Gasdynamics Laboratory. Dr. Baer is currently with Informed Diagnostics, Inc.

References

1. R. K. Hanson, P. A. Kuntz and C. H. Kruger, “High-Resolution

Spectroscopy of Combustion Gases Using a Tunable Infrared

Diode Laser,” Applied Optics 16, 2045-2048 (1977).

2. E. C. Rea, Jr., S. Salimian and R. K. Hanson, “Rapid-Tuning

Frequency-Doubled Ring Dye Laser for High Resolution

Absorption Spectroscopy in Shock-Heated Gases,” Applied

Optics 23, 1691-1694 (1984).

3. R.K. Hanson, D.S. Baer, C. Morris, M. Thurber, E. Furlong

and S. Wehe, “Recent Advances in Laser-Based Diagnostics

for Gaseous Flows,” invited plenary paper, SPIE/VSJ,

Yokohama, December 7-9, 1998; also J. Visualization 2, No. 3/

4 (1999).

4. R.J. Cedolin, W.A. Hargus Jr., P.V. Storm, R.K. Hanson and

M.A. Cappelli, “Laser-Induced Fluorescence Study of a Xenon

Hall Thruster,” App. Phys. B. 65, 459-469 (1997).

5. S.D. Wehe, D.S. Baer and R.K. Hanson, “Tunable Diode-Laser

Absorption Measurements of Temperature, Velocity and H2O

in Hypersonic Flows,” paper AIAA-97-3267 at 33rd Joint

Propulsion Conference, July 7-9, 1997, Seattle, WA.

6. D. S. Baer and R. K. Hanson, “Tunable Diode Laser

Absorption Diagnostics for Atmospheric Pressure Plasmas,”

Jour. of Quant. Spectrosc. and Radiative Transfer 47, 455-475

(1992).

7. H.A. Chang, D.S. Baer, and R.K. Hanson, “Semiconductor

laser diagnostics for simultaneous determination of kinetic and

population temperatures in high-enthalpy flows,” 19th

International Symposium on Shock Waves, Marseilles, France

(1993).

8. E.R. Furlong, R.M. Mihalcea, M.E. Webber, D.S. Baer and

R.K. Hanson, “Diode Laser Sensor System for Closed-Loop

Control of a 50-kW Incinerator,” paper AIAA-97-2833 at 33rd

Joint Propulsion Conference, July 7-9, 1997, Seattle, WA.

9. T.P. Parr, E.J. Gutmark, K.J. Wilson, D. M. Hanson-Parr, K.

Yu, R.A. Smith, and K.C. Schadow. Twenty-Sixth Symposium

(International) on Combustion, p. 2033-2055, The Combustion

Institute, Pittsburgh, 1996.

10. T.P. Parr, K.J. Wilson, K. Yu, R.A. Smith, and K.C. Schadow.

15th International Conference on Incineration and Thermal

Treatment Technologies, May 1996, Savannah, GA.

11. E.R. Furlong, D.S. Baer and R.K. Hanson, “Real-Time

Adaptive Combustion Control Using Diode-Laser Absorption

Sensors,” 27th Symp. (International) on Combustion, The

Combustion Institute, pp. 103-111 (1998).
24 Naval Research Reviews
Volume 51, Number 3 & 4 25

T

he applications of thrust-vectoring

are obvious. The AV-8B Harrier is a

combat star, and it achieves its

extraordinary performance through its

thrust-vectoring engines. Short or even

vertical take-off and landing capabilities

are invaluable, especially to aircraft that

operate from ships, or from small,

damaged, or improvised airfields. Thrustvectoring can compensate for loss of

aerodynamic lift, and help a pilot keep

an aircraft under control during stall

conditions.

But thrust-vectoring in operational

craft has always come at a high price in

terms of mechanical complexity and

weight. Gimbaled engines or elaborate

systems of vents and louvers impose all the

usual penalties of moving parts. They are

also heavy, and weight is always at a

premium in aerospace engineering. If

there were a way of altering the direction

of a reaction engine’s thrust without these

mechanical impediments, that would

change everything.

Strykowski and Krothapalli describe

a fluidic approach to thrust-vectoring that

overcomes these disadvantages. It turns

out that there are indeed ways of using

the fluids present in the system to direct

the jets that provide thrust. - J.P.
26 Naval Research Reviews

P. J. Strykowski

University of Minnesota

Minneapolis, Minnesota

Vectoring Thrust Using Confined

Shear Layers

Abstract

A fluidic scheme is described which exploits a confined

countercurrent shear layer to achieve multiaxis thrust vector

response of high speed jets in the absence of moving parts.

Continuous control of the thrust vector angle is demonstrated

in jet exhaust up to Mach 2 for nozzle geometries including

rectangular and axisymmetric cross sections. Studies

conducted jointly between Florida State University and the

University of Minnesota indicate that thrust vector angles up

to at least 20 ̊ can be achieved at slew rates in excess of 180

degrees per second in laboratory scale nozzles. Complementary studies carried out at China Lake Naval Air Warfare

Center on large scale hardware (ten fold laboratory scale)

and at high temperatures (over 3000 ̊F) suggest that scaling

and temperature issues will not preclude the fluidic approach

from working. In both studies, secondary mass flow requirements for control were found to be less than approximately

2% of the primary jet mass. The confined shear layer is

susceptible to both attached and unattached flow regimes, the

nature of which will be discussed. Finally, the performance

of the fluidic approach will be examined in the presence of

external coflow to evaluate the potential of the concept under

flight conditions.

Introduction

The ability to redirect the thrust of an aircraft engine or

rocket exhaust offers several advantages to the aerospace

industry. It provides the potential for a vertical component of

thrust which may be used, especially at low speeds, to

augment the lift force generated by the wings. This allows

A. Krothapalli

Florida A&M University and Florida State University

Tallahassee, Florida

the aircraft to take off in a shorter distance, and ascend at a

higher rate. During landing, vectored thrust can be used to

supplement the lift force generated by the wings, and

approach speeds may be reduced without changing the rate

of descent. The benefits of short take off and landing aircraft

are especially attractive for landing on aircraft-carriers or on

damaged airfields. Traditionally the wings are the sole

mechanism for generating lift. However they have aerodynamic limitations, namely airfoil stall, which causes a

dramatic decrease in airfoil performance and must be

avoided to maintain adequate control of the aircraft. Thrust

vectoring can be used to maintain or re-establish control

under stalled conditions, thus enhancing the overall maneuverability of the plane. In missile applications, multiaxis

thrust vector control could be employed for steering control

at potentially considerably lower expense in terms of vehicle

weight and cost.

There are various methods to vector the exhaust thrust of a

jet engine or missile system. One way is to tilt the entire exit

plane of the nozzle. While this generally requires complex

actuation hardware, it has been implemented successfully on

the Harrier jet for years. For vertical take-off the thrust is

directed almost fully downwards, and once in flight the

nozzle locks in place horizontally and the plane performs in a

conventional manner. Another method which has been

employed on the Lockheed X-31 and the NASA F-18 High

Alpha Research Vehicle, among others, employs hinged

turning vanes downstream of the nozzle exit. The exhaust

impinges on these mechanically actuated vanes and is

deflected. This method has proven to greatly enhance the

maneuverability of the test aircraft, as well as considerably
Volume 51, Number 3 & 4 27

shorten take-off and landing distances. However, these

advantages come at the expense of additional weight and

reduced thrust due to the wall interaction. Similar advantages

can be achieved, in principle, without intricate mechanical

control systems and with improved thrust recovery, by

fluidically vectoring the jet.

The present study concerns a method of fluidic vector

control which employs a confined countercurrent shear layer.

This approach, also known as Counterflow Thrust Vector

Control, or CF-TVC, attempts to combine the continuous

and proportional features of mechanical based systems with

the simplicity of fluidic control, and has recently been

demonstrated experimentally for both subsonic and supersonic jets.1,2 Successful fluidic vectoring must achieve

proportional and non-hysteretic control over a significant

portion of the operation envelop, while minimizing the

secondary mass flow required for actuation. The basic

concept of counterflow thrust vectoring can best be illustrated by referring to the sketch in Figure 1a, where the side

view of the short dimension of a rectangular jet is shown.

