John C. DeLaat, George Kopasakis, Joseph R. Saus, and Clarence T. Chang

Glenn Research Center, Cleveland, Ohio

Changlie Wey

ASRC Aerospace Corporation, Cleveland, Ohio

Active Combustion Control for Aircraft Gas-Turbine 

Engines—Experimental Results for an Advanced, 

Low-Emissions Combustor Prototype

NASA/TM—2012-217617

July 2012

AIAA–2012–783
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John C. DeLaat, George Kopasakis, Joseph R. Saus, and Clarence T. Chang

Glenn Research Center, Cleveland, Ohio

Changlie Wey

ASRC Aerospace Corporation, Cleveland, Ohio

Active Combustion Control for Aircraft Gas-Turbine 

Engines—Experimental Results for an Advanced, 

Low-Emissions Combustor Prototype

NASA/TM—2012-217617

July 2012

AIAA–2012–783

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

50th Aerospace Sciences Meeting

sponsored by the American Institute of Aeronautics and Astronautics

Nashville, Tennessee, January 9–12, 2012
Acknowledgments

The authors would like to thank Dr. Daniel Paxson for his invaluable insights concerning the phenomenon of combustion instability. 

The authors would also like to thank the staff of Test Cell CE5 at NASA Glenn Research Center for their excellence in conducting 

the experimental part of this effort. Finally, the authors thank the NASA Fundamental Aeronautics Supersonics Project and the 

Environmentally Responsive Aircraft Project for supporting this work.

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at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217617 1 

Active Combustion Control for Aircraft Gas-Turbine Engines—

Experimental Results for an Advanced, Low-Emissions 

Combustor Prototype 

John C. DeLaat, George Kopasakis, Joseph R. Saus, and Clarence T. Chang 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Changlie Wey 

ASRC Aerospace Corporation 

Cleveland, Ohio 44135 

Abstract 

Lean combustion concepts for aircraft engine combustors are prone to combustion instabilities. 

Mitigation of instabilities is an enabling technology for these low-emissions combustors. NASA Glenn 

Research Center’s prior activity has demonstrated active control to suppress a high-frequency combustion 

instability in a combustor rig designed to emulate an actual aircraft engine instability experience with a 

conventional, rich-front-end combustor. The current effort is developing further understanding of the 

problem specifically as applied to future lean-burning, very low-emissions combustors. A prototype 

advanced, low-emissions aircraft engine combustor with a combustion instability has been identified and 

previous work has characterized the dynamic behavior of that combustor prototype. The combustor 

exhibits thermoacoustic instabilities that are related to increasing fuel flow and that potentially prevent 

full-power operation. A simplified, non-linear oscillator model and a more physics-based sectored 1-D 

dynamic model have been developed to capture the combustor prototype’s instability behavior. Utilizing 

these models, the NASA Adaptive Sliding Phasor Average Control (ASPAC) instability control method 

has been updated for the low-emissions combustor prototype. Active combustion instability suppression 

using the ASPAC control method has been demonstrated experimentally with this combustor prototype in 

a NASA combustion test cell operating at engine pressures, temperatures, and flows. A high-frequency 

fuel valve was utilized to perturb the combustor fuel flow. Successful instability suppression was shown 

using a dynamic pressure sensor in the combustor for controller feedback. Instability control was also 

shown with a pressure feedback sensor in the lower temperature region upstream of the combustor. It was 

also demonstrated that the controller can prevent the instability from occurring while combustor operation 

was transitioning from a stable, low-power condition to a normally unstable high-power condition, thus 

enabling the high-power condition. 

Nomenclature 

ASPAC Adaptive Sliding Phasor Average Control 

FAR Fuel/Air Ratio 

ṁ Mass flow 

P Pressure 

P3 Combustor inlet pressure 

P3DynA Sensed dynamic combustor inlet pressure, first location 

P3DynB Sensed dynamic combustor inlet pressure, second location 

P4 Combustor pressure 

P4DynUp Sensed dynamic combustor pressure, upstream location 

P4DynDn Sensed dynamic combustor pressure, downstream location 
NASA/TM—2012-217617 2 

T Temperature 

T3 Combustor inlet temperature 

 Equivalence ratio 

Introduction 

Future aircraft engines must provide low emissions and high efficiency at low cost while maintaining 

the reliability and operability of present day engines. The demands for increased performance and 

decreased emissions have resulted in advanced combustor designs that are critically dependent on 

efficient fuel/air mixing and lean operation. However, all combustors, but most notably lean-burning 

ultra-low-emissions combustors, are susceptible to combustion instabilities. These instabilities are 

typically caused by the interaction of the fluctuating heat release of the combustion process with naturally 

occurring acoustic resonances (Refs. 1 and 2). If not mitigated, these interactions can produce large, highfrequency pressure oscillations within the combustor that could reduce component life and potentially 

lead to premature mechanical failures. 

Combustion instability has been a persistent issue in ground-based gas turbines using premixed 

combustors (Refs. 3 and 4) and will become more challenging as aero-engine combustor development 

continues to move toward leaner direct injection schemes. Effective suppression of the high-frequency 

combustion instabilities which result from the relatively short aero-engine combustor geometries is a 

critical enabling technology for lean-burning ultra-low-emission combustors and requires several key 

issues to be addressed. Dynamic models play a vital role in understanding the instability phenomenon and 

also for developing controls. Sensors and algorithms able to detect and interpret the instability need to be 

developed. Actuator devices that can introduce controlled-perturbations into the combustor to affect 

change in the instability are needed. And lastly, suitable control algorithms are needed to drive the 

actuators to obtain the necessary effect. 

In recent years, there has been considerable activity addressing active combustion control. 

Government, academia, and industry research efforts, through analysis and the use of laboratory 

combustors, have shown the considerable potential for active control (Refs. 5 to 15). At the NASA Glenn 

Research Center, the Active Combustion Control Technology activity aims to demonstrate active control 

in a realistic environment relevant to aircraft engines by providing experiments tied to aircraft gas turbine 

combustors. The intent is to advance the technology maturity of active combustion control toward 

eventual demonstration in an engine environment. Previous work at NASA Glenn has shown that active 

combustion control utilizing advanced algorithms working through high frequency fuel actuation can 

effectively suppress combustion instabilities in a conventional, rich-front-end combustor. The combustor 

was developed in conjunction with Pratt & Whitney and United Technologies Research Center and 

emulated a conventional aircraft gas turbine engine and its instability behavior (Refs. 16 to 19). For this 

effort, a suitable high-frequency fuel modulation device was developed by Georgia Institute of 

Technology and employed for modulating the combustor’s main stage fuel injector. A high-temperature 

single crystal silicon carbide (SiC) pressure sensor was able to detect the existence of the thermoacoustic 

instability in the combustor test rig while operating at 780 °F (420 °C) (Ref. 20). Two combustion 

instability control algorithms were each able to demonstrate up to 90 percent suppression of the 

combustor instability (Refs. 17 and 18). A dynamic model was utilized to evaluate instability control 

methods, and also to gain insights into the instability phenomenon (Ref. 21). It is desired to extend these 

active control technologies to advanced ultra-low-emissions combustors such as those employing multipoint lean direct injection (Ref. 22). 

In order to continue to mature the active combustion control technologies, a near term, but technically 

advanced low emissions combustor test bed is desired. Emissions testing of an advanced low-emissions 

combustor prototype was performed in 2006 in one of the combustion test cells at the NASA Glenn 

Research Center. During this testing, combustion instability pressure oscillations were observed. In 

addition, the instability pressure oscillations were observed to grow with increasing fuel/air ratio (FAR). 

