[american institute of aeronautics and astronautics 44th aiaa aerospace sciences meeting and exhibit...

American Institute of Aeronautics and Astronautics

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Acoustic Interactions of a Pulse Detonation Engine Array with a Gas Turbine

Nicholas Caldwell*, Aaron Glaser†, and Ephraim Gutmark‡

University of Cincinnati, Cincinnati, Ohio, 45221

Peak pressure attenuation data is presented for pulse detonation combustor (PDC) firing into a single stage axial flow turbine. Cases are presented for both single PDC tubes and for an annular array of six PDC tubes. Pressure attenuation is characterized for a wide array of PDC operating parameters, including fill fraction, equivalence ratio, nitrogen dilution percentage, and firing frequency. Effects of additional cold air flow mixed into the PDC exhaust prior to the turbine inlet are also quantified. It is shown that certain PDC operating conditions lead to a maximum peak pressure attenuation, which should correspond to the greatest energy extraction by the turbine from the PDC exhaust flow.

Nomenclature β = bypass ratio f = firing frequency ff = fill fraction φ = equivalence ratio p = peak pressure pref = reference pressure (20 µPa) SPL = sound pressure level TL = peak pressure attenuation (transmission loss) 1 = subscript representing location upstream of PDC tube exit 2 = subscript representing location upstream of turbine inlet 3 = subscript representing location downstream of turbine exit

I. Introduction HE concept of integrating an array of pulse detonation engines (PDEs) with a conventional gas turbine engine as the primary combustion system is piquing the interest of the research community due to the

potentially significant performance benefits that are implied. Taking advantage of the detonative mode of combustion promises greater thermal efficiency over its deflagrative counterpart, resulting in reduced specific fuel consumption and greater specific thrust. The difficulty in incorporating a pulse detonation array as a combustor is the innate unsteadiness associated with the PDE. Extracting power from such a time-dependent cycle is problematic due to the fact that gas turbines have always been designed for steady flow inlet conditions. Questions also arise regarding the high temperatures and the impulsive nature of the pulse detonation cycle leading to reduced lifetimes of turbine blades. On top of this, the pulse detonation combustor (PDC) generates high acoustic emissions compared to a stable combustion process. Before the future advent of the pulse detonation combustor, these issues and many others must be addressed through extensive experimental, theoretical, and computational research. The aim of this current work is to characterize the noise associated with a pulse detonation combustor. Due to the large pressure ratio across the detonation wave, this form of combustion is inherently much louder than a typical form of burning. For example, a typical gas turbine combustor generates a sound

* Graduate Research Assistant, Department of Aerospace Engineering, ML0070, AIAA Student Member. † Graduate Research Assistant, Department of Aerospace Engineering, ML0070, AIAA Student Member. ‡ Professor and Ohio Eminent Scholar, Department of Aerospace Engineering, ML0070, AIAA Associate Fellow.

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44th AIAA Aerospace Sciences Meeting and Exhibit9 - 12 January 2006, Reno, Nevada

AIAA 2006-1233

Copyright © 2006 by Nicholas Caldwell. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.


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Figure 2. University of Cincinnati PDC/Turbine Facility

Figure 1. PDE Tube Number Designations (detonation wave

travels into the page)

Turbine Mean Radius

#4 -- 240º

#5 -- 180º

#6 -- 120º #1 -- 60º

#2 -- 0º Turbine Mean Radius

#3 - 300º

pressure level (SPL) on the order of 170-180 dB1,2 in the instability range and a much lower level during a regular stable operation, whereas a PDE may exhibit an SPL closer to 200 dB at the tube exit plane3. Although the overall sound pressure level (OASPL) reduces significantly to the range of 120-150 dB at a distance from the PDE, the noise problem that must be solved is reducing the pressure that is exiting the system4,5. Previous work by Allgood et al has shown that the OASPL is largely affected by the operating parameters of the PDE6. Confining this combustion inside a burning chamber undoubtedly helps reduce this excessive noise; however, a pulse detonation combustor will continue to be louder than its conventional counterpart. Previous measurements have shown that expanding the PDE exhaust flow into a larger chamber prior to the turbine inlet can reduce its peak pressure level by about 7 dB7. This certainly depends on the size of the expansion and the operating parameters of the PDE, but is useful in showing the order of magnitude of the reduction. The question then remains regarding how effectively the turbine blades can be used to further reduce the noise issuing from the PDCs to levels acceptable for human ears. By extracting power from the PDC exhaust flow, the amount of energy released by the system in the form of acoustics should be lessened8. The power portion of this double benefit is presented elsewhere9.

