Combustion Instability Additive Investigation Dr. Fred S. Blomshield and Naval Air Warfare Center China Lake, CA

Prof. Merrill W. Beckstead Brigham Young Upiversity Provo, Utah

Richard A. Stalnaker 2ooo/ Large Amplitude , Motor Failure

t Oscillations Motor without Stability Additive

._ 1500 :: il 4 1000 is

f

ii Bulses I ov ’ ’ ’ ’ ’ lL=-+“J ’ ’ ’ ’ 0 1 2 3 4 5 6 Time -seconds

35th AIAA/ASME/SA~/ASEE Jdint Propulsion Conference and Exhibit

June 20-24,1999 Los Angeles, CA

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics, 1801 Alexander Bell Drive, Suite 500, reston, VA 22091-4344. ,

.

Coinbustion Instability Additive Investigation*

F.S. Blomshield’, R.A. Stalnaker and M.W. Be&stead+ Naval Air Warfare Center

Code 4T4320D, Propulsion Research Branch China Lake, CA 93555-6100

ABSTRACT

Combustion stability additives like zirconium carbide (ZrC), aluminum oxide (A&OJ and zirconium ortho- silicate (ZrSiOJ have long been known to suppress combustion instability in reduced smoke, composite propellant solid rocket systems. Often, as little as 0.5 percent additive can stabilize an otherwise unstable rocket motor. The additives appear to have effects on both linear and non-linear pulsed instabilities. Although several theories have been proposed, the actual mechanism on how stability additives work remains unknown. The common belief that additive particle damping alone stabilizes rocket motors is

. not true. Somehow, the additives change the response behavior of the propellant. Past studies have shown that the additive effect is a combination of particle damping and a reduction in the combustion response of the propellant. In the past study, four pro- pellants were studied containing 0, 1,3, and 5 percent ZrC. In this study, the 3 percent propellant used hefore will be used again, except 3 percent HMX will be used in one formulation and 3 percent ultra fine aluminum or ALEX will be used in another. The emphasis here is to examine the combustion response changes. This paper will present the results of T-Burner response testing and compare the results to past propellants containing additives. The reason for the work is that recent evidence suggests that traditional additives may not work as well when solid motors are operating at higher pressures. In addition, additives like the ones proposed, add energy to the propellant which would be a performance advantage over classical additives like ZrC.

INTRODUCTION AND BACKGROUND

Stability additives are crucial ingredients in nearly every modem reduced smoke, composite propellant foxmu- lation in use today. They have been proven very effective in reducing and/or eliminating combustion instability in many solid rocket systems. In one recent study numerous reduced smoke composite propellant tactical sized rocket motors were intentionally pulsed. ‘*z3 Many of the motors containing pro- pellants without additives were pulsed into violent unstable non-linear instability, accompanied by extensive DC pressure shifts, and often resulted in nozzle or motor case failure. These same motors, cast with propellants containing 1 per- cent ZrC, Sum f&O3 or 90pm Alz03 in place of l-percent the ammonium perchlorate (AP), could not be pulsed into un- stable behavior even when pulsed twice as hard. Fig. 1 shows this graphically.

The effect of particles on acoustics has been studied for many years, and specific application to solid rockets has evolved over the past forty years. The theoretical models describing particle damping are primarily based on the early work by Epstein and Carhart, and later refined by Ten&in and Dobbins.5*6 These modeling efforts have been validated by comparison with data gathered on aerosols by Dobbins and Ten&in’** and Zink and Delasso.g The normal expression for particulate damping due to viscous effects as expresses by Culick” is:

Here, C,,, is the weight ratio of particle to gas; o is the angular frequency, ro is the particle relaxation time:

(2)

Motor without

identical Motor

0 0 1’ 2 3 4 5 6

Time - seconds

Fig. 1. Pulsed Motor Firing with and without a Stability Additive

* This effOrt was sponsored by Independent Research Funding of the Naval Air Warfare Center. ’ Research Scientist, Research and Technology Division, Senior Member AJAA + Consultant to the,Naval Air Warfare Center, Chemical Engineering Department, Brigham Young University, Provo, Utah, Senior Member AIAA

Approved for Public Release; Distribution is Unlimited.