This instability growth prevented the combustor from reaching full power operation. Dynamic data 
NASA/TM—2012-217617 3 

acquired from that combustor was analyzed in order to characterize the instability behavior of the 

combustor. Specifically, trends in instability amplitude versus operating conditions were observed and 

documented (Ref. 23). The dynamic characterization identified the combustor prototype as a suitable test 

bed for active controls research. 

A combustion instability simulation, previously developed at NASA Glenn (Ref. 21), was modified to 

capture the specific geometry and operating conditions of the advanced combustor prototype. The 

simulation layout embodied the relevant physical features of the combustor and test rig in order to provide 

a simulation of the experimentally observed instability behavior. The simulated combustion instability 

closely matched that of the combustor for steady fuel flow and also replicated the growth in instability 

amplitude as the FAR increased (Ref. 23). 

Most recently, in 2010, one of the previously developed instability control methods, the Adaptive 

Sliding Phasor Averaged Control (ASPAC) (Ref. 18) was updated for the low-emissions combustor 

prototype. Active combustion instability suppression using the ASPAC control method has been 

demonstrated experimentally with this combustor prototype in a NASA combustion test cell operating at 

engine pressures, temperatures, and flows. Dynamic pressure sensors in the combustor and in the 

combustor inlet provided combustor pressure feedback to the controller. A high-frequency fuel valve was 

utilized to perturb the combustor fuel flow. Successful instability suppression was shown. It was also 

demonstrated that the controller can prevent the instability from occurring while combustor operation was 

transitioning from a stable, low-power condition to a normally unstable high-power condition. 

This paper provides a description of the combustor rig experimental setup and synopsizes the 

experimental dynamic characterization. The sensors and fuel actuator used for active control are then 

described. An overview of the instability control method and how it was adapted for the low-emissions 

combustor is then given. Data from the experimental demonstration of combustion instability control are 

provided and discussed. The paper concludes with a summary of the experimental results and a discussion 

of the continuing technical challenges and long range plans for NASA’s efforts in active combustion 

control. 

Description of the Experimental Hardware and Software 

Low-Emissions Combustor Prototype 

Several advanced low emissions combustor concepts for aircraft gas turbine engines have been 

developed in conjunction with NASA’s industrial partners. Thermoacoustic instability was observed 

during emissions testing of one of these advanced combustor prototypes in 2006. This combustor 

prototype operates on liquid jet fuel (typically JetA or JP8) and consists of a pilot and main stage. The 

pilot operates locally rich and the main stage operates locally lean, however the overall fuel/air ratio 

(FAR) is always lean with typical equivalence ratio of  = 0.3 to 0.5 (FAR = 0.02 to 0.034). Here  = 1 is 

the stochiometric FAR (for liquid jet fuel,  = 1 is equivalent to FAR = 0.068). 

The combustor was installed in a flame tube in test cell CE5 at NASA Glenn (Fig. 1). The combustor 

and rig were operated at the range of conditions shown in Table I. The combustor was instrumented with 

dynamic pressure transducers at the locations shown in Figure 1. There were two water cooled dynamic 

(AC-coupled) pressure transducers upstream of the fuel injector (P3DynA, P3DynB). Downstream of the fuel 

injector, two uncooled dynamic pressure transducers (P4DynUp, P4DynDn) were each installed in a purged, 

semi-infinite line configuration in order to isolate the transducers from the hot combustion products. The 

downstream transducer, P4DynDn, had less background noise and is the one most often used for the 

dynamic characterization and feedback control presented in this paper. 
NASA/TM—2012-217617 4 

Figure 1.—Advanced low-emissions combustor prototype installed in NASA flame tube. The pilot fuel line is shown in 

red and the main stage fuel lines are shown in green. Dimensions shown are in inches. Drawing is not to scale. 

TABLE I.—RANGE OF COMBUSTOR 

OPERATING CONDITIONS

Inlet pressure (psia) 65 to 250 

Inlet temperature, F 400 to 1000 

Air flow, lbm/s 0.9 to 4.0 

Fuel flow, lbm/hr approx. 100 to approx. 400 

Combustor ΔP/P, percent approx 4.0 to 4.5 

Fuel/air ratio 0.02 to 0.034 

Characterization of Combustion Instability Versus Operating Conditions 

During emissions testing with this combustor in 2006, dynamic pressure data were collected and the 

dynamic behavior of the combustor characterized (Ref. 23). Observations from this test series showed an 

instability that occurred at approximately 530 Hz. This instability was only present when operating on the 

main stage of the fuel injector, that is, it was not present for low power, pilot-only operation. Further, the 

instability was seen to generally increase in amplitude with increasing FAR. Due to concerns about 

damage to the fuel injector as the instability amplitude increased, the combustor was not operated at its 

full power FAR. The dynamic pressure data were recorded on a dSPACE Inc. real-time data acquisition 

system along with operating condition data. This allowed data to be acquired during instability, and also 

during the onset of instability. For all of the analyses shown in this paper, combustor dynamic pressure 

data were acquired with a sample time of 5 kHz. For steady fuel flow conditions, 10 sec of data were 

typically acquired. The data were processed to obtain a power spectral density (PSD) of the pressure. 

The MATLAB (The Mathworks, Inc.) PSD function, which implements Welch’s method, was applied to 

the data. A Fast Fourier Transform (FFT) length of 4096 was used and the data were segmented with a 

Hanning window of length 4096 and a window overlap of 75 percent of the window size. The spectral 

data were scaled to provide spectral amplitude in psi. A plot of the resulting peak value from the spectral 

data (the value at the instability frequency) versus FAR is shown in Figure 2. A simple curve-fit has been 

added for each set of data points to aid identification of trends in the data. Repeats during later test runs 

are included and are shown in Figure 2 as hollow symbols with dashed-line curve fits. The curves are 

Dynamic pressure transducer, P4DynUp

Dynamic pressure transducer, P4DynDn

Dynamic pressure transducer, P3DynB

Dynamic pressure transducer, P3DynA

1

2.

0

1

2

12.0

1

2

.

0

1

2

3.0
NASA/TM—2012-217617 5 

Figure 2.—Combustor pressure peak instability amplitude versus fuel/air ratio for multiple test runs 

shows increase in combustion instability amplitude with increasing fuel/air ratio (from Ref. 23). 

labeled according to run number, combustor inlet pressure, combustor inlet temperature, combustor 

airflow, and pilot/main fuel split percentage. The overall trend of instability peak amplitude increasing 

with FAR can be seen in all the data sets. Further details on the dynamic characterization of the 

combustor can be found in Reference 23. 

Due to the repeatable nature of the exhibited combustion instability across multiple operating 

conditions, as well as the ability to size the instability by varying FAR and operating conditions, this leanburn combustor prototype was selected as a suitable test bed for the combustion control effort described 

later in this paper. 