II. Experimental Setup The baseline setup at the University of Cincinnati is a single tube pulse detonation engine with a 1”

inner diameter and an internal length of 25”. Three valves feed the oxygen portion of the oxidizer to the system, while two additional valves feed in a nitrogen dilution. A final valve supplies the ethylene fuel. An MSD automotive ignition system is utilized as the spark source, delivering up to 40 mJ of energy per spark. This baseline system allows for a wide range of testing, including pressure diagnostics along the length of the tube, shadowgraph/schlieren visualizations, particle image velocimetry (PIV), and a host of other diagnostics. In translating this system to a PDC/Turbine facility, this single tube pulse detonation engine was duplicated into an annular array of pulse detonation combustors. An array of six PDCs has been designed and fabricated, along with a facility to integrate them as a combustion device feeding into a low pressure turbine. Details of this system have been presented in previous work, so only the most important features will be

repeated here7. Each PDC has an inner diameter of 1” and an internal length of 25”. Currently the maximum firing frequency of each is 20 Hz, using ethylene as the fuel and a variable mixture of oxygen and nitrogen as the oxidizer. Every valve and spark plug

in the system is independently controlled, allowing for every PDC to be fired at any time relative to the others, and at a different fill fraction and frequency if desired. In this manner, a multitude of firing patterns may be investigated. The PDC array is annular, with a diameter of 5” to match the mean diameter of the downstream turbine blades. Each PDC can pass a mass flow rate of approximately 0.04 lbm/sec at 20 Hz. The PDE tube number designations and azimuthal angle relative to tube 2, are shown in Figure 1.

The turbine being used for this experiment


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Figure 3. University of Cincinnati PDC/Turbine Facility Schematic, showing locations of pressure transducers

Figure 4. Sample Pressure Traces

is an Allied Signal model JFS-100-13A axial flow power turbine, designed for a mass flow rate of 1.6 lbm/sec. The turbine has one removable stator section and one rotor section. The rotor blade height is 1.15”, closely matching the PDC diameters, and the mean diameter of the blades is 5”. At its design point, the turbine produces 90 shp at 60,400 RPM. A bypass air system has been installed to supply the remainder of the mass flow rate that cannot by supplied by the PDCs. It can deliver over 0.75 lbm/sec of air at a stagnation pressure of 80 psi. This air is not preheated; hence the work produced by the turbine is expected to be much less than its intended application. The experimental facility is shown in Figure 2.

PDC #5 (in the 180º position) contains four pressure transducer ports, in which are placed PCB model CA113A dynamic pressure sensors with a range from 0 to 3000 psi. Slightly downstream of the PDE exhaust (about 2.5”), after the expansion into the integrating chamber, a PCB model 102A06 dynamic pressure sensor with a range of 0 to 500 psi measures the effect of the expansion from the detonation tube and the addition of the bypass air. Prior to the stator section, at a location 6.125” from the PDE exhausts, ports are available for four static pressure transducers and four type K thermocouples. A final pressure transducer, PCB model 102A07 dynamic pressure sensor, is located immediately behind the rotor section of the turbine. This transducer quantifies the effect of the detonation wave structures passing through the rotor and stator blades, as well as after the turbine expansion. The range on this sensor is 0-50 psi.

Measurement of the peak pressures at each of these locations provide insight into the mechanisms of noise attenuation due to the flow path associated with this pulse detonation combustor integration. First, the effect of the PDC exhaust expanding into the integrating chamber is quantified and understood as a function of the system operating parameters. Next, the peak pressure attenuation of the detonation wave passing through the rotor and stator blades is analyzed. Under conditions of varying fill fractions, firing frequencies, nitrogen dilutions, equivalence ratios, bypass ratios, the maximum peak pressure attenuation conditions are of interest. This will hopefully correspond to conditions under which the most power is extracted from the pulse detonation flow. Measurements are also made of the effect on the flow of the expansion directly following the rotor blades in the turbine by relocating the most downstream pressure sensor. The main locations of these pressure measurements are shown in the schematic view of the facility in Figure 3.