The primary jet exhausts from the nozzle between symmetric

curved surfaces called collars, placed on either side of the

primary stream. To achieve upward thrust vectoring at an

angle δv

, a secondary counterflow must be established in the

upper shear layer of the jet. This can be accomplished by

applying suction to the plenum chamber located between the

upper curved surface and the primary nozzle. The action of

counterflow in the upper shear layer gives rise to asymmetric

entrainment and a cross-stream pressure gradient sufficient to

vector the jet.

The ability of the jet to vector in the presence of a stationary

surface has been commonly referred to as the Coanda effect.3

As a nominally two-dimensional jet issues from the nozzle,

the shear between the jet and the surrounding quiescent fluid

gives rise to lateral momentum transport which is accentuated by the turbulent mixing of the shear layer. This process

leads to momentum mixing between the two fluids and is

commonly referred to simply as entrainment. Because the jet

is gathering fluid through entrainment, its boundaries diffuse

outward and, in the process, weaken the jet’s ability to

further entrain the surrounding fluid. Figure 1b depicts the

two-dimensional jet with a curved surface of radius R

aligned tangentially to the edge of the nozzle exit. The fluid

entrained along the lower collar surface by the pumping

action of the shear layer is constrained, giving rise to

subatmospheric pressures in that region. This differential

pressure draws the jet toward the curved surface. It is this

change in the pressure field within the vicinity of a solid

surface that is known as the Coanda effect. (First described

by Young,3

 the Coanda effect is named for the French

engineer who patented the “effect” in 1932.) If the entrainment rate is sufficiently high, the pressure continues to drop

until the jet becomes attached to the surface. It is this

attachment process that is most often associated with the

Coanda effect. However, attachment is not a necessary

condition for the Coanda effect, but rather a consequence of

it.

Once the flow becomes attached, the jet continues to entrain

fluid between itself and the curved surface leading to further

wrapping of the jet around the surface. This process continues until the shear layer structures have weakened sufficiently and can no longer entrain the fluid necessary to force

the jet against the surface. The point of separation of the jet

from the surface is therefore dependent on the nature of the

shear layer structures and their ability to entrain ambient

fluid. Once the jet has established itself in the configuration

shown in Figure 1b under steady conditions, it remains

stable, i.e., perturbations to this flow result in the jet reestablishing itself to the condition prior to the perturbation. While

the attached portion of the jet is obviously no longer entraining fluid against the curved wall, the pressure there is kept

low as demanded by the equations of motion. Under these

conditions, the curved jet can be described by two-dimensional cylindrical coordinates r and θ which, respectively, are

the local radius of the curved jet and its angular position. The

corresponding local velocities in the r and θ directions are

then ur

 and uθ. If we assume that the flow is nominally

parallel in the vicinity of the collar, namely that uθ = uθ

(r)

and ur = 0, we obtain for a steady, inviscid and incompressible fluid the following relationship:

Fig. 1 – (a) Schematic of fluidic thrust vectoring nozzle

when counterflow is activated in the upper shear layer

of the jet; (b) jet attachment to a curved surface caused

by natural entrainment commonly referred to as the

Coanda effect.
28 Naval Research Reviews

∂

∂

P ρ θ

r

u

r =

2

(1)

Consequently, the pressure gradient in the r-direction is seen

to increase with the square of the velocity. Therefore, the

pressure will increase from some minimum value at the

curved surface to atmospheric pressure near the outer

boundary of the jet, ensuring that the centripetal acceleration

is balanced by the pressure field.

The distribution of vacuum pressure maintained on the

curved surface will determine the net forces acting on the

surface. However, one of the inherent difficulties of controlling the Coanda effect as described above, is its passive

nature. For a given jet momentum flux and nozzle-collar

geometry, a fixed pressure distribution will be created on the

collar, and therefore a fixed side force will be generated.

Hence, the Coanda effect leads to hysteretic jet attachment

and bistable operation, which could be used for thrust vector

control if the Coanda surface were designed to be movable.

But this configuration would defeat the initial attractiveness

of fluidic control using stationary hardware, and hence why

the passive Coanda effect as shown in Figure 1b was not

examined as a likely candidate for aircraft and missile

control applications. However, if the jet curvature can be

controlled by, for instance, manipulating jet entrainment, the

cross stream pressure field can be altered continuously, and

thereby generate the side forces necessary to vector thrust

and avoid hysteresis.

Strykowski, Krothapalli & Jendoubi4

 demonstrated that

counterflow, when applied to the periphery of an

axisymmetric jet, greatly enhanced the entrainment characteristics of the shear layer. By applying counterflow on only

one side of the jet, the excited shear layer entrains mass more

effectively than the opposing free shear layer. This imbalance

in entrainment causes a cross-stream pressure gradient,

which deflects the jet in the direction of the applied

counterflow, resulting in a vectored primary flow. Van der

Veer & Strykowski1

 were able to proportionally pitch vector

control a subsonic rectangular jet up to Mach 0.5 at angles up

to 20°. Strykowski, Krothapalli & Forliti2 achieved similar

pitch vectoring performance in a supersonic jet at Mach 2.

No moving parts are directly required to steer the jet when

CF-TVC is employed. Consequently, reliability can, in

principle, be greatly enhanced. Furthermore, there are no

surfaces in direct contact with the high temperature, high

velocity exhaust gases. This may eliminate the need for

expensive materials or exotic cooling schemes, which would

be necessary with turning vanes. Because of the potential for

minimal weight addition to the vehicle, CF-TVC systems

may be retrofitted to an existing aircraft’s engine without

significant structural alteration of the airframe. Similar

benefits are expected for the design of new exhaust nozzle

concepts for both aircraft and missile applications, namely a

reliable, robust system with a relatively low initial cost, and

minimized maintenance costs.

Facilities

The majority of measurements were conducted in the blowdown compressed air facility of the Fluid Mechanics

Laboratory located at Florida State University. (The complementary studies at China Lake Naval Air Warfare Center will

only be summarized here.) The facility is driven by a highdisplacement reciprocating air compressor which is capable

of supplying air at a maximum storage pressure of 160 bars.

Large storage tanks provide a total capacity of 10 m3 and are

capable of driving the Mach 2 primary jet flow examined in

this study continuously for up to 30 minutes. The supply air

can also be heated by being passed through an array of

resistive tank heaters having a maximum power output of

450 KW and capable of achieving stagnation temperatures in

excess of 750 Kelvin. Visualization of the jets was achieved

by creating laser sheet images of the vectored plane along the

jet axis using frequency doubled Nd:YAG pulsed lasers. Fine

condensation ice particles formed in the mixing region

between the dry cold air of the primary jet and the moist

ambient air scatter the laser light thus rendering the shear

layers visible.

For the pitch-axis vectoring study, the blow-down facility

was fitted with a rectangular nozzle having an exit aspect

ratio of 4:1 (52 mm by 13 mm ). The Mach 2 axisymmetric

nozzle used for multiaxis thrust vectoring was designed to

have the same minimum throat and exit plane cross-sectional

area as the rectangular nozzle to ensured that, for the same

stagnation pressure and temperature, the flow conditions for

both jets were nominally the same. Contours of both nozzles

were generated by a method of characteristics for a design

Mach number of two and were run at their design pressure

ratios. The jets were operated at stagnation temperatures

between 300 K and approximately 700 K, corresponding to

exit plane Reynolds numbers from 0.4 to 1.2 x 106

.

Pitch Vector Control

To achieve thrust vector control, a secondary counterflowing

stream needs to be established along the outer surface of one

of the jet shear layers as shown previously in Figure 1a. This

secondary stream is created by connecting a vacuum pump

and manifold to a cavity placed along the periphery of the

jet. To assure that the counterflowing stream acts along the

appropriate shear layer of the jet, suction chambers must be

created, the nature of which will depend on the nozzle

geometry under consideration. This can be achieved quite

easily in the rectangular geometry by the positioning of side

plates attached to the ends of the collars, effectively isolating

the secondary flow to either the upper or lower shear layer.