Instability Control Sensors 

As mentioned previously, and as shown in Figure 1, four high-frequency pressure sensors were used to 

measure thermoacoustic pressure oscillations in the combustor prototype during the 2006 combustor 

testing. For the instability control demonstration described in this paper, similar dynamic pressure sensors 

were used at similar locations. For the combustor inlet pressure (P3), two dynamic (AC-coupled) pressure 

sensors (P3DynA, P3DynB) were located approximately 17 in. (43 cm) upstream of the fuel injector face at 

two different circumferential locations. These sensors are PCB Piezotronics type 123Mxx pressure 

sensors with 100 psi range. Water cooling is used to keep the sensors below their maximum operating 

temperature of 250 F (120 C). For the pressure inside the combustor (P4), two dynamic pressure sensors 

(P4DynUp, P4DynDn) were located at approximately 7 in. (18 cm) and 27 in. (68.5 cm) downstream of the fuel 

injector face, respectively. These sensors are PCB type 112Axx sensors with 100 psi range. These 

sensors are each located approximately 12 in. (30 cm) from the flowpath on a purged, semi-infinite line. 

A comparison of how well the sensors at P3 are able to measure combustion instability versus those at P4

is given later in the experimental results section. 

0

0.2

0.4

0.6

0.8

1

1.2

0.0220 0.0240 0.0260 0.0280 0.0300 0.0320 0.0340 0.0360

Fuel/Air Ratio

Combustor Pressure Amplitude at Instability Freq., psi 

Run 905 - 135psia, 1000degF, 1.84pps, 10/90

Run 905 - 151psia, 1000degF, 2.05pps, 10/90

Run 905 - 151psia, 1000degF, 2.05pps, 20/80

Run 901 - 166psia, 1000degF, 2.26pps, 10/90

Run 906 - 166psia, 1000degF, 2.26pps, 10/90

Run 901 - 192psia, 950degF, 2.66pps, 10/90

Run 906 - 192psia, 950degF, 2.66pps, 10/90

Run 901 - 250psia, 1000degF, 3.4pps, 10/90

Run 905 - 250psia, 1000degF, 3.4pps, 10/90

Run 901 - 250psia, 1000degF, 3.4pps, 20/80

Run 905 - 250psia, 1000degF, 3.4pps, 20/80

Suspect Data Point
NASA/TM—2012-217617 6 

High-Frequency Fuel Actuator 

For the open-loop fuel modulation testing and closed-loop instability control testing, a high-frequency 

fuel modulation valve developed previously (Refs. 18 and 19) was utilized. The valve is actuated by a rod 

comprised of a magnetostrictive material. The fuel valve has an associated control system with a mean 

flow control loop that keeps the mean fuel flow constant while the fuel flow is perturbed. The design 

application for the valve was to modulate the main stage fuel at up to 1000 Hz. For the previous 

combustion control application (Refs. 16 to 18), the combustor main fuel injector Flow Number was 

about 110. Flow Number for an orifice is defined as: 

P

m

   Flow Number (1) 

with ṁ being the fluid mass flow rate in lbm/hr (abbreviated pph) and ∆P being the pressure differential 

in psi across the orifice. The resulting units for Flow Number are (pph/psi0.5), though they are rarely 

stated.1

 

It was desired for the current test that the valve modulate the fuel flow through the fuel injector pilot 

which carries a smaller portion of the overall fuel flow (typically 10 to 20 percent) through a 

correspondingly smaller Flow Number of approximately 8. Using a dynamic flow test rig (Ref. 24), the 

fuel valve was dynamically characterized to try to predict how well it would be able to modulate the pilot 

fuel. Fuel line lengths and sizes were set in the characterization rig to approximate the lengths and 

volumes for the fuel feed system in the combustor rig. The results of the valve dynamic characterization 

(Ref. 25) suggested that the fuel valve, operating through the pilot, might not have enough actuator 

authority to impact the combustion instability. However, since no appropriately sized valve was available 

at the time, it was decided to proceed to combustion testing. As a fallback position, the valve would be 

used to modulate the main stage fuel, which for this fuel injector had a Flow Number of approximately 

30. The performance of the valve during combustor testing is described in the experimental results section 

of this paper. 

Control Method Description 

The Adaptive Sliding Phasor Averaged Control (ASPAC) algorithm was developed previously and 

used to control combustion instability in a conventional, rich-burn combustor. The control algorithm and 

previous results are described in detail in Reference 18. As depicted in Figure 3, combustor 

thermoacoustic instability is a closed-loop, self-excited process that occurs when combustion excites the 

combustor acoustics, and the acoustics feed back into the combustion process with the appropriate phase. 

To suppress combustion instability, a controller must interrupt this naturally occurring feedback. The 

ASPAC algorithm senses combustor pressure fluctuations, uses a band-pass filter to isolate the instability 

frequency, and modulates the fuel valve. The pressure oscillations generated by combustion of the 

ASPAC-modulated fuel oppose in phase the pressure oscillations generated by the instability, reducing 

the net amplitude of the combustor pressure oscillations. As shown diagrammatically in Figure 4, the 

sensed pressure is phase shifted with a phase that slides back and forth inside a restricted control region 

that the controller has identified as being favorable to suppression of combustion instability. This phase 

shifted pressure is used to generate the command to the fuel valve. The combustor pressure resulting from 

heat addition due to the fuel modulation, vectorially adds to the combustor pressure due to the instability 

to provide the overall dynamic combustor pressure. The dashed line region (the “effective stability 

region”) within the shaded area in Figure 4 is the stable region in which the phase of the fuel modulation 

 1

The effective flow area of a pressure atomizer such as a fuel injector nozzle is often described in terms of Flow 

Number. However, Flow Number is dependent on the fluid density (Ref. 1). For reference, with jet fuel, a Flow 

Number of 110 correlates to an effective cross-sectional area of approximately 6.510–3 in. 
NASA/TM—2012-217617 7 

Figure 3.—ASPAC combustion instability control block diagram showing filter and 

phase shift controller that are the heart of the ASPAC algorithm. 

Figure 4.—ASPAC Phasor diagram depicted in a stationary frame of reference shows 

how pressure from properly phased fuel modulation decreases overall combustor 

pressure oscillations. 

relative to the instability pressure is such that the power of the combustor pressure oscillations is reduced. 

The controller phase slides back and forth within the effective control region; starting from one end of its 

boundary, marching to the other end and then reversing direction. This continuous dithering of the phase 

allows the controller to respond quickly to any changes in phase required to maintain instability 

suppression. 

The controller command is generated at a 10 KHz rate and the controller calculates and applies a new 

phase shift at a rate of 40 Hz. Also, this controller optionally employs discontinuous exponential gain 

modulation control. In this control mode the gain toggles on and off with an exponential decay in order to 

counteract the effective proportional gain variability produced by the large dead time phase delay of the 

plant. In addition, controller parameter adaptation is employed to tune some of the key parameters of the 

controller. 

For the current effort, the ASPAC controller has been updated to provide active control of combustion 

instability in the low-emissions combustor prototype. The controller was modified via the band-pass filter 

design, to be more sensitive to the instability frequencies observed in the low-emissions combustor. 

Initially, a simple non-linear oscillator model was used to tune and exercise the controller. In order to 

conduct more detailed simulation evaluations of instability control methods, and also to gain insights into 

the instability phenomenon, the NASA Sectored 1-D combustion instability simulation method was 

Phase Shift

Controller

Fuel

Valve

Filter

Combustor

Acoustics Flame

Closed‐Loop Self‐Excited System

Naturalfeedback process

Fuel‐air

Mixture

system

Combustor

Pressure

Fuel Valve

Command

ASPAC Algorithm
NASA/TM—2012-217617 8 

utilized. The simulation layout embodies the relevant physical features of the combustor and test rig. As 

described in Reference 23, modifications were applied to the basic simulation in order to match the 

physical layout shown in Figure 1, and capture the experimentally observed instability behavior of the 

low emissions combustor prototype as shown in Figure 2. The Sectored 1-D simulation was used in a 

limited fashion for evaluation of the controller prior to experimental testing of the controller in the 

combustor rig. Further runs with the Sectored 1-D simulation are planned with the results to be compared 

to the experimental data described next. However, as will be discussed, the behavior of the combustor rig 

changed with time so that the Sectored 1-D simulation, which captured the relatively static behavior of 

the rig in the 2006 testing, no longer captured this changing combustor instability behavior. The 

discussion of experimental results later in this paper briefly explores options on how to proceed with postrun simulation evaluations. 