III. Peak Pressure Attenuation Analysis The attenuation measurements presented here reflect the reduction in the peak sound pressure level as evaluated at different locations in the PDC/Turbine flow path. The sound pressure level (SPL) at any location in the system is given by the following equation10:

Flow Direction

p1

p2 p3

=

=

refref pp

ppSPL log20log10 2

2

p4

0 0.5 1 1.5 2

x 104

-200

0

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400

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x 104

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ssur

e, p

si

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x 104

-2

0

2

4

Relative Time, samples

p1

p2

p4

x104


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In this equation the reference pressure is taken to be 20 µPa, commonly used in acoustic work as the quietest sound at 1000 Hz that can be heard by the human ear. The peak pressure attenuation from one point to another is then given by the difference in the sound pressure level at the two points based on the respective peak pressures. For example, the transmission loss computed from location 2 to location 3, TL23, is given by the following equation, eliminating the need altogether for the reference pressure:

A sample pressure trace for the location upstream of the turbine blades is shown in Figure 4. The sharp peak that is seen is due to the pressure rise across the leading shock wave, and is followed by a much longer duration blowdown cycle during which the pressure in the tube is relaxed back to atmospheric. The length of this blowdown time is dependent on both the PDC parameters and any downstream boundary conditions which may affect the relaxation, but even at a 20 Hz operating frequency, the blowdown portion of the cycle is well completed before the next detonation event occurs. The noise resulting from the blowdown portion of the cycle is only a fraction of that due to the peak pressure, and therefore is not the focus of the current work.

IV. PDC Operating Parameter Results The following section describes the effects on peak pressure attenuation that result from the

measurement of a single pulse detonation tube firing into the turbine. For these tests the PDC #5 (180º position) is used, and fired for at least five detonations, the number of cycles determined by Rasheed et al as required to achieve a steady peak pressure measurement11. The downstream pressure transducer is placed at the exit plane of the turbine (the p4 location), and hence these results quantify the transmission loss due to the stationary rotor and stator vanes, as well as the loss due to the expansion in the turbine following the rotor. The majority of this testing was conducted with a stationary rotor, but to quantify the effect of adding bypass flow into the turbine (Section IV-E), the testing was performed at various rotor speeds.

A. Fill Fraction The fill fraction, defined as the ratio of the volume of the PDC tube filled with a detonable mixture

prior to combustion to the total volume of the tube, can be seen to have a significant impact on how much attenuation is experienced by the detonation waves. Increasing the fill fraction has been shown to be linearly proportional to the amount of thrust produced by a PDE until the point at which the tube is completely filled12. Further increasing the fill fraction generates a free cloud of detonable mixture outside the tube exit. In an unconfined environment downstream of the PDE exhaust, this excess fuel and oxidizer would go to waste or lead to a secondary external detonation, and hence contributing nothing to the PDE thrust13. However, in a confined volume this spillage would maintain the detonation wave for a short distance from the PDE exhaust plane before it decays. Therefore it is expected that larger fill fractions would lead to stronger detonation waves, and it makes sense that a strong shock wave would be less affected by an obstacle than a weak one. Previous work performed by Schauer et al has found that lower fill fraction cases have weaker interactions with the turbine blades than high fill fraction cases14, and similar results are presented here in Figure 5. It can be seen that for full detonation conditions, i.e. fill fraction cases greater

=

−

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2323223 log20loglog20

pp

pp

ppSPLSPLTL

refref

Figure 5. Effect of Fill Fraction on Peak Pressure Reduction

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Figure 7. Effect of Nitrogen Dilution on Peak Pressure Reduction

than 0.77, pressure attenuation decreases with fill fraction, while the peak pressure that defines the strength of the detonation wave is increasing. It must be noted here that the first two data points shown were incomplete detonation events, but it is interesting to see that they are significantly more affected by the turbine than full detonation waves. The greatest attenuation is achieved at a fill fraction of 0.77, with a magnitude of 28.92 dB. Making the trend lines include only those cases in which full detonations were achieved, both the peak pressure and the reduction exhibit a linear trend in opposite directions.

B. Equivalence Ratio The strength of the initial detonation

wave is known to depend greatly on the equivalence ratio of the gaseous mixture prior to combustion. In much the same way as the fill fraction, reducing the equivalence ratio effectively lessens the chemical potential of the mixture before ignition. In Figure 6 the pressure axis shows that the strongest detonation wave is achieved at stoichiometric conditions, and that the attenuation drops off with increasing equivalence ratio. With the strength of the detonation wave maximized, the effect of passing through the blades of the turbine will be least potent. For this reason, the stoichiometric condition leads to the smallest attenuation of the peak pressure.