Similar partitioning can be employed in axisymmetric jets to

achieve multiaxis thrust vector control as will be discussed

later.
Volume 51, Number 3 & 4 29

To demonstrate the CF-TVC concept for the pitch vectoring

of a rectangular jet, a base study was conducted with flow

visualization using planar laser scattering. The primary jet

was unheated having a stagnation temperature of 300K,

resulting in a low static temperature after expansion to Mach

2. Under these conditions, the entrainment of relatively moist

air into the shear layer of the cold jet gives rise to ice crystal

formation in the mixing region and thereby illuminates the

jet shear layers during vectoring. The series of photographs

in Figure 2 were taken over the streamwise distance of

approximately 7H to 23H , and were captured by illuminating the jet with a 10-nsec exposure of the laser sheet positioned along the jet axis. When the vacuum system,

connected to the upper collar, is deactivated and air is

allowed to entrain freely into both the upper and lower shear

layers, the jet exhausts along its geometrical axis as shown in

Figure 2a. Under these conditions the collar arrangement

operates analogously to a jet ejector configuration, albeit

inefficiently due to the rapid divergence of the collar walls.

When the vacuum system is activated creating secondary

counterflow in the gap between the jet and the collar, thrust

vectoring can be achieved as shown by the photographs in

Figures 2b,c. The static pressure measured on the upper

collar surface in the jet exit plane (essentially the secondary

plenum pressure PB) was reduced from -0.061 bar in Figure

2b, to -0.29 bar in Figure 2c, corresponding to vector angles

estimated from the photographs and corroborated using

direct thrust measurements of 6 ̊ and 16 ̊, respectively. The

jet could be vectored toward either the upper or lower collar

by activating the appropriate vacuum system.

Due to the intense mixing which occurs in the shear layer of

the vectored jet, it is not obvious whether the secondary flow

drawn into the vacuum system originates primarily from

ambient fluid or from the jet itself. To obtain a more complete physical understanding of the countercurrent flowfield

in the vicinity of the collar, the mean velocity was captured

using particle image velocimetry, a non-intrusive technique

requiring that the primary and secondary streams be seeded

with submicron particles and illuminated using a pulsed

laser.2

 The mean velocity-vector field and corresponding

isovelocity contours obtained for a Mach 2 jet vectored at δv

= 8 ̊ are shown in Figure 3. A close examination of the

velocity vectors reveals that the maximum countercurrent

velocity occurs near the midpoint of the collar, at x/H ≈ 4 in

Figure 3, and corresponds to nearly 30% of the primary jet

velocity. These high secondary velocities are due to both the

vacuum system as well as the natural entrainment characteristics of the primary jet. What is perhaps most interesting is

that the majority of the counterflow is eventually entrained

by the jet and carried downstream, resulting in relatively

small amounts of secondary flow entering the vacuum

system. The maximum amount of mass drawn into the

secondary plenum is less than 2% of the primary jet mass

flow rate.

Casting the experimental thrust performance data into a

common format, which can be used for nozzle design,

requires an examination of the parameters influencing jet

curvature in the proximity of the collar surface. For purposes

Fig. 2 – Light scattering

from the jet centerplane

showing thrust vectoring

at (a) 0 ̊, (b) 6 ̊, and (c) 16 ̊

using a collar geometry of

G/H = 0.38, L/H = 6.9, and

R/H = 15.7.

Fig. 3 – Velocity vector field and isovelocity contours

obtained using particle image velocimetry for a jet at

Mach 2 for G/H = 0.38, L/H = 6.9 and R/H = 15.7.
30 Naval Research Reviews

of discussion, we will considering the pitch vectoring of a

rectangular jet as shown schematically in Figure 1a and

visually in Figure 2, and subsequently extend the scaling to

other nozzle geometries, in particular the axisymmetric

nozzle. The exit vector angle, δv, depends on the length of the

collar surface, L, the forward momentum of the jet per unit

depth, ρ1U2

1

H, and the pressure distribution on the collar

surface, where the subscript 1 indicates mean quantities in

the jet exit plane. Experimental work1,2 has shown that the

pressure distribution on the collar surface is nominally

uniform over the forward half of the collar where the

majority of the side force is taken up, and may be proportionally controlled by varying the secondary plenum pressure PB.

The relationship between the plenum pressure and vector

angle is required for nozzle design, and has been examined

over a wide range of conditions for both subsonic2 and

supersonic3,5 jets. (Note that the scaling arguments and

experimental results presented here are relevant for jets

operated near their design pressure ratios only.)

The two-dimensional momentum equation in cylindrical

coordinates provides insight into the physical phenomena

affecting the relationship between δv

 with PB. Recognizing

that viscous terms are small compared to inertial effects in a

free jet, and assuming that the flow is nominally parallel in

the streamwise direction the momentum equation can be

simplified to the expression given previously in Eqn. (1). To

develop a scaling equation for CF-TVC performance, it is

assumed that the jet uniformly arcs at a constant radius of

curvature, R. This assumption is fairly accurate as long as R

is significantly larger than the jet dimension H; hence, the

analysis applies to small vector angles.

Integrating from one side of the jet to the other, the right

hand side of Eqn. (1) represents the streamwise momentum.

Since streamwise momentum is conserved in a free jet flow,

this quantity is equal to the momentum flow in the jet exit

plane per unit depth, ρ1

U2

1

H. The integral on the left-hand

side is known from the boundary conditions. On the radially

outward shear layer, the pressure is equal to that of the

ambient, P∞, on the inner shear layer, the pressure is approximately that of the secondary plenum PB. Thus we obtain:

P PB

U H

∞ R − = ρ1 1

2

(2)

Upon reaching the streamwise extent of the collar, L, the

counterflow ceases and with it the mechanism for sustaining

the cross-jet pressure gradient. As a result, the jet exits

straight into the atmosphere, at an angle δv

, as seen in Figure

1a. The jet curvature can be related to the exit vector angle

and the length of the collar using the expression:

R L

v

= sinδ (3)

Combining the last two expressions and solving for the

vector angle in terms of the control pressure, we have:

sin

( )

δ

ρ v

P PB L

U H = ∞ −

1 1

2 (4)

Experimental observations of pitch vectoring in rectangular

jets, for both subsonic and supersonic conditions show that

this expression holds true, to a good approximation, over a

wide range of flow conditions and collar parametrics as

shown in Figure 4. The data presented in the figure include a

variety of studies of supersonic flow at Mach 2, where nozzle

gap G, collar length L, radius of curvature R, and jet stagnation temperature have been varied; thrust vector angles were

determined using a six-component thrust stand and were

corroborated with control volume analysis. The pitch vector

data, indicated by the solid symbols enclosed within the

hatched region, demonstrate a nearly linear relationship

between normalized plenum pressure and jet response for

primary stream stagnation pressures held fixed at the design

pressure ratio for Mach 2. Under these conditions, weak

shocks are present due to the asymmetric pressure distribution in the jet exit plane caused by underexpansion in the

direction of secondary counterflow.

The scaling relationship observed in Figure 4 can be used to

design nozzle-collar systems in other geometries by considering that the expression on the right hand side of Eqn. (4) is

effectively the ratio of the side forces acting on the collar

surface (P∞ – PB)Aside to the axial force imposed by the jet

ρ1U2

1

Ajet, where the ratio Aside/Ajet can be simplified to L/H for

the rectangular geometry. Hence, the basic scaling law we

want to consider for more complicated geometries is given

Fig. 4 – Counterflow thrust vectoring performance at

Mach 2 for rectangular and axisymmetric nozzles.
Volume 51, Number 3 & 4 31

by Eqn. (5), valid for nozzles operated near their design

pressure ratios.

δ

ρ v fcn

P PB Aside

U Ajet

= ∞ − 



















( )

1 1

2 (5)

Multiaxis Thrust Vector Control

Extending the CF-TVC concept from pitch vectoring to

multiaxis operation requires a collar design which effectively

directs counterflow along the jet shear layer. Since the

vacuum source alone cannot direct the flow, the collar must

be partitioned to created chambers through which the

secondary flow can pass and thereby establish a countercurrent shear layer. Figure 5 indicates the partitioning strategy

employed for multiaxis vector control of an axisymmetric

Mach 2 nozzle. When the vacuum source is connected to a

particular chamber, a secondary counterflow is developed —

bounded by the collar surface and a segment of the jet shear

layer. This, in principle, causes the primary jet to vector in a

direction normal to the shear layer along which counterflow

is established. A single vacuum source was connected to the

chambers via control valves allowing independent vectoring

in any of the six directions as desired.