Instability Control Experimental Demonstration 

The updated ASPAC controller was implemented on a dSPACE Inc. real-time control computer and 

operated with the low emissions combustor prototype to experimentally demonstrate combustion 

instability control. Experimental testing was conducted in the NASA Glenn Research Center combustor 

test cell CE5 from September through November of 2010. There were five research objectives for the 

experimental program: 

A. Determine if the combustion instability behavior of the low-emissions combustor prototype is 

similar to that observed during the 2006 testing; 

B. If Objective A yields a suitable operating condition for combustion instability control 

investigations, attempt to demonstrate that the ASPAC controller is able to suppress combustion 

instabilities in the low-emissions combustor prototype and, as a result, extend the combustor 

operating range into previously unstable regions; 

C. If Objective B is successful, determine if combustion instability control can be accomplished 

using the dynamic pressure at P3 for feedback; 

D. Also, if Objective B is successful, determine if combustion instability control can be accomplished 

through modulation of the pilot fuel flow; and 

E. Obtain dynamic characterization data during the closed-loop, instability control activities in 

Objectives B through D to allow construction of a closed-loop version of the Sectored 1-D 

combustor simulation that can be used as a benchmark problem for evaluating other combustor 

instability control methods and/or other combustion instability simulation methods. 

Experimental results addressing each of these research objectives are presented and discussed in the 

following sections. 

Comparison of Combustion Instability Behavior Versus Prior Testing 

As discussed previously, the combustion instability behavior of the low-emissions combustor 

prototype as shown in Figure 1 was dynamically characterized to obtain the relationship between 

combustion instability and operating conditions such as air temperature, air pressure, and fuel/air ratio. 

The results from testing in 2006 are reported in Reference 23 and are summarized in Figure 2. Since 

several years and several rig configuration change-outs had occurred, the first set of test runs in 2010 

attempted to replicate the 2006 behavior. The expectation was that the combustor would behave similarly. 

Sensor P4DynDn was again used to measure instability pressure oscillations. As was seen in 2006, the 

instability was only present when operating on the main stage of the fuel injector, that is, it was not 

present for low power, pilot-only operation. Comparisons were then run for combined pilot/main 

 
NASA/TM—2012-217617 9 

Figure 5.—Combustion instability pressure peak amplitude versus FAR for a single representative operating 

condition shows decreasing instability amplitude versus time in 2010 as compared to 2006 instability behavior. 

The data shown is for P3 = 192 psia, T3 = 950 F, ṁair =2.66 lbm/s, 10/90 fuel split (Pilot/Main). 

operation. Figure 5 shows, for one of the operating conditions shown in Figure 2, the 2006 instability 

behavior versus FAR compared to the same operating condition during the 2010 runs. As can be seen in 

the figure, during the first 2010 run, the instability behavior (amplitude vs FAR) seemed similar to that 

observed in 2006. However, as also seen in Figure 5, after the first run, the instability amplitude 

decreased from run to run. Eventually the combustor was essentially quiet and there was no increase in 

instability amplitude with increasing FAR. Similar behavior (that is, a decrease in susceptibility to 

instability) was observed for all other operating conditions as well. It was found that to obtain instability 

pressure oscillations of sufficient amplitude to conduct the combustion instability control experiments 

described next, it was necessary to operate the combustor off-design. For example, decreasing the 

combustor ΔP/P for the same FAR was found to increase the instability amplitude. The resulting actual 

conditions used for combustion control testing will be described in the next section. 

A definitive cause for the continual decrease in susceptibility to instability was not identified during 

the testing. It was suspected that the fuel injector might be accumulating some damage that was changing 

how the fuel and air were delivered to the combustor. However, examination of the fuel injector hardware 

between runs showed no obvious physical changes. Near the end of the test cell entry, it was noticed in 

the test data that the main stage fuel injector Flow Number had been decreasing steadily over the weeks of 

testing (Fig. 6). It was suspected that this may have been due to coking. Between two of the runs, 

detergent was run through the fuel injector and the Flow Number partially (but only temporarily) 

recovered, as did the instability amplitudes. It was hypothesized, but not verified, that the change in fuel 

injector Flow Number may have changed how the fuel was delivered to the flame, thus impacting the 

instability mechanism. Since the attempted fix was only partial, and temporary, it remained necessary to 

run the fuel injector off-design to get sufficient instability amplitudes for controls testing. 

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.022 0.024 0.026 0.028 0.03 0.032 0.034 0.036

Combustor PressureAmplitude atInstability Freq., psi

Fuel/Air Ratio

Run 901 ‐ 7/26/2006

Run 906 ‐ 9/12/2006

Run 1007 ‐ 9/1/2010

Run 1008 ‐ 9/2/2010

Run 1014 ‐ 11/4/2010
NASA/TM—2012-217617 10 

Figure 6.—Main fuel injector average flow number from each test date, normalized to 

initial flow number, shows flow number decreasing over time. Flow number partially 

recovered when the fuel injector was flushed with detergent to remove possible coking. 

Demonstration of Combustion Instability Control 

From the testing described in the previous section, the operating condition P3 = 166 psia, T3 = 

1000 F, with pilot/main fuel split percentage of 10/90 was chosen for instability control investigations. 

Depending on the FAR and combustor ΔP/P chosen, this point exhibited an instability with peak pressure 

amplitude of between 0.5 and 1.0 psi peak amplitude (1 to 2 psi peak-to-peak) at approximately 620 Hz. 

Figure 7 shows the amplitude spectra and time history plot of the combustor dynamic pressure, P4DynDn, 

for a FAR of 0.037 and combustor ΔP/P of 3 percent. The pressure oscillations can clearly be seen in the 

amplitude spectra as well as in the time history. There is also a small harmonic response. The 620 Hz 

instability frequency observed here is higher than the 530 Hz instability observed in the 2006 testing. This 

is believed to be due to the higher FAR required to elicit the instability producing a higher temperature 

and thus a higher speed of sound in the combustor. 

The high-frequency fuel modulation valve was installed in the main stage fuel injector supply line as 

close to the fuel injector as the rig installation would allow (about where the word “Fuel” is shown on 

Fig. 1). This was about 3 in. from the rig pressure vessel. However, due to stress relief bends in the fuel 

line inside the pressure vessel, there was still approximately 32 in. from the fuel valve to the fuel injector 

tip. Open-loop, sinusoidal valve command variations were provided to the valve to observe the response 

of combustor pressure to main stage fuel injector perturbations. The commanded frequency and voltage 

level to the valve were varied and combustor pressure P4DynDn monitored. The limits on valve command 

voltage were set to values found in prior testing (Ref. 25) to provide maximum linear valve travel without 

valve overheating. Fixed frequency and sweep frequency perturbations from 100 to 1500 Hz were 

commanded to the valve. Representative results are shown in Figure 8. Shown are the valve commanded 

voltage and the combustor dynamic pressure for commanded frequencies of 500 and 800 Hz. Along with 

the instability pressure oscillations (as shown previously in Fig. 7), the response to the fuel modulations 

(and a small harmonic response) can also be seen in the combustor pressure. 