C. Nitrogen Dilution As mentioned before, the oxidizer in this PDE system is diluted with nitrogen to ensure detonation

within the small diameter tube at temperatures that can sustain the detonation process for long periods of time without auto ignition. Reducing the amount of nitrogen with which the oxidizer is diluted has the effect of increasing the temperature of the detonation, but also results in a stronger detonation wave with an increased wave velocity. The effect seen in Figure 7 is very slight, but a positive trend is seen. The strongest wave, with the strongest wave velocity, occurs at the lowest percentage of nitrogen dilution, and hence this is the end of the curve at which the lowest pressure attenuation is achieved. Wave speed measurements as a function of nitrogen dilution percentage for this mixture show a variation of less than 200

m/sec for the entire range examined here, and the peak pressure for this range shows a good degree of scatter and only a slight reduction with nitrogen dilution percentage, so it makes sense that the attenuation trend is as slight as it appears.

D. Firing Frequency The final operating variable that could be varied in the PDC alone was the frequency at which the PDC

was fired. Results of this testing are shown in Figure 8. By firing the combustor more frequently, less time is available between detonation events for the system to exhaust combustion products and to prepare for the next cycle. Hence, increasing the firing frequency leads to higher temperatures and pressures in the system directly upstream of the turbine prior to the start of the next detonation event. The addition of more

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Figure 6. Effect of Equivalence Ratio on Peak Pressure Reduction

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Figure 9. Effect of Turbine Speed on Peak Pressure Reduction

Figure 8. Effect of Firing Frequency on Peak Pressure Reduction

dramatic upstream conditions with increasing firing frequency must be accountable for the decreasing pressure attenuation, i.e. the strengthening of the detonation waves. It may be that at the higher frequencies the system is reaching a state of thermal stability at a quicker pace than at the lower frequencies, and reaches a point at which the detonation waves are stronger in a shorter span of time. This is something that must be investigated in the future to determine if a better criterion for steady pressure attenuation measurements can be found.

E. Bypass Ratio Similar to that of a typical gas turbine

engine, the bypass ratio β is defined as the ratio of the mass flow rate of cold air added into the system downstream of the PDC exit plane to the mass flow rate through the PDC tubes. The addition of this bypass air has the effect of spinning the rotor at a nearly constant velocity prior to firing the detonation tubes. For the short bursts of detonations analyzed here (typically five detonations to achieve a steady peak pressure), it is assumed that the angular speed of the rotor is not greatly affected. Hence, each bypass ratio is associated with a constant angular speed of the rotor. Rasheed

et al15 measured the attenuation at variable turbine speed, the results of which are overlaid with the data collected for this test, shown in Figure 9. The significant variation between the results of the two groups may be attributed to several differences between to two experimental facilities including the use of different turbines. A discrepancy in blade solidities could explain why more attenuation would be seen in one case over another. The turbine used at the University of Cincinnati has a blade solidity of approximately 2.0, whereas that used at GEGR is closer to 1.51. Further discrepancies between this data and that of GEGR arise due to a difference in measurement techniques, including pressure sensor locations. According to Saravanamuttoo et al16, shock

losses (and hence peak pressure reduction) in a steady flow case should increase with blade solidity. This statement seems as though it should carry over to the unsteady case. A higher solidity turbine would generally have a smaller pitch between adjacent blades, and hence obstruct the flow to a larger degree. The present data was taken at a fill fraction of 1 under stoichiometric conditions and maximum nitrogen dilution percentage, 40%. In this case also the downstream pressure transducer was placed immediately behind the spinning rotor, so the effect of the expansion in the turbine to be discussed later must also be considered. The facility employed by GEGR operates on an ethylene-air mixture, utilizing a greater percentage of nitrogen in the oxidizer than the facility at the University of Cincinnati. The tests compared in the figure were run at a frequency of 20 Hz, with a fill fraction varying between 0.9 and 1.4. The strength of the detonation waves do not vary in the cases presented here, so the effect of the turbine spinning is itself responsible for the variation in attenuation. Figure 9 shows that with the turbine spinning there is a significant increase in the amount of attenuation that is achieved over the stationary case. This result is confirmed in previous testing done by GEGR17. However, there seems to be an optimum turbine speed that attenuates the signal more than the others. After this speed is reached, the attenuation begins to drop off.