Instantaneous planar laser scattering images showing the

cross-sectional view of the unvectored and vectored

axisymmetric jet at x/D ≈ 2.5 are shown in Figure 6; note

that the camera is positioned off-axis to avoid interaction

with the high-speed jet, creating the observed oblique view.

A close examination of the images reveals that not only does

the jet move toward the chamber with counterflow, but the jet

shape is altered considerably. We believe this is primarily

due to the difference in the mixing rates of the countercurrent

shear layer (upper shear layer in Figure 6b) and the shear

layer regions isolated from secondary counterflow. For

instance, if we compare the top portion of the jet shear layer

in Figure 6b with the lower portion, it is apparent that the

upper layer has diffused over a significantly larger area. This

observations may allow the CF-TVC concept to be used not

only for vector control but for mixing enhancement as well,

which may be desirable for reducing jet noise or thermal

signature. The crosshairs provide a frame of reference to

compare the unvectored and vectored flow. There is a slight

vertical shift in the lower shear layer, but the upper layer is

considerably displaced. Thrust stand measurements indicate

that the effective jet deflection is ~10 ̊, which is difficult to

detect in the oblique view of Figure 6 due to the relatively

small streamwise distance where the images were captured.

A summary of the axisymmetric jet vectoring as a function

of the static pressure parameter is provided in Figure 4 to

allow comparison to the pitch vectoring data; multiaxis thrust

vector data are shown for studies conducted at Florida State

University/University of Minnesota and China Lake.

Although the axisymmetric nozzle data show reasonably

linear behavior, the pitch vectoring results obtained using the

rectangular nozzle appear to be slightly more efficient,

namely larger vector angles can be achieved with lower

vacuum pressures developed in the secondary plenum. We

Fig. 5 – Schematic of the end view of the axisymmetric

CF-TVC nozzle showing six-sector partitioning.

Fig. 6 – Planar laser

scattering images

of a Mach 2

axisymmetric jet

cross section at x/D

≈ 2.5 for (a)

unvectored jet and

(b) jet vectored into

uppermost sector

at δv = 10 ̊.
32 Naval Research Reviews

believe the primary reason for these differences is due to

leakage between adjacent chambers of the collar assemblies

for the axisymmetric arrangement. This was determined by

monitoring the static pressure distributions in the chambers

adjacent to the primary chamber where vacuum is applied.

These measurements indicated that counterflow is not

completely isolated from other parts of the shear layer,

resulting in a degradation of performance due to the transverse pressure gradient across the shear layer. It is also

difficult to determine the appropriate side area Aside which is

used to develop the scaling in Figure 4; a simple projection

of the active sector was used when plotting the data. The

maximum vector angle of 10 ̊ achieved for the axisymmetric

study compared to nearly 20 ̊ for the rectangular jet, is not

due to an inherent limitation of the concept, but rather caused

by the smaller side area of the axisymmetric collar compared

to the rectangular collar when acted upon by the same

available vacuum pressure (P∞ - PB). Although these threedimensional effects will undoubtedly influence the collar

performance, we believe that they can be minimized by

improvements in collar design.

Performance Considerations

Performance of the multiaxis CF-TVC system was also

documented in terms of secondary mass flow rate, jet

dynamic response and hysteretic behavior. The mass flow

drawn through the secondary vacuum system was determined

to be consistently less than 2% of the primary jet mass for

the unheated primary flow (stagnation temperature of 300

K), and significantly lower when the primary jet was heated.

In the FSU/UM experiments described above, the primary

mass flow rate was between approximately 0.5 and 0.75 kg/

sec, for hot and cold jet conditions respectively. The secondary mass flow rate was below 0.015 kg/sec for all conditions

examined. Dynamic performance of the multiaxis system

was also evaluated for the axisymmetric configuration. A

solenoid valve was used to activate and deactivate

counterflow to a desired chamber and two different measurements were used to determine the dynamic response of the

jet. In the first approach, the planar laser scattering images

were recorded on video during this transient process and a

time stamp was superimposed on this record. The video

record shows the jet being deflected from δv

 = 0 ̊ to 9 ̊ and

vice versa in significantly less than a tenth of a second, the

lowest subdivision on the time stamp. This places a lower

limit of 90 degrees/sec on the jet deflection rate. A second

method was then used to determine the jet dynamic response

more precisely. The unsteady plenum pressure PB on the

collar surface was monitored using a fast response pressure

transducer (minimum flat frequency response of 40 kHz)

while counterflow was cycled on and off. These measurements indicate that the jet could be deflected from zero to 9

degrees in less than 50 ms, resulting in jet vectoring rates of

at least 180 degrees/sec.

Due to the entrainment differential, a small degree of jet

turning is possible, in principle, without a collar surface,

however it is the addition of an extended collar surface of

length L that makes CF-TVC a viable technology. First, it

channels the secondary flow in parallel with the primary

flow, effectively inducing a countercurrent mixing layer over

a finite extent of the jet column. Second, it restricts the

natural entrainment of the jet, intensifying the cross-stream

pressure gradient. Finally, it gives the pressure forces a

surface over which to act. A longer collar has more surface

area for the pressure forces to act on thus a smaller pressure

differential will impart an equal transverse force on the

nozzle-collar assembly. In other words, with a longer collar,

the same vector angle is achievable with a smaller amount of

vacuum from the secondary flow system. The net effect is a

reduction in pumping power losses.

While the collar is vital to the efficiency of the CF-TVC

system, it poses a potential problem that can disrupt the

continuity of the operating curve. This condition, a result of

the bistable interaction between a free jet and a wall, is one

in which the jet attaches to, and reaches a stable equilibrium

on, the wall. It has been observed, under certain operating

conditions by previous CF-TVC investigators,1,2 that the jet

attaches to the collar surface during vectoring. This is

unacceptable from a design standpoint as, in a flight situation, this may cause a loss of control. Furthermore, to release

the jet from the collar, the differential control pressure must

be reduced beyond that which was required to cause attachment. This hysteretic behavior makes a continuous vectorcontrol system very difficult (if not impossible) to

implement. Fortunately, this situation can be prevented in

many instances by correctly tailoring the geometry of the

collar.6

The important thing to bear in mind when designing a CFTVC system is that wall attachment can only occur if an

equilibrium can be sustained as shown in Figure 1b. This

means that the entrainment mechanisms within the jet must

be able to sustain the low pressure necessary to hold the jet

attached to the wall. The shorter the collar, the sharper the jet

must turn in order to attach; thus a lower plenum pressure

will be required. Furthermore, with a short collar, the shear

layer has less contact with the secondary stream, making it

more difficult for the pumping mechanism within the jet to

generate the low pressure required to hold itself to the wall.

Consequently, if the collar is sufficiently short, attachment

will not occur.

There is a unique equilibrium location where the attaching

streamline intersects the collar. This location depends on the

gap width, the mixing dynamics of the shear layer, and the

amount of secondary flow leaving (or entering) the recirculation zone. A large gap places the collar surface farther away

from the jet, resulting in a longer attachment length. Enhanced shear layer growth rates mean the jet is more effective in sustaining low pressures along the collar surface.

Likewise, counterflow being drawn from this region by a
Volume 51, Number 3 & 4 33

secondary pump assists the jet in sustaining the low pressure

necessary to hold itself to the wall. These two latter conditions enable the jet to turn with a smaller radius of curvature,

resulting in a shorter attachment length. If the collar is longer

than this equilibrium attachment length, Latt, the jet will

merely attach at the designated location, and follow the

contour of the wall until boundary layer separation occurs.

However, if the collar length is shorter than the equilibrium

attachment length, i.e. L < Latt, the jet will not attach to the

wall. When designing a collar, it is important to be able to

estimate the attachment length based on the intended

operating conditions. Ultimately, designing a CF-TVC

system for aircraft or missile propulsion requires that a collar

geometry be found which is capable of achieving the

required thrust vector angle, δv, while minimizing external

drag, secondary mass flow pumping demands, and provide

attachment-free operation over the entire operating domain

of the vehicle.