 

0.84

0.86

0.88

0.9

0.92

0.94

0.96

0.98

1

1.02

Normalized Main Injector Flow Number

Testing Date

Injector Flushed w/ Detergent
NASA/TM—2012-217617 11 

Figure 7.—Amplitude spectra and time history plot of combustor dynamic 

pressure showing the combustion instability at 620 Hz for operating 

conditions P3 = 166 psia, T3 = 1000 F, FAR = 0.037 and combustor 

∆P/P = 3 percent. 

a) 500 Hz b) 800 Hz 

Figure 8.—Combustor dynamic pressure (P4DynDn) shows both the combustion instability and the response to open 

loop fuel perturbations commanded to the valve at two different frequencies. 

101 102 103 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

Combustor Pressure, psi

Frequency, Hz

1 1.01 1.02 1.03 1.04 -1.5

-1

-0.5

0

0.5

1

1.5

2

psi

Time, sec

Instability peak at 620 Hz

Combustor Pressure, psi

psi

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

Com bustor Pressure, psi

-1.5

-1

-0.5

0

0.5

1

1.5

2

psi

101 102 103 0

0.5

1

1.5

2

2.5

3

3.5

4

Valve Com m and (REFA), Volts

Frequency, Hz

1 1.01 1.02 1.03 1.04

-4

-3

-2

-1

0

1

2

3

4

Volts

Time, sec

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

Com bustor Pressure, psi

-1.5

-1

-0.5

0

0.5

1

psi

101 102 103 0

0.5

1

1.5

2

2.5

3

3.5

Valve Com m and (REFA), Volts

Frequency, Hz

1 1.01 1.02 1.03 1.04

-4

-3

-2

-1

0

1

2

3

4

Volts

Time, sec

Combustor 

Pressure 

Response to Fuel 

Modulation

Combustor 

Pressure 

Response to Fuel 

Modulation

Valve Command, Volts 

Valve Command, Volts 
NASA/TM—2012-217617 12 

While the amplitude of the response to fuel modulation appears small relative to the amplitude of the 

instability, it was observed during testing that at certain frequencies, especially those close to the 

instability frequency (e.g., within 25 Hz), the instability amplitude would get smaller. Also, for 

frequencies very close to the instability (e.g., within 10 Hz), the instability frequency would shift to that 

of the fuel modulation frequency. Both of these effects indicate that the fuel modulation is affecting the 

instability and will thus be able to impact the instability via closed-loop control. It was thus concluded 

that fuel modulation via the main stage injector had sufficient authority to proceed to closed-loop testing. 

With the instability behavior just described present in the combustor, the ASPAC controller was 

turned on, allowing it to command the fuel valve. The sensor P4DynDn was used as the controller feedback. 

Initially, the gain for the controller was manually set to zero, which commanded zero voltage 

perturbations to the valve. The gain was then slowly increased until good instability suppression was 

observed and then until the commanded voltage to the valve approached its limits. This was done to give 

the controller the maximum available actuator authority. However, it was found that increasing the gain 

beyond a certain value caused the instability suppression to be erratic. This is expected as a result of the 

large time delays and noise in the system. As the instability gets sufficiently small compared to the 

system noise, the noise will randomize the phase of the pressure oscillations. A large gain will then tend 

to excite the instability during those times when the ASPAC algorithm is going through larger-than-usual 

phase adaptations (i.e., when the phase is not completely correct for suppression). 

Some manual and automated tuning of the controller’s parameters (e.g., gain, filter pass-band, phase 

adaptation limits, instability growth allowed before large scale phase adaption kicks in, etc.) was then done 

to obtain the best instability suppression as determined from real-time spectral analysis of the combustor 

pressure. Figure 9 shows a combustion instability control result using P4DynDn as feedback and modulating 

fuel through the main stage injector. As shown in the spectral plots on the left, when the controller is on, the 

Figure 9.—Combustor pressure amplitude spectrum (left) and time 

history (right) shows the combustion thermoacoustic instability 

being suppressed by the application of closed-loop control. For 

this data, P4DynDn was used as controller feedback, and the fuel 

valve was modulating fuel through the main stage fuel injector. 

a) Without control

b) With control

101 102 103 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

Combustor Pressure, psi

Frequency, Hz

1 1.01 1.02 1.03 1.04 -2

-1.5

-1

-0.5

0

0.5

1

1.5

2

psi

Time, sec

101 102 103 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

Combustor Pressure, psi

Frequency, Hz

1 1.01 1.02 1.03 1.04 -2

-1.5

-1

-0.5

0

0.5

1

1.5

2

psi

Time, sec
NASA/TM—2012-217617 13 

amplitude of the pressure oscillations at the instability frequency is reduced by nearly 95 percent from its 

uncontrolled value (from 0.84 to 0.05 psi). The time history plot on the right shows a decrease in the peakto-peak pressure oscillations from about 2.5 psi to about 1 psi. A second oscillatory mode around 500 Hz 

was sometimes observed in the combustor pressure. This resonance can be seen in the left plots in Figure 9. 

Since the controller was not designed to address this other resonance, this sometimes limited the amount the 

time history of the pressure oscillations (as shown in the right side of Fig. 9) could be reduced. 

A useful result would be to quantify how much fuel mass flow modulation is required to accomplish 

this instability suppression. However, there are no dynamic mass flow measuring devices in the fuel lines 

for this combustor rig, so the flow modulation must be estimated from the fuel injector ΔP. Further, since 

no dynamic ΔP measurement across the fuel injector is available, the fuel injector ΔP is also estimated 

from the modulation valve exit pressure minus the time average P4. The result is a very rough estimate of 

the size of the fuel modulations entering the combustor. The following example is for the case shown in 

Figure 9. First, Equation (1) is used to calculate the main injector Flow Number. Using the time-average 

value for valve exit pressure of 210 psia, for P4 of 150 psia, and for fuel flow of 220 lbm/hr, the main 

injector Flow Number is found to be 28. This is consistent with the expected value stated earlier in the 

paper. Next, the fuel mass flow modulation is estimated from the fuel injector ΔP and Flow Number by 

inverting Equation (1). From valve exit pressure modulations of approximately 20 psi about the mean, 

the fuel mass flow modulations required for instability suppression are estimated to be around 35 lbm/hr 

or approximately 8 percent of the mean fuel flow. 

While it is interesting to know if the controller is able to suppress an instability that is already 

occurring, the more desirable controller action would be to prevent the instability before it has a chance to 

get large enough to be of concern. To test the ability of the ASPAC controller to prevent a combustion 

instability from developing, a fuel/air ratio transient was commanded by the facility operators to transition 

the combustor from a region of stable operation to a region of unstable operation (the facility controls the 

overall fuel and air through the combustor). This FAR transient, without the ASPAC controller active, is 

shown in the left part of Figure 10. Shown are the fuel modulation valve command (inactive in this case), 

the combustor dynamic pressure, and the FAR. The combustor pressure has been band-pass filtered about 

the instability frequency (the filter pass-band is about 100Hz around a center frequency of 620 Hz) in 

order to deemphasize noise and accentuate the pressure oscillations due to the instability. 