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Rasheed et al (Ref. 15)University of Cincinnati


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Figure 11. Transmission Loss Due to Flow Turning and Expansion in Turbine – Variation with Firing

Frequency

Figure 10. Transmission Loss Due to Flow Turning and Expansion in Turbine – Variation

with Fill Fraction

Until the drop off point, this trend matches that found by Rasheed, which shows a monotonically increasing attenuation with turbine speed. The difference is turbines and facilities is likely to account for the magnitude discrepancies. It may also be that the most recent data presented here is above the range tested by Rasheed et al15.

V. Turbine Flow Turning and Expansion Effect The particular turbine used in this study has an axial flow inlet but exhausts radially. This requires both

a turning of the flow, and an expansion back to the atmosphere. By measuring the difference between p3 and p4 the change in the peak pressure reduction achieved by this turning and expansion may be determined. These results are presented here for two cases discussed previously; the effects of the fill fraction and firing frequency. It is shown that the expansion effect is largely dependent on PDC parameters, and in some cases can account for 23% of the total transmission loss.

A. Fill Fraction Figure 10 shows the peak pressure reduction due to the flow turning and expansion within the turbine.

A similar trend is seen as that of the fill fraction alone in Figure 10. The low fill fraction cases (0.38 and 0.58) do not represent full detonation events, and the remaining data points seem to show a linear trend with the peak pressure attenuation decreasing with increasing fill fraction. At a fill fraction of 0.77, the largest attenuation is seen due to the expansion effects, 6.5 dB, while the overfilled case exhibits a reduction of less than 1 dB. Since the pressure leaving the rotor section varies with fill fraction, it should be expected that the pressure drop that ensues due to the exhaust geometry of the turbine should not be constant, much in the same way that a solid nozzle is specifically designed for one inlet pressure. It is interesting to note here that for the overfilled case, the stator vanes and rotor blades are the source of the majority of the pressure reduction, whereas for lower fill fractions, a larger percentage of the reduction is due to the turbine expansion and flow turning. At a fill fraction of 1.35, the expansion accounts for only 3% of the attenuation, while at a fill fraction of 0.77,

the expansion is responsible for 23% of the transmission loss.

B. Firing Frequency The effect on the peak pressure

reduction due to flow turning and expansion is found to increase with firing frequency, as shown in Figure 11. There is a definite positive trend, and from the low end of the frequency curve to the high end, a significant increase in transmission loss of over 4 dB. Firing at only 1 Hz, the expansion and turning attenuate the pressure by only 3.5 dB, whereas at the maximum firing frequency, the expansion results in a 7.5 dB reduction. This suggests that at lower

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Figure 12. Separated Effects of the Transmission Loss Components

firing frequencies, more of the pressure is attenuated by the stator vanes and rotor blades than in the higher frequency cases. Since firing a at a lower frequency allows more time for the completion of the blowdown cycle, it is reasonable to assume that the flow has had more of a chance to expand at the lower frequency conditions. Hence, the stator vanes and rotor blades have ample opportunity to affect the flow before the next detonation event occurs. However, when this next detonation event takes place before the blowdown portion of the cycle is complete, the back pressure generated inside the system and inside the turbine affects the pressure directly behind the rotor, and the flow undergoes an expansion in the turbine that is more significant than in the lower pressure cases.

VI. Separation of Stator and Rotor Effects Results presented in this portion look at the effect of removing the entire turbine from the system except