Flight Effects

The generally favorable characteristics of CF-TVC performance as described above are encouraging, however successful system integration of the concept requires at a minimum

that the technology can be implemented during actual flight

conditions. To examine this aspect of performance, the

influence of a coflowing external stream was imposed on a

rectangular jet (aspect ratio of 4:1) operated at a Mach

number of 1.4; external coflow Mach numbers were examined between 0.3 and 0.7. A schematic of the CF-TVC

system and coflow hardware is shown in Figure 7. The

coflow was established by placing a rectangular duct around,

and extending downstream of the nozzle-collar hardware.

The duct was equipped with glass windows to allow optical

access for flow visualization; surface pressures were used

together with momentum flux measurements to evaluate the

effective thrust vector angle using a control volume analysis.

A summary of the CF-TVC performance of a rectangular jet

for various counterflow and external coflow levels is

provided in Figure 8 together with the pitch vector control

presented in Figure 4 for comparison. Again PB represents

the secondary plenum pressure measured on the collar

surface in the jet exit plane. At first glance it appears that

thrust vector performance with external coflow is very

similar to the behavior without coflow, in that the jet vector

angle increases nearly linearly with increasing vacuum

pressure. However, a closer examination reveals that, albeit

small, there is a systematic degradation in the performance

with coflow. For a given differential pressure across the

collar, the jet response indicated by δv decreases as the

coflow Mach number increases. This effect can be understood if we consider the effect of vectoring in one direction,

for instance up, on the flow regime on the opposite side, in

this case the region below the curved primary jet. In the

configuration used in our experiments, when counterflow is

actuated in the upper shear layer, there is no flow through the

lower gap between the jet and lower collar. But the lower

shear layer of the primary jet and the upper shear layer of the

lower coflowing stream both entrain fluid from the region

bounded by these two shear layers creating a low pressure

zone — essentially a wake — on the lower side of the

primary jet. This in turn reduces the side force on the nozzlecollar assembly and the effective angle δv

. As the entrainment

rate increases due to increased coflow Mach number, so does

the wake effect and the detrimental influence on performance. However, at higher thrust vector angles the pumping

action of the primary jet has a reduced impact on the surface

pressures of the lower collar for a fixed coflow Mach

number. Hence, as δv increases the difference between coflow

and no-coflow experiments should diminish, which is the

trend seen in Figure 8.

Fig. 8 – Thrust vector performance of a rectangular jet

operated in the presence of an external coflowing

stream.

Fig. 7 – Side view of a Mach 1.4 rectangular jet

positioned in a coflow duct to simulate forward flight

effects.
34 Naval Research Reviews

The degradation in CF-TVC performance can be alleviated,

in principle, if atmospheric pressure is maintained in the

regions not actively experiencing counterflow. This could be

achieved by ducting some of the freestream flow into the

inactive secondary plenum, and thereby “filling” this region

to reduce the wake effect. Although the actual hardware used

to accomplish this will depend on the overall vehicle

configuration, we believe that the effect of coflow on CFTVC performance can be implemented using this approach.

Closing Remarks

We have provided a brief overview of the issues which must

be considered in the design of a thrust vector system based

on counterflow. It has been shown that through CF-TVC it is

possible to vector thrust in multiaxes in a continuous fashion

using a fluidic approach. The performance of the nozzle

system is relatively insensitive to jet stagnation temperature,

at least over the range studied to date (300 to 1930 K), and

the thrust vector angle is approximately a linear function of

the static pressure developed in the counterflowing stream.

Although many issues need to be examined further, such as

the presence of shock waves within the collar region for jets

operated off-design, and further study at high temperatures,

the inherent simplicity of this method makes it a promising

concept for vectoring thrust.

Acknowledgments

The authors would like to thank the Office of Naval Research

for support of this work and the guidance we have received

from technical monitor Dr. Gabriel Roy. We would also like

to acknowledge our colleagues Professor F. Alvi, D. Forliti,

G. Schmid, and M. Van der Veer.

References

1. Van der Veer, M.R. and Strykowski, P.J., “Counterflow thrust

vector control of subsonic jets: continuous and bistable

regimes,” J. Prop. & Power, Vol. 13, No. 3, 1997, pp. 412-420.

2. Strykowski, P.J., Krothapalli, A. and Forliti, D.J.,

“Counterflow thrust vectoring of supersonic jets,” AIAA J., Vol.

34, No. 11, 1996, p. 2306-2314.

3. Young, T., “Outlines of experiments and inquiries respecting

sound and light,” Lecture to the Royal Society, Jan. 1800; also

Journal of the Royal Aeronautical Society, Vol. 61, 1957, p.

157.

4. Strykowski, P.J., Krothapalli, A. and Jendoubi, S., “The effect

of counterflow on the development of compressible shear

layers,” J. Fluid Mech., Vol. 308, Feb. 1996, p. 63-96.

5. Washington, D.M., Alvi, F.S., Strykowski, P.J. & Krothapalli,

A., “Multiaxis fluidic thrust vector control of a supersonic jet

using counterflow,” AIAA J., Vol. 34, No. 8, 1996, p. 1734-

1736.

6. Schmid, G.F., “Design and optimization of a counterflow

thrust vectoring system,” M.S. Thesis, University of

Minnesota, Minneapolis, June 1996.
Volume 51, Number 3 & 4 35

Designing gas turbines has long depended on a lot of empirical work—

“cut and try,” in engineers’ parlance. But now we’re reaching a point in

computational power and modeling sophistication where that may all change.

Advances in computational fluid dynamics and related areas of

computational combustion mechanics are revolutionizing the field of gas turbine

design. Engineers have successfully demonstrated simple models—semianalytical mechanistic models—for various combustors. They are now ready to

validate advanced models for combustion design.

Hukam Mongia describes how the state of aero-thermal design has

advanced over the last quarter century, and what we can expect for the future

as engineers continue to push the frontiers of combustion modeling. -J.P.
36 Naval Research Reviews

Aero-Thermal Design Methodology

for Gas Turbine Combustion

Systems: Future Direction

Hukam C. Mongia*

Combustion Center of Excellence

GE Aircraft Engines, Cincinnati, Ohio

GE Aircraft Engines (GEAE) has introduced several new

aeropropulsion gas turbine engines with improved combustion systems since 1994 including the CFM56 dual annular

combustor (DAC), the CF6-80 low-emissions single annular

combustor (LEC), the GE90 DACII, in addition to the dry

low emissions (DLE) for the LM1600, LM2500 and the

LM6000. Development and validation of advanced turbulence and radiation models as part of the National Combustor

Code (NCC), is now pursued in collaboration with NASA.

These activities pave the way for combustion design process

that will use computational combustion dynamics-based

methods unlike the most conventional “cut-and-try” design

practice.

Introduction

The empirical/analytical gas turbine combustor design

methodology has progressively improved over the last 25

years in regard to: a) the design process, b) the aero-thermal

combustion models, and c) simulations including accurate

representation of the combustion geometry, numerics and

other critical steps. Since 1986, General Electric (GE) has

developed two techniques (namely, hybrid modeling and

anchoring methodology) for getting quantitatively good

agreement with gas turbine combustor data in regard to

combustion efficiency, HC, CO, NO, lean blowout, burner

exit temperature profile, pattern factor, and liner wall

temperatures.

The empirical/analytical combustion design methodology

has been used in the designs of numerous combustors1,2.

Simple models, termed as semi-analytical mechanistic

models (SAM), have been formulated and their feasibility

demonstrated for both the diffusion-flame and premix

combustors3,4. Anchored Computational Combustion

Dynamics (CCD) based approach introduced in 1996, that

has become a standard practice in the combustion design

procedure5. With recent advances in grid generation, computational speed, and combustion models, validation of

advanced models as the next logical step of combustion

design tools evolution, has been started.