As can be seen, with no control, the instability develops and the pressure oscillations due to the 

instability grow to approximately the levels shown in Figure 9. However, as shown in the right side of 

Figure 10, with the ASPAC controller active, repeating the same FAR transient shows no increase in 

pressure oscillations. Interestingly, the valve command shows very little increase in activity when the 

instability would be developing due to the fact that the controller is already monitoring combustor 

pressure and modulating the fuel valve to minimize the energy in the combustor pressure. These two FAR 

transients (with and without the controller active) show that the ASPAC controller is able to prevent the 

growth of instability before it can reach an unacceptable level. As discussed earlier, since the instability 

amplitude for the combustor grows with increasing FAR, in order to achieve higher power operation, the 

instability growth must be reduced or eliminated. The results shown here, then, are highly encouraging. 

They show that, with combustion instability control, the combustor can operate in regions that had 

previously been avoided due to instability, thus enabling an expanded operating range for the combustor. 

Another experiment was conducted to observe how quickly the ASPAC controller is able to suppress 

an instability when the controller is switched on. This is of interest because, in contrast to the previous 

test, it would be operationally desirable to only have the controller active when an instability is starting to 

develop. This would decrease the activity of the high-frequency valve and extend its life. The left side of 

Figure 11 shows the ASPAC controller being turned on in the presence of an already developed 

instability. This would be the worst case, since, as was already discussed, it would normally be 

undesirable to let the instability become fully developed. As seen in the figure, after the controller is 

switched on at a little after 4 sec, the controller starts modulating the fuel valve command (the top plot) 

and is able to start impacting the instability pressure oscillations (the bottom plot) almost immediately. 
NASA/TM—2012-217617 14 

a) Without ASPAC control b) With ASPAC control active 

Figure 10.—Growth of the combustion thermoacoustic instability is prevented by the application of the ASPAC 

closed-loop control during an increase in fuel/air ratio. Shown (bottom to top) are combustor rig experimental 

data for the fuel/air ratio, combustor pressure, and modulation valve command voltage.

a) ASPAC controller switched on at ~4 sec b) ASPAC controller turned off at ~4 sec 

Figure 11.—Time history of the fuel modulation valve command and the combustor pressure shows the time 

required for the ASPAC controller to suppress an already developed instability; and time required for an 

instability to re-develop when the ASPAC controller is switched from on to off. 

-5

0

5

Valve Command, Volts -2

-1

0

1

2

Combustor Pressure, psi

0 5 10 15 20 0.03

0.035

0.04

0.045

Fuel/Air Ratio

Time, sec

-5

0

5

Valve Command, Volts -2

-1

0

1

2

Combustor Pressure, psi

0 5 10 15 20 0.03

0.035

0.04

0.045

Fuel/Air Ratio

Time, sec

-5

0

5

Valve Command, Volts

0 5 10 15 20

-2

-1

0

1

2

Combustor Pressure, psi

Time, sec

-5

0

5

Valve Command, Volts

0 5 10 15 20

-2

-1

0

1

2

Combustor Pressure, psi

Time, sec
NASA/TM—2012-217617 15 

However, this early suppression is of poor quality in that it allows the instability to re-grow while the 

controller narrows the region of phase shifting required for suppression. The abilities of the controller 

early on are also limited due to the fuel valve saturating and not providing as much authority as the 

controller is requesting. As the controller converges on the optimum phase shift region, the quality of the 

suppression improves until the suppression is able to be maintained. The right side of Figure 11 shows 

that if the controller is turned off again, the instability quickly re-develops. These results show that the 

relative time for the instability to fully develop is shorter than the time the controller takes (from an 

entirely off state) to suppress this fully developed instability. This reinforces the importance of the 

previous results whereby the instability is prevented from developing. This also points to a desired mode 

of operation where the controller monitors the combustor pressure continually (or at least in the vicinity 

of conditions known to be prone to instability). The controller could then “get a head start” on instability 

suppression by observing the instability as it is starting to develop. This mode of controller activation was 

not able to be tested during this test cell entry, but is considered to be a highly desirable feature for future 

implementations of the ASPAC controller. 

Combustion Instability Control With P3 Dynamic Pressure as Feedback 

The typical temperature just upstream of the combustor (T3  1300 F) is considerably less harsh than 

the typical temperature inside the combustor (T4 > 3000 F). Thus it is highly desirable to be able to 

utilize a pressure sensor at P3 for controller feedback rather than one at P4. Since pressure sensor 

technology is moving in the direction of pressure sensors that can operate at temperatures in excess of 

1000 F, being able to use P3 for feedback instead of P4 would enhance the technical viability of 

combustion instability control. To investigate the use of P3 dynamic pressure for controller feedback, the 

measured dynamic pressure at P3 (P3DynA) was compared to the dynamic pressure at P4 (P4DynDn) (see 

Fig. 1 for the locations of these sensors). Figure 12 shows a comparison of the measured values for these 

two pressure sensors for the combustion instability condition shown previously in Figure 7. As can be 

seen, the pressure oscillations upstream of the combustor at P3 are about one-third the size of those inside 

the combustion at P4. However, the same spectral information about the instability is present. A similar 

relationship between P3 and P4 was seen at multiple operating conditions. An additional item of interest 

seen in Figure 12 is that the relative phase of sensed P3 and P4 (recorded at the exact same time) are not 

the same. This phase difference points to the assumed mode shape of the combustion instability of 

roughly a half-wave with a node in the vicinity of the fuel injector. Since the ASPAC controller sets a 

relative phase based on observed instability suppression, the phase difference between P3 and P4 should 

not negatively impact the controller. Based on these results, it was decided to attempt to repeat some of 

the earlier shown results, but instead using P3 as the feedback sensor to the ASPAC controller. 

Figure 13 shows a combustion instability control result using P3DynA as feedback and modulating fuel 

through the main stage injector. For consistency with the presentation of earlier results, P4DynDn is shown 

as the “instability pressure” even though P3DynA was used as the controller feedback. As shown in the 

spectral plots on the left, when the controller is on, the amplitude of the pressure oscillations at the 

instability frequency is reduced by roughly 85 percent (from 0.43 to 0.06 psi) from its uncontrolled value. 

The time history plot on the right shows a decrease in the peak-to-peak pressure oscillations from about 

2.0 psi to about 0.8 psi. Note that the uncontrolled peak pressure here is smaller than in results shown 

earlier due to the continued decrease in instability amplitude versus time discussed earlier. This was 

considered to be a promising demonstration of the feasibility of using P3 for controller feedback. 

However, it should be noted that how well the pressure oscillations at P3 represent the instability pressure 

oscillations inside the combustor, and thus these results, may be configuration dependent. Further 

investigations with other configurations are planned. 
NASA/TM—2012-217617 16 

Figure 12.—Comparison of the measured pressure oscillations upstream of the 

combustor (P3DynA) and inside the combustor (P4DynDn) shows that the combustion 

instability pressure oscillations can be sensed in the lower temperature 

environment upstream of the combustor. 

Figure 13.—Combustor pressure amplitude spectrum (left) and time 

history (right) shows the combustion thermoacoustic instability 

being suppressed by the application of closed-loop control using 

combustor inlet pressure, P3DynA, as controller feedback.