for the stator section. The downstream pressure transducer is then placed approximately 1” downstream of the trailing edge of the stator vanes, and measures the peak pressure reduction that is achieved by the stator alone. From this and previous data collected, the individual effects of the stator and rotor can be determined, with the losses due to the flow expansion and turning already known. A plot of this separation for various fill fraction cases is shown in Figure 12. It can be seen that the expansion accounts for only a small portion of the total pressure attenuation, while the greatest losses occur in the stator section, especially at the highest fill fraction case, where the stator accounts for 68% of the total transmission loss. With increasing fill fraction, the pressure attenuation occurs to a greater degree in the stator section than in the rotor. In fact, the attenuation in the rotor diminishes as the fill fraction increases. This is the same trend that was found previously for the fill fraction effect. On the other hand, the stator exhibits a totally different trend, namely an increase in peak pressure reduction with fill fraction. It may be possible to relate these trends to the pressures upstream of each station, and this merits a more detailed investigation. The pressure loss mechanisms across the stator vanes and rotor blades must be similar, but Figure 12 suggests that these mechanisms are different enough to act on detonation waves of varying strength in dissimilar manners. However, as found before the overall effect of the turbine components is lessened attenuation with increasing fill fraction. Knowing how each component of the turbine attenuates the pressure may be useful information in the design of a turbine suitable to a pulse detonation combustor in that the blade rows may need to be loaded more evenly to achieve a better reaction. In the case that a steady flow turbine such as is used in this facility were to be used downstream of a pulse detonation combustor, a slightly reduced fill fraction would need to be used to equally divide the work between the stator and the rotor sections.

VII. Detonation Wave Strength Analysis The peak pressure of the detonation wave exiting the PDC is a function of all of the operating

parameters presented in this work. As such, one particular peak pressure at this location may be achieved through any combination of these variables. For example, a peak pressure in the range of 450 psi may be achieved by either incompletely filling the tube or reducing the equivalence ratio from stoichiometric. Varying the amount of nitrogen diluting the oxidizer may also give a peak pressure in this same range. This being the case, it is probable that the PDC exit peak pressure, p1, may be the sole determining factor in how much the peak pressure is attenuated through the turbine. The strength of the detonation wave is extremely dependent on this parameter, so it is likely that all data can collapse onto a single graph. In fact, Figure 13 shows that all PDC operating parameters collapse onto nearly the same line. All parameters show a decreasing trend with increasing PDC exit pressure, implying that as the initial strength of the detonation

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Figure 13. Peak Pressure Reduction as a Function of PDC Exit Pressure for Various PDC Operating

Parameters

wave increases, it is more capable of sustaining itself across the obstacle presented by the turbine stator and rotor sections. This also explains the trends observed for the individual PDC operating parameters. A stronger detonation wave, formed by a high fill fraction, high equivalence ratio, or low nitrogen dilution, will survive the passage through the blades better than a weak wave. Specifically a higher pressure detonation wave will be less attenuated through the turbine sections, no matter how that detonation wave strength is achieved. Reducing the initial strength of the wave will lead to a greater acoustic attenuation, and eventually less noise exhausting from the system as a whole.

VIII. Conclusions The effect of PDC operating parameters on the attenuation of the detonation wave peak pressure

through the stator and rotor blades of a turbine was investigated. It has been shown through these variables that the attenuation experienced by a detonation wave due to the obstruction presented by the turbine is heavily dependent on the initial strength of the wave. A stronger detonation wave, characterized by a high peak pressure at the exit of the PDC tube and higher wave velocity, is more capable of weathering the obstruction than a weaker one. For example, high fill fraction cases are known to generate stronger detonation waves than low fill fraction cases, and the results presented here in fact show that peak pressure attenuation is minimized at high fill fractions. Similarly, transmission loss is lowest at stoichiometric conditions and low nitrogen dilution percentages. Increasing the firing frequency seems to have a measurable effect in reducing the transmission loss, though the mechanism is not fully understood. Adding bypass air into the system, and hence spinning the turbine, has the effect of significantly increasing the peak pressure reduction. However, as the turbine speed increases this attenuation effect is reduced. A study was then conducted to separate the effect of the expansion and flow turning that takes place in the turbine directly after the rotor blades. The transmission loss in this portion of the system also seems to be a function of the PDC operating parameters, contributing a widely variable fraction of the peak pressure reduction, between 3% and 23% for a range of fill fractions alone. When the fraction of the total reduction due to the flow turning and expansion is low, it must be assumed that the stator vanes and rotor losses must account for the remaining percentage of the effort in reducing the peak pressure amplitudes. However, in the worst case the turbine blading accounts for 77% of the attenuation. Removing everything but the stator section from the facility, the effect of only the stator vanes was quantified to further separate the effects of the individual components in the attenuation of the pressure amplitudes. It was found that for all fill fractions, the stator vanes account for the bulk of the peak pressure reduction, while the rotor also has a significant impact. The remainder of the pressure attenuation is achieved through the flow turning and expansion behind the rotor, as described earlier. Since the information that has been learned suggests that the initial strength of the detonation wave is the only factor determining the attenuation that is experienced as it passes through the turbine, an attempt was made at collapsing all of the data onto a single graph of pressure attenuation plotted against PDC exit pressure. The data was found to collapse nicely, thereby proving that the initial strength of the detonation wave, whether achieved by variations in fill fraction, equivalence ratio, or nitrogen dilution, is all that is needed to predict how the peak pressure will be reduced through the turbine.