Advanced Combustion Code

Having established Generation 1 SAM and anchored CCD

methodology, several activities have been initiated recently

to bench mark and/or improve:

a) CFD capability of the design code

b) Baseline turbulent combustion models based on eddybreakup (EDC) and scalar transport PDF

c) Turbulence models relevant to gas turbine combustion

d) Radiation model

e) Spray model

f) Laminar flamelet modeling with reduced and skeletal

mechanisms

g) Large eddy simulation (LES).

The following paragraphs give brief description of the

activities in each of these eight areas.

An Advanced Combustion Code (ACC) is currently being

developed to further advance the use of Computational

Combustion Dynamics (CCD). ACC has complex geometry

representation capability via unstructured single block grids, *Copyright 1998 by the author. Published by ONR with permission.
Volume 51, Number 3 & 4 37

and comprises of a number of advanced combustion/spray/

turbulence models that are being developed in a modularized

fashion. The ACC solver is based on a fully implicit pressure

correction algorithm with co-located grids. An overview of

an early version of ACC is given by Tolpadi et al.

6

 This code

allows a more accurate representation of staged combustion

systems including DLE.

Since the first use of CCD in combustion design 1, the need

to improving turbulence models for gas turbine combustion

was emphasized. In spite of concerted effort by several

turbulence modeling groups over the last 25 years, the

predictive accuracy for turbulent combustion field has not

improved. Even after having expended a considerable effort

on carefully measuring and modeling isolated sub-component flow-fields relevant to gas turbine combustors, it was

not possible to improve accuracy of turbulent combustion

models. For example, refer to the numerical calculations of

the jet diffusion flame performed by Tolpadi et al.6 using the

eddy breakup, eddy dissipation and the scalar PDF transport

combustion models with the standard k-ε model. Even

though one can offer several good reasons for the failure of

the models to achieve acceptable level of quantitative

agreement for such a simple test case, it is nevertheless a

discouraging state of the turbulent combustion modeling.

Encouraged by recent advances in turbulence modeling7-9,

this author’s group recently started to reassess the strategy

for improving turbulence modeling for the following three

areas of interest in gas turbine combustors.

There is an urgent need to predict accurately the pressure

recovery and total pressure loss distribution in gas turbine

combustor diffusion system so that the designer can calculate

the distribution of airflow, static and total pressure distribution around the combustor dome and inner and outer liners.

Karki et al.

10 demonstrated that 3-D diffusion system

calculations can be done if a) one had an accurate turbulence

model for flows involving adverse pressure gradients; and b)

sufficiently fine grids were used for simulating production

gas turbine combustor diffusion system.

Nikjoo and Mongia11 recently evaluated the performance of

the standard k-ε model, a nonlinear Yang and Shih k-ε model

(YSM), k-ω model (high and low-Reynolds number formulation), and shear stress transport (SST) model for welldefined experimental data for adverse pressure gradients

(involving separated boundary layer) documented by

Driver12. Fine-grid calculations were done for the region

between the first measuring station at x/R0=-6.531, and the

exit plane at x/R0=11.4 where R0=70mm is the cylinder

radius and the boundary layer separation takes place at X =

0.0.

Figure 1 shows comparison between between the measured

CP and predicted CP by the standard k-ε model, YSM, a

nonlinear k-ε model, the low-Reynolds k-ω model, and SST

model. It shows that a significantly improved Cp correlation

is achieved by the SST model. All the models predicted

similar Cp

 in the first half of the flow in the region with

adverse pressure gradient. However in the recovery region,

the performance of the k-ω and YSM models are not

satisfactory. The SST model’s predictions are in very good

agreement with data in both the front and recovery regions.

Fig. 1 – Comparison between predicted (by Nikjoo and

Mongia, 1999) and measured (by Driver, 1991) CP

distribution.

Fig. 2 – Comparison between predicted (by Nikjoo and

Mongia, 1999) and measured (by Driver, 1991 ) axial

velocity profiles in a separating boundary layer.
38 Naval Research Reviews

In other words, when when the eddy viscosity is properly

corrected for the adverse pressure gradients (as done in the

SST model), it is possible to correctly predict Cp distribution

in a diffusion system, a pre-requisite for claiming to have an

acceptable diffusion system model.

Comparisons of the measured and predicted mean axial

velocity component by the SST, k-ω model, and YSM model

are shown in Figure 2 for different axial stations. The axial

velocity component calculated by the SST model closely

resembles the experimental data. There are substantial

differences between the measured profiles and the predictions with the k-ω and YSM models. The YSM and the k-ω

models failed to predict the velocity profiles correctly in the

separation zone, X ≥ 0. The mean axial velocity profiles

predicted by the SST model are in good agreement with

experiment both in the inner and outer flow regions. In the

downstream stations, the results of YSM and k-ω models are

almost similar but they are significantly different than the

measurement.

Co-annular sub- and super-critical swirling flows are

routinely used in gas turbine combustors. The standard k-ε

model fails to capture critical features of these flows. Nikjoo

and Mongia13 have recently evaluated the performance of the

standard k-ε model, quadratic and cubic YS models for the

two sets of swirling flows, namely, a single swirling jet with

no sudden expansion, and a co-annular confined swirling jet.

Comparisons of the measured and predicted mean swirl

velocity and profiles shown in Figure 3 co-annular jets.

As expected, the standard k-ε model captures qualitatively

the flow field characteristics. But there are substantial

differences in regard to swirl velocity (which model overpredicts). Similar discrepancies are noted in axial velocity, as

well.

The performance of the quadratic nonlinear k-ε model is

essentially comparable with that of the standard k-ε model.

However, the cubic non-linear k-ε model does significantly

better in regard to simultaneously matching swirl velocity

profile, jet spreading rate (i.e., location of the peak) and the

peak value of the axial velocity. The major deficiency of this

model is the under-prediction of the jet centerline velocity.

More modeling work is planned to get better agreement with

the data.

In order to meet the combustion system life requirements

within the constraints of significantly reduced cooling air

availability, designers are being challenged to reduce the

structure temperature margin used in the past applications.

Fig. 3 – Comparison between predicted (by Nikjoo and

Mongia, 1998) and measured (by Nikjooy et al., 1993 )

swirl velocity profiles of a confined annular jet a

Fig. 4 – Radial profiles of mean axial velocity, shear

stress and turbulence KE of a turbulent wall jet. a) Heat

transfer coefficient along jet center line. b) Radial

profiles of heat transfer coefficient at X/D = 3.0. c)

Radial profile of heat transfer coefficient at X/D =10.
Volume 51, Number 3 & 4 39

This means a more optimized utilization of the available

cooling air and significantly improved prediction capability.

Work is therefore needed to improve both the radiation and

convective heat transfer models similar to what is described

in the following paragraphs.

A non-gray sum-weighted, narrow-band radiation model

consistent with finite-volume 3-D gas turbine combustion

model has been formulated and coded for incorporation as a

module to the National Combustion Code being developed

jointly by NASA and industry. Kumar et al.

13 have done

preliminary assessment of this model. The wall temperature

levels calculated by the wide band model (which is

computationally faster than the narrow-band model by a

factor of more than 5) are quite close to those of the narrowband model for a typical liner at 40 atm. and 0.03 fuel/air

ratio; the resulting root-mean-square (RMS) error is quite

small. On the other hand, the differences between the

narrow- band and the conventional state-of-the-art models

are substantial. For nonluminous flames requiring no

estimates for soot concentration, a quasi-3D anchored

reduction model14 is about as good as one needs for calculating temperature levels of typical combustor structure.

The formulation and validation of a CFD-based convective

heat transfer model is being done in three steps. Step one

involves the validation of the model with benchmark quality

data in regard to flow field, film effectiveness and heat

transfer coefficients. Several promising turbulence models

have been evaluated.15 Figure 4 shows typical comparison

between predicted and measured data for turbulent wall jet in

regard to mean axial velocity, shear stress and turbulence

kinetic energy. Clearly, So and Sarkar turbulence model

(SSA) is superior to other turbulence models (Figure 4)

including standard k-ε, and models proposed by Chien, and

Yang and Shih (YS). Work is in progress for Step 2 that

involves making film effectiveness predictions for typical gas

turbine combustors, followed by (in Step 3) validation with

measured temperature levels and gradients of combustor

dome and liners.