0

0.2

0.4

0.6

0.8

P3DynA_psi

-2

-1

0

1

2

psi

101 102 103 0

0.2

0.4

0.6

0.8

P4DynDn_psi

Frequency, Hz

1 1.01 1.02 1.03 1.04 -2

-1

0

1

2

psi

Time, sec

b) With control

a) Without control

101 102 103 0

0.1

0.2

0.3

0.4

0.5

Combustor Pressure, psi

Frequency, Hz

0 0.01 0.02 0.03 0.04 -1.5

-1

-0.5

0

0.5

1

1.5

psi

Time, sec

101 102 103 0

0.1

0.2

0.3

0.4

0.5

Combustor Pressure, psi

Frequency, Hz

0 0.01 0.02 0.03 0.04 -1.5

-1

-0.5

0

0.5

1

1.5

psi

Time, sec
NASA/TM—2012-217617 17 

Combustion Instability Control Using Pilot Fuel Injector Modulation 

The high-frequency fuel valve was also installed in the combustor pilot fuel line, which typically 

carries 10 to 20 percent of the overall combustor fuel flow. The valve installation was very similar to that 

for the main stage fuel injector. There was slightly less fuel line length, approximately 25 in., between the 

fuel valve and the fuel injector tip than with the main stage injector. Open-loop, sinusoidal valve 

command variations were again provided to the valve to observe the response of combustor pressure to 

the pilot fuel injector perturbations. As with the main stage fuel injector, the commanded frequency and 

voltage level to the valve were varied and combustor pressure P4DynDn monitored. Fixed frequency and 

sweep frequency perturbations of the valve showed very little response in the combustor pressure. Unlike 

with the main stage fuel injector modulations, no response was seen in the spectral data at the perturbation 

frequency. Further, there was no apparent interaction with the instability at frequencies near the instability 

frequency. There was also very little response in the valve exit pressure to the valve perturbations, further 

indicating small, if any, fuel flow perturbations into the combustor. This essentially validated the 

suspicions from dynamic characterization testing of the valve prior to combustion testing that the valve, 

designed for large Flow Number applications, would have difficulty modulating the flow through a small, 

pilot fuel injector (Ref. 25). It’s possible that even large pilot fuel modulations would be ineffective at 

instability suppression, although this scenario could not be investigated. Several optimizations of the fuel 

feed line and valve configuration were attempted with little improvement in the response. None-the-less, 

it was decided to attempt closed-loop control with the ASPAC controller modulating the pilot fuel. 

Unfortunately, as described earlier, the combustor instability behavior continued to become harder to 

elicit, and the FAR required to get an instability to occur exceeded the safe operational limits of the 

facility. This combined with schedule pressure from other users of the facility ended this combustion 

control test cell entry and thus prevented further investigation of combustion instability control using pilot 

fuel modulation. This remains an important future objective. 

Closed-Loop Combustor Data for Development of Combustion Control Simulations 

The last research objective was to obtain sufficient combustion instability control data to allow 

extension of the previously described combustion instability simulation to include closed-loop control. 

Unfortunately, as described earlier, the instability behavior of the combustor was changing from run to 

run. As a result, closed-loop instability control results were only obtained at a single operating condition. 

Further, for this condition, the combustor was being operated differently than in the 2006 testing. Most 

notably, the FAR was higher and the injector ΔP/P lower than in any previous testing. Even without 

control applied, this requires some investigation on how the Sectored 1-D simulation of the combustion 

instability described in Reference 23 can be updated to capture this new behavior. When/if the simulation 

is suitably updated to capture the uncontrolled combustor instability behavior, the next effort would be to 

incorporate the ASPAC controller and try to match the controlled behavior. This is the subject of ongoing 

work and is not reported on further in this paper. 

Concluding Remarks 

Lean combustion concepts for aircraft engine combustors are prone to combustion instabilities. 

Mitigation of instabilities is an enabling technology for these low-emissions combustors. Active control 

of combustion instability has been demonstrated on a prototype low-emissions aircraft engine combustor 

installed in the NASA Glenn CE-5 flametube rig. Dynamic pressure sensors provided the feedback signal 

for control. A high-frequency fuel modulation valve perturbed the fuel into the combustor. The NASA 

GRC-developed Adaptive Sliding Phasor Averaged Control (ASPAC) algorithm, utilizing a highfrequency fuel modulation valve, was able to detect and suppress a thermoacoustic combustion instability 

occurring in the prototype combustor. The ASPAC algorithm determines the frequency at which the 

instability occurs and modulates the fuel at a phase that interferes with the instability. The phase is 
NASA/TM—2012-217617 18 

constantly adapted in order to maintain control as conditions in the combustor change. It was also 

demonstrated that the controller can prevent the instability from occurring while combustor operation was 

transitioning from a stable to a normally unstable, high-power condition, thus enabling the high-power 

operation. A pressure sensor inside the combustor was initially used for feedback. A pressure sensor in 

the less-harsh environment upstream of the combustor was also able to sense combustion instability 

inside the combustor and was successfully used as a feedback sensor for the controller. Successful 

instability control was demonstration with the controller modulating the fuel to the combustor’s main 

stage fuel injector. Testing to demonstrate instability control with the controller modulating the fuel to the 

combustor’s pilot fuel injector was investigated, but was unsuccessful due to inadequate fuel modulation 

strength. Future plans are to extend an existing simulation of the uncontrolled combustion instability to 

include the controlled case. Future work is also planned to develop fuel actuators sized for pilot injectors 

and to apply combustion instability control technologies via pilot fuel modulation to increasingly 

advanced lean-burn combustors. 

References 

1. Lefebvre, A.H.: “Gas Turbine Combustion,” 2nd edition, Taylor & Francis, 1999. 

2. Zinn, B.T.; and Lieuwen, T.C.: “Combustion Instabilities: Basic Concepts,” Combustion Instabilities 

in Gas Turbine Engines: Operational Experience, Fundamental Mechanisms, and Modeling, edited 

by T. Lieuwen and V. Yang, Progress in Astronautics and Aeronautics, AIAA, Reston, VA, 2005, 

Chap 1, pp. 3-26. 

3. Mongia, H.; Held, T.; Hsiao, G; Pandalai, R.: “Challenges and Progress in Controlling Dynamics in 

Gas Turbine Combustors,” Journal of Propulsion and Power, Vol. 19, No. 5, p. 822-829, SeptemberOctober 2003. 

4. Lieuwen, T.C., and Yang, V., Combustion Instabilities in Gas Turbine Engines: Operational 

Experience, Fundamental Mechanisms, and Modeling, Progress in Astronautics and Aeronautics, 

AIAA, Reston, VA, 2005, Chaps. 2-8. 

5. Schadow, K.; Yang, V., Culick, F.; Rosfjord, T.; Sturgess, G.; Zinn, B.: “Active Combustion Control 

for Propulsion Systems,” AGARD Report 820, September 1997. 

6. Zinn, B.T.; Neumeier, Y.: “An Overview of Active Control of Combustion Instabilities,” AIAA Paper 

97-0461, January 1997. 

7. McManus, K.R.; Magill, J.C.; Miller, M.F.; Allen, M.G.: “Closed-Loop System for Stability Control 

in Gas Turbine Combustors,” AIAA Paper 97-0463, January 1997. 

8. Annaswamy, A.M.; El Rifai, O.M.; Fleifil, M.; Hathout, J.P.; Ghoniem, A.F.: “A Model-based 

Self-tuning Controller for Thermoacoustic Instability,” in Combustion Science and Tech., Vol. 135, 

pp. 213-240, 1998. 