Acknowledgments The authors would like to acknowledge the financial support of both NASA Glenn Research Center in

the development of this facility, as well as General Electric Global Research (GEGR) for continued funding

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of related projects. Thanks are specifically due to Ed Envia at NASA Glenn and Tony Dean and Adam Rasheed at GEGR for guidance pertinent to this work. Thanks are also due to Russell DiMicco, the laboratory supervisor for the Gas Dynamics and Propulsion Laboratory at the University of Cincinnati, for his technical input and assistance.

References 1 Mongia, H., Held, T., Hsiao, G., and Pandalai, R., “Challenges and Progress in Controlling Dynamics in Gas Turbine Combustors”, Journal of Propulsion and Power, Vol. 19, No. 5, 2003, p. 823. 2 Pandalai, R., and Mongia, H., “Combustion Instability Characteristics of Industrial Engine Dry Low Emission Combustion Systems”, AIAA 98-3379. 3 He, X., and Karagozian, A., “Reactive Flow Phenomena in Pulse Detonation Engines”, AIAA 2003-1171, 41st AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, 2003. 4 Glaser, A., Caldwell, N., and Gutmark, E., “Experimental Investigation into the Acoustic Performance of a Pulse Detonation Engine with Ejector”, 43rd AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV, 2005, AIAA 2005-1345. 5 Dittmar, J.H., Elliott, D.M., and Horne, S.E., “The Far Field Noise of a Pulse Detonation Engine”, NASA TM-2003-212307, May 2003. 6 Allgood, D., Glaser, A., Caldwell, N., and Gutmark, E., “Acoustic Measurements of a Pulse Detonation Engine”, AIAA 2004-2879, 10th AIAA/CEAS Aeroacoustics Conference, Manchester, Great Britain, May 2004. 7 Caldwell, N., Glaser, A., Dimicco, R., and Gutmark, E., “Acoustic Measurements of an Integrated Pulse Detonation Engine with Gas Turbine System”, AIAA 2005-0413, 43rd AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, 2005. 8 Hudson, Kate. Private Communication. Dec 2, 2005. 9 Glaser, A., Caldwell, N., and Gutmark, E., "Experimental Performance Measurements of a Pulse Detonation Engine Integrated with an Axial Flow Turbine", 44th AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV, 2006, AIAA 2006-1232. 10 Hassall, J.R., and Zaveri, K., Acoustic Noise Measurements, 5th ed., Bruel-Kjaer, Denmark, June 1988, p. 33. 11 Rasheed, A., Tangirala, V., Vandervort, C., and Dean, A., “Interactions of a Pulsed Detonation Engine with a 2D Blade Cascade”, 42nd AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV, 2004, AIAA 2004-1207. 12 Schauer, F., Stutrud, J., and Bradley, R., “Detonation Initiation Studies and Performance Results for Pulsed Detonation Engine Applications”, 39th AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV, 2001, AIAA 2001-1129. 13 Allgood, D., Gutmark, E., Meyer, T., Hoke, J., Katta, V., and Schauer, F., “Computational and Experimental Studies of Pulse Detonation Engines”, 41st AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV, 2003, AIAA 2003-0889. 14 Schauer, F., Bradley, R., and Hoke, J., “Interaction of a Pulse Detonation Engine with a Turbine”, 41st AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV, 2003, AIAA 2003-0891. 15 Rasheed, A., Furman, A., and Dean, A., “Wave Attenuation and Interactions in a Pulsed Detonation Combustor- Turbine Hybrid System”, 44th AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV, 2006, AIAA 2006-1235. 16 Saravanamuttoo, H.I.H., Rogers, G.F.C., and Cohen, H., Gas Turbine Theory, 5th ed., Prentice Hall, Harlow, England, 2001, p. 248. 17 Rasheed, A., Furman, A., and Dean, A., “Experimental Investigations of an Axial Turbine Driven by a Multi-Tube Pulsed Detonation Combustor System”, 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Tucson, AZ, 2005, AIAA 2005-4209.

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