Convective heat transfer model validation for multi-hole

liners (also called effusion cooling) is similarly being

conducted in the three steps as explained above for the

nugget liners. Kumar and Mongia16 compare their predictions

with data from Sen17, and Sinha18 for a single row of inclined

jets to a well-defined cross-flow stream. Typical comparison

for heat transfer coefficient (Figure 5) and film effectiveness

(Figure 6) with the SSA, YS and the standard k-ε models

clearly show preference for the SSA model for near-wall heat

transfer and fluid flow calculations.

Numerical calculations of a jet diffusion flame were performed by Tolpadi et al.19 using the eddy breakup, eddy

dissipation and the scalar PDF transport combustion models

with the standard k-ε model. The agreement between data

and calculations was rather unacceptable. Even though one

can offer several good reasons for the failure of the models to

achieve acceptable level of quantitative agreement for such a

simple test case, it is nevertheless a discouraging state of the

turbulent combustion modeling.

Even though a variety of the two-step eddy break-up model

first used by Reynolds et al. (1977) are being used in

industry, no one is quite happy with its demonstrated

accuracy levels. Hura et al.

20 have expended considerable

effort in establishing how far one can go in anchoring a twostep eddy breakup model in regard to matching a single cup

flow field data from the LM6000 DLE mixer. As shown in

Figure 7, one can indeed get a good anchored agreement

with measured centerline axial velocity using standard k- and

2-step eddy breakup models.

Fig. 5 – Comparison between measured (by Sen, 1995)

and calculated heat transfer coefficient (by Kumar and

Mongia, 2000) for a single 35-degree inclined jet in

cross-flow. a) Film effectiveness along the jet

centerline. b) Radial profiles of film effectiveness at X/D

= 1.0. c) Radial profiles of film effectiveness at X/D = 15.
40 Naval Research Reviews

It is logical that a skeletal mechanism is needed for accurately predicting HC, CO and NOx emissions. One had to

split up this in two steps – flow field calculations with

simpler turbulence/chemistry interaction model (e.g., a 2-

step eddy breakup model), followed by a reactor network

with a skeletal mechanism. Now it is possible to formulate a

single-step approach. As an example, a laminar flamelet

model has been used by Held and Mongia21 to demonstrate

its feasibility for predicting the flow-field, NOx and CO

emissions levels of the LM6000 DLE mixer. Figure 8

demonstrates the NOx prediction.

Under partial sponsorship of the AFOSR Focused Research

Initiative, Menon and his associates have been working on

the use of large eddy simulation with linear eddy (LESLE)

subgrid modeling. The feasibility of this approach for the

LM6000 DLE mixer has recently been demonstrated22, 23.

They used the same inlet profiles as anchored by Hura et

al.20(Their predicted centerline axial velocity (Figure 9)

should be compared with the results of Hura et al., shown in

Figures 7). This is a significant step forward with LESLE,

considering that the first set of calculations were made by

Kim and Menon22 with no prior experimental values, and no

attempt has been made to “further tune”. In subsequent

calculations, Kim and Menon23 have made further improvements to the models and obtained better agreement with data

as shown in Figure 9. One of the significant advantages of

the LESLE’s calculations is that it provides dynamic

structure of the flame. This type of information along with

measured data gives the research tools for providing the

insight required to develop the technology in a more costeffective and timely manner.

Combustion dynamic data from DLE combustor (premixed

combustion) and CFM family (diffusion combustion) were

collected and used to validate the Advanced Combustion

Dynamics (ACD) model. Figure 10 illustrates the comparison of predicted and measured organ tone frequencies. Two

discrete frequencies, namely mode 1 and mode 2 organ

tones, have been observed in some CFM engines. The hollow

triangles are the measured frequencies of mode-1 organ tone,

while the hollow circles represent the mode-2 frequencies.

Fig. 6 – Comparison between measured (by Sen et al.,

1991) and calculated film effectiveness (by Kumar and

Mongia, 2000) for a single 35-degree inclined jet in

cross-flow.

Fig. 7 – Comparison between measured and

“anchored” centerline and corner axial velocity

component of the LM6000 DLE mixer with a two-step

eddy breakup model (from Hura et al., 1998).

Fig. 8 – Feasibility demonstration of NOx prediction

with a laminar flamelet model and comparison with

data of the LM6000 DLE mixer (from Held and Mongia

(1998)).
Volume 51, Number 3 & 4 41

Since mean flow properties of combustor are required as

inputs for the ACD model, the steady state CFD analysis

with given cycle conditions were first conducted. Several

variables, including temperature, velocity vectors, and steady

state heat release, were extracted from steady state CFD flow

field data and piped into a common interface of ACD model.

The resonant frequencies were then obtained from the power

spectrum density generated by the model. The predicted

frequencies are presented by the solid symbols. The results

from ACD model predict the increase in resonant frequency

with increasing fan speed, the identical trend as observed

from the engine data.

The current production DLE combustor is a triple annular

premixed design. Two discrete frequencies, one at 450 Hz

and the other from 600 to 650 Hz have always been observed

in high power operating mode. Figure 11 shows the comparison between measured and predicted dynamic response. The

predicted response (shown by the thin line in figure) is

compared with a line drawn through the individual peaks of

the measured spectrum plot (shown by the thick line in

figure). The predicted value is obtained by taking the

maximum of the individual responses from each ring,

spectrum (b)-(e). The response characteristics of three

different acoustic modes (plane wave, m=1, and m=2), are

plotted in solid, dotted, and dashed lines, respectively, in

these figures. Within the frequency range of interest, the

model predicts two dominant frequencies, one at 500 Hz and

the other at 600 Hz. These frequencies are reasonably close

to those observed from engine tests at the specified operating

condition. It is self-illustrated from these figures that the 500

Hz mode is more active compared to the 600 Hz dynamics.

Based on the current analysis, the dynamic pressures

response in the vicinity of 500 Hz is mainly contributed by

the plane waves (m=0) from the middle and inner domes.

The 600 Hz acoustic mode is a result of the first order

circumferential mode arising in the outer ring. This analysis

did not indicate the presence of a second-order mode in the

combustor. In the high frequency region (above 1000 Hz) the

current model over predicts the dynamic response. It is

probably due to the effect of neglecting viscous effects in the

governing equations.

Summary

The R&D effort has resulted in significant progress in the

development and application of empirical/analytical design

methodology that is used routinely in the design process, and

in the semi-analytical mechanistic models formulation of

their feasibility demonstration for both the diffusion-flame

and premix combustors, further the feasibility and validation

of anchored CCD methodology is well established and its

use in the combustion design procedure is now a standard

practice in CCoE.

Good progress has been made for validating the Advanced

Combustion Code (w/ more accurate combustion models)

with benchmark quality database with target completion by

the end of next year.

It is anticipated that .the advanced combustion code effort

Fig. 9 – Comparison between measured and

“anchored” centerline and corner axial velocity

component of the LM6000 DLE mixer with Large Eddy

Simulation (from Kim and Menon, 1999).

Fig. 11 – Comparison between predicted and measured

power spectrum density plot for a premixed

combustion system.

Fig. 10 – Comparison between predicted and measured

organ tone frequencies for a diffusion flame

combustor.
42 Naval Research Reviews

will be completed by the end of the year 2000 providing a

significant advancement in combustor design.

Biography

Dr. Hukam Mongia graduated in 1965 from Punjab Engineering College and later received an MS and Ph.D. degrees

from the University of Massachusetts. Dr. Mongia joined

Garret Engine Company in 1972, where he was promoted to

Senior Supervisor, Combustors, in January 1981. He, later,

joined Allison Gas Turbines as Chief, Combustors R&D and

CFD. He then moved to GE Aircraft engines in January of

1994, where he is employed as the Manager, Technology and

Transition, combustion Center of Excellence. In this

capacity, Dr. Mongia is responsible for the provision of

leadership in technology, transition of technology into

combustion products, and design methodology and tools. He

has been involved in the design and development of numerous advanced technology and engine combustors, as well as

the formulation and successful application of empirical/

analytical design methodology in the combustion design

process

References

1. Mongia, H. C., and Smith, K. F., “An empirical/analytical

design methodology for gas turbine combustor”, AIAA Paper

78-998.