9. Hibshman, J.R.; Cohen, J.M.; Banaszuk, A.; Anderson, T.J.; Alholm, H.A.: “Active Control of 

Combustion Instability in a Liquid-Fueled Sector Combustor,” International Gas Turbine & 

Aeroengine Congress and Exhibition, June 1999. ASME Paper 99-GT-215. 

10. Allgood, D.; Campos-Delgado, D.U.; Acharya, S.; and Zhoo, K.: “Acoustic Control of 

Thermoacoustic Instabilities Using Experimental Model-Based Controllers,” Proceedings of ASME 

Turbo Expo, New Orleans, LA, 2001. 

11. Johnson, C.E.; Neumeier, Y.; Nuemaier, M.; Zinn, B.T.: “Demonstration of Active Control of 

Combustion Instabilities on a Full-Scale Gas Turbine Combustor,” Proceedings of ASME Turbo 

Expo, New Orleans, LA, 2001. 

12. Kiel, B.: “Review of Advances in Combustion Control, Actuation, Sensing, Modeling and Related 

Technologies for Air Breathing Gas Turbines,” 39th Aerospace Sciences Meeting and Exhibit, Reno 

NV, AIAA-2001-0481, January 2001. 

13. Barooah, P.; Anderson, T.J.; and Cohen, J.M.: “Active Combustion Instability Control with Spinning 

Valve Actuator,” ASME Turbo Expo 2002, Amsterdam, The Netherlands. ASME Paper GT-2002-

30042, June 2002. 
NASA/TM—2012-217617 19 

14. Park, S.; Ghosh, A.; Diao, Q.; and Ye, K.H.: “Optimizing Combustion Instability Suppression Using 

Secondary Fuel Injection,” 47th Joint Propulsion Conference & Exhibit, San Diego, CA, AIAA-2011-

5627, August, 2011. 

15. Goeke, J.L., and Overman, N.T.: “Active Combustion Control System Rig Testing,” 46th Joint 

Propulsion Conference & Exhibit, Nashville, TN, AIAA-2010-6865, July 2010. 

16. DeLaat, J.C.; Chang, C.T.: “Active Control of High Frequency Combustion Instability in Aircraft 

Gas-Turbine Engines,” 16th International Symposium on Airbreathing Engines, Cleveland, OH, 

ISABE-2003-1054, NASA/TM—2003-212611, September 2003. 

17. Le, D.K.; DeLaat, J.C.; Chang, C.T.; Vrnak, D.R.: “Model-Based Self-Tuning Multiscale Method for 

Combustion Control,” 41st Joint Propulsion Conference and Exhibit, Tucson, Arizona, AIAA-2005-

3593, July 2005. 

18. Kopasakis, G.; DeLaat, J.C.; Chang, C.T.: “Adaptive Instability Suppression Controls Method For 

Aircraft Gas Turbine Engine Combustors,” Journal of Propulsion and Power, Vol. 25, No. 3, May–

June 2009, pp. 618-627. 

19. Cohen, J.M.; Proscia, W; and DeLaat, J.C.: “Characterization and Control of Aeroengine Combustion 

Instability: Pratt & Whitney and NASA Experience,” Combustion Instabilities in Gas Turbine 

Engines, Operational Experience, Fundamental Mechanisms, and Modeling, Progress in Astronautics 

and Aeronautics series, AIAA, Reston, VA, 2005, Chap 6, p. 113-145. 

20. Okojie, R.S.; DeLaat, J.C.; Saus, J.R.: “SiC Pressure Sensor for Detection of Combustor 

Thermoacoustic Instabilities,” Proceedings of the 13th International Conference on Solid-State 

Sensors, Actuators and Microsystems, Seoul, Korea, Volume 1, p. 470-473, June 2005. 

21. Paxson, D.: “A Sectored-One-Dimensional Model for Simulating Combustion Instabilities in Premix 

Combustors,” 38th Aerospace Sciences Meeting & Exhibit. AIAA-2000-0313, NASA TM-1999-

209771, January 2000. 

22. Tacina, R.; Wey, C.; Laing, P.; Mansour, A.: “A Low NOx Lean-Direct Injection, Multipoint 

Integrated Module Combustor Concept for Advanced Aircraft Gas Turbines,” NASA/TM—2002-

211347, April 2002. 

23. DeLaat, J.C.; Paxson, D.E.: “Characterization and Simulation of the Thermoacoustic Instability 

Behavior of an Advanced, Low Emissions Combustor Prototype,” 44th Joint Propulsion Conference 

and Exhibit, Hartford, Connecticut, AIAA-2008-4878, NASA/TM—2008-215291, July 2008. 

24. Saus, J.R.; Chang, C.T.; DeLaat, J.C.; Vrnak, D.R.: “Design and Implementation of a Characterization 

Test Rig for Evaluating High Bandwidth Liquid Fuel Flow Modulators,” Presented at the 45th Joint 

Propulsion Conference and Exhibit, Denver, CO, AIAA-2009-4886, NASA/TM—2010-216105, 

August 2009. 

25. Saus, J.R.; Chang, C.T.; DeLaat, J.C.; and Vrnak, D.R.: “Performance Evaluation of a High Bandwidth 

Liquid Fuel Modulation Valve for Active Combustion Control,” 50th AIAA Aerospace Sciences 

Meeting, Nashville, TN, AIAA-2012-1274, January 2012 . 
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01-07-2012 

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Technical Memorandum 

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4. TITLE AND SUBTITLE

Active Combustion Control for Aircraft Gas-Turbine Engines-Experimental Results for an 

Advanced, Low-Emissions Combustor Prototype 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

DeLaat, John, C.; Kopasakis, George; Saus, Joseph, R.; Chang, Clarence, T.; Wey, Changlie 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 984754.02.07.03.19.04 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

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E-18205 

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NASA 

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NASA/TM-2012-217617 

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

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

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Lean combustion concepts for aircraft engine combustors are prone to combustion instabilities. Mitigation of instabilities is an enabling technology for these low-emissions 

combustors. NASA Glenn Research Center’s prior activity has demonstrated active control to suppress a high-frequency combustion instability in a combustor rig designed to 

emulate an actual aircraft engine instability experience with a conventional, rich-front-end combustor. The current effort is developing further understanding of the problem 

specifically as applied to future lean-burning, very low-emissions combustors. A prototype advanced, low-emissions aircraft engine combustor with a combustion instability 

has been identified and previous work has characterized the dynamic behavior of that combustor prototype. The combustor exhibits thermoacoustic instabilities that are related 

to increasing fuel flow and that potentially prevent full-power operation. A simplified, non-linear oscillator model and a more physics-based sectored 1-D dynamic model have 

been developed to capture the combustor prototype’s instability behavior. Utilizing these models, the NASA Adaptive Sliding Phasor Average Control (ASPAC) instability 

control method has been updated for the low-emissions combustor prototype. Active combustion instability suppression using the ASPAC control method has been 

demonstrated experimentally with this combustor prototype in a NASA combustion test cell operating at engine pressures, temperatures, and flows. A high-frequency fuel 

valve was utilized to perturb the combustor fuel flow. Successful instability suppression was shown using a dynamic pressure sensor in the combustor for controller feedback. 

Instability control was also shown with a pressure feedback sensor in the lower temperature region upstream of the combustor. It was also demonstrated that the controller can 

prevent the instability from occurring while combustor operation was transitioning from a stable, low-power condition to a normally unstable high-power condition, thus 

enabling the high-power condition.

15. SUBJECT TERMS

Aircraft engines; Combustion stability; Combustion control; Engine control 

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