2. Mongia, H. C., Reynolds, R. S., and Srinivasan, R., 1986,

“Multidimensional gas turbine combustion modeling:

Applications and limitations”, AIAA J., Vol. 24, No. 6, June

1986, pp. 890-904.

3. Rizk, N. K., and Mongia, H. C., 1993, “Semianalytical

correlations for NOx, CO and UHC emissions”, J. Eng. Gas

Turbines and Power, July 1993, Vol. 115, pp. 603-611.

4. Tonouchi, J. H., Held T. J., and Mongia, H., 1997, “A semianalytical finite rate two-reactor model for gas turbine

combustors”, ASME Paper 97-GT-126

5. Danis, A. M., Burrus, D. L., and Mongia, H. C., 1997,

“Anchored CCD for gas turbine combustor design and data

correlation”, J. Eng. Gas Turbines and Power, 1997, Vol. 119,

pp. 535-545.

6. Tolpadi, A. K., Prakash, C., Hura, H., and Mongia, H. C.,

1998, “Advanced combustion code: Overall description,

prediction of a jet diffusion flame and combustor flowfields,”

ASME Paper 98-GT7. Launder, B. E., and Sharma, B. I., 1974, “Application of the

energy dissipation model of turbulence to the calculation of

flow near a spinning disc,” Letters in Heat and Mass Transfer,

1, pp. 131-138.

8. Yang, Z., and Shih, T. H., 1993, “New time-scale based k-ε

model for near-wall turbulence,” AIAA J., 31, pp. 1191-1198.

9. Wilcox, D. C., 1993, “Comparison of two-equation turbulence

models for boundary layers with pressure gradients,” AIAA J.,

31, pp. 1414-1421.

10. Karki, K. C., Oechsle, V. L., and Mongia, 1992, “A

computational procedure for diffuser-combustor flow

interaction analysis”, J. Eng. Gas Turbines and Power, January

1992, Vol. 114, pp. 1-7.

11. Nikjoo, M., and Mongia, H., 1999, “Predictions of flows with

adverse pressure gradients,” AIAA Paper 99-2817.

12. Driver, D. M., 1991, “Reynolds shear stress measurements in a

separated boundary layer,” AIAA Paper 91-1787.

13. Kumar, Ganesh N., Moder, Jeffrey, P., Mongia, Hukam C., and

Prakash, Chander, 1998, “Development of a three dimensional

radiative heat transfer computational methodology for aircraft

engine combustors,” AIAA Paper 98-0855

14. Kumar, G. N., and Mongia, H. C., 1998, “A quasi-3D anchored

radiation model for gas turbine combustors,” AIAA Paper 98-

3985.

15. Kumar, G. N., and Mongia, H., 1999, “Validation of turbulence

models for wall jet computations as applied to combustor

liners,” AIAA Paper 99-2250.

16. Kumar, G.N., and Mongia, H., 2000, “Assessment of advanced

turbulence models and unstructured code for calculating film

effectiveness of a modern film-cooled combustor.” Paper to be

presented at the 38th AIAA Aerospace Sciences Conference,

Reno, NV, January 2000.

17. Sen, B., 1995, “Effects of injection hole geometry, high free

stream turbulence, and surface roughness on film cooling heat

transfer,” Ph.D. Dissertation, University of Texas at Austin,

1995.

18. Sinha, A.K., Bogard, D.G., and Crawford, M.E., 1991, “FilmCooling Effectiveness Downstream of a Single row of Holes

with Variable Density ratio,” Transactions of the ASME,

Journal of Turbomachinery, Vol. 113, June 1991, pp. 442-449

19. Tolpadi, A. K., Prakash, C., Hura, H., and Mongia, H. C.,

1998, “Advanced combustion code: Overall description,

prediction of a jet diffusion flame and combustor flowfields,”

ASME Paper 98-GT-229.

20. Hura, H. S., Joshi, N. D., Mongia, H. C., and Tonouchi, J.,

1998, “Dry Low Emissions Premixer CCD modeling and

validation,” ASME Paper 98-GT-444.

21. Held, T. J., and Mongia, H. C., 1998, “Application of a

partially premixed laminar flamelet model to a low emissions

gas turbine combustor,” ASME Paper 98-GT-217

22. Kim, W.-W., Menon, S., and Mongia, H. C., 1998, “Large eddy

simulations of reacting flow in a dump combustor,” AIAA

Paper 98-2432.

23. Kim, W.-W., and Menon, S., 1999, “Large-Eddy simulation of

turbulent premixed flames in the thin reaction zone regime,”

AIAA Paper 99-2816.
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(fold on line and tape to secure)
Professor Ronald K. Hanson has been affiliated with the Mechanical Engineering Department of Stanford University since 1971, where he is the Woodward Professor and Department Chairman. Professor Hanson’s research has been in the fields of laser diagnostics, chemical kinetics, combustion, and propulsion. He is an internationally recognized

leader in the development of laser-based diagnostic methods for combustion and propulsion.

In the 1970s he pioneered the use of tunable IR diode laser absorption spectroscopy for non-intrusive measurements of

species and temperature in combustion environments. In the early 1980’s his group established a method, known as PlanarLaser-Induced Fluorescence (PLIF) and now used worldwide, for obtaining instantaneous 2-D images of combustion propertiesby means of laser excitation and digital camera recording. PLIF provides unique capability to observe complex flow structureswhich are currently beyond computation or measurement by other methods. Measurable quantities include species,temperature, pressure, and velocity. In the 1990s, Prof. Hanson and his students began to utilize room temperature tunablediode lasers (TDLs) as light sources in absorption diagnostics. These lasers offer high potential for rugged, compact, andeconomical diagnostics suitable for use in research and development. The potential of these diode lasers for sensing andcontrol of practical combustion and propulsion systems was recently validated through a collaboration with Naval AirWarfare Center (NAWC), China Lake, CA. Prof. Hanson’s group designed and implemented a multiple-wavelength diodelaser system which provided rapid sensing of temperature and key species in a forced combustor at NAWC. The temperaturedata were used in a fast, active control scheme to optimize combustor performance, with a particular objective of achievinghigh combustion efficiencies and minimum emissions of pollutants such as CO and unburned hydrocarbons. This pioneeringimplementation was highly successful, and has led to current research on sensors for other pollutant species.

A key advance in Prof. Hanson’s diode laser research was the development of measurement concepts based on the

shapes and positions of absorption lines. This measurement strategy enables simultaneous determination of multipleparameters, including quantities such as mass and momentum flux, which have not been measured previously by opticalmeans. During the past few years, under support from ONR and AFOSR, these TDL diagnostics have been applied assensors in research on combustion control, yielding near-real-time sensing ability which has aided development of newtypes of forced-combustion devices of interest to the Navy. At present, Prof. Hanson’s group is working to extend TDLdiagnostics to control gas turbine combustors, and to new applications in advanced propulsion systems such as pulsedetonation engines.

Professor Ronald K. Hanson

Mechanical Engineering Department

Stanford University

Stanford, CA

 could be understood, the

s can be validated, and the

ility can be improved. To this

various diagnostic technique

sition of state-of-the-art

ombustion. Prof. Ron Hanson

in his article, addresses the

ach he has taken to make

with the temporal and spatial

ol involves control of plume

ure and emission, as well as

n control of the jet plume with

lted in the invention of a novel

tor control. Here the exhaust

al fins, is performing the thrust

minating the drag induced by

s fins or bend-bodies. Prof. A.

e University and Prof. Paul

of Minnesota address this

on generated by the ONR basic

ssfully applied in various

gns such as ducted rockets,

ors, and stator-rotor designs.

Mongia of General Electric

he challenges in the design of

 combustion systems. The

onal combustion dynamics are

 probability density function

on and validation of various

described.

ogram is well coordinated with

and industry so that research

into applications. A balanced

 experimental approach is

 excellent contributions in the

eas with innovative concepts,

d experimentation.

Gabriel D. Roy

Office of Naval Research
Volume 51, Number 3 & 4/1999

and more . . .

Air Breathing ChemicalPropulsion: Meeting theChallenges

Timing is Everything p. 2

Laser Solved Mysteries p. 14

Going with the Coflow p. 24