Here, pP is the particle density; p is the gas viscosity; and D is the particle diameter. There is also a corresponding term due to thermal attenuation, which is normally an order of magnitude smaller than the viscous attenuation term. This equation gives an optimum particle damping at @rd = 1, which defines an optimum particle diameter of:

For a frequency of 300 Hz, typical of a tactical missile, an optimum particle size is - 7 to 20 pm depending on the particle density.

Fig. 2 shows the relative damping efficiency versus particle size for 10 different frequencies. Equations 1-3 were used to create this plot. Notice that the damping is plotted on a log scale and that the damping variation with particle size and frequency is several orders of magnitude. This plot was generated using typical values for aluminum oxide in an inert binder, composite propellant formulation. The total mass of aluminum oxide in the propellant products was held constant. Fig. 3 is a plot of frequency versus optimum particle size to damp at that frequency. Fig. 3 is the peaks from Fig. 2 plotted for many different frequencies and particle sizes.

Qualitative verification of Equation (1) was made by Horton and McGie” using a T-burner. Varying aluminum concentration in a series of propellants from 0,0.5, 1, 1.5 and

- 5000 Hz -- 2000 Hz - 1000 Hz - 500Hz - 300 Hz --- 200 Hz - IOOHz

Particle Size - microns Fig. 2. Relative Particle Damping Efftciency

versus Particle Size

Optimum Particle Diameter - microns Fig. 3. Optimum Particle Size versus Frequency

2 percent, they observed that the tests went liom spontan- eously unstable (a - 130 sed’) for the zero percent aluminum propellant, to stable for those with greater than 1.5 percent aluminum. The size of the AlaO, generated during the com- bustion was not known, thus giving qualitative verification of the particle damping effect. It is significant that the data did not show a significant change in the measured response function of the propellants due to varying the aluminum concentration.

A quantitative verification of the Ten&in and Dobbins equation was later established by workers at China Lake in a series of papers’213*14 comparing data from a T-burner with calculations based on Equation (1). They incorporated 5 and 10 percent Al203 into propellants that they tested in a T- burner. After firing they collected and carefully sized the residue, and then used the measured particle size distribution to calculate the particle damping. The calculated attenuation values compared well with the measured T-burner values.

Most reduced smoke propellants that are fielded in solid rockets contain a small percentage of additives, incor- porated into the propellant formulation for suppressing .,,: _ potential combustion instability. Typically, these additives are 0.5 percent graphite flake and 1 percent ZrC. These two ingredients apparently became somewhat standardized as a result of a study performed by Aerojet and reported in two papers15~‘6 and a fmal report” in 1973-74. As part of the study, T-burner and 2.75 inch motor tests were performed to evaluate stability additives in general. In the 1973 paper, they concluded that any form of graphite provided improved damping in both T-burner and motor. However, they also concluded that their T-burner damping data were incon- clusive, and did not show any particle damping trends. Unfortunately, they used cup samples in their T-burner, which result in a small fillet of burning propellant after the main sample has burned out at precisely the time, when damping data are obtained. Thus, the use of cup samples in a T-burner totally invalidates damping data from the testing.

In the 1974 paper and report they” continued T-burner and motor tests with a variety of additives, but focused on ZrC and Al203 (in conjunction with carbon). The motor tests showed all of the additives reduced the potential for insta- bility, but the ZrC resulted in better light transmission. Thus, the combination of ZrC and carbon as stability additives was apparently established. However, part of the basis for their conclusions was their T-burner damping data, which did not show any particle damping trends. Again, this is not a just- ifiable‘conclusiqn due to their use of cup samples in their T- burner. The data did apparently indicate a slightly reduced response function for the additive propellants, thus indicating that the stabilization mechanism was apparently due to some kind of negative catalytic effect of the ZrC at the propellant burning surface. It would appear that these effects and con- clusions need to be re-evaluated with greater precision, to determine optimum additives for instability suppression and to understand the mechanism whereby they function.

In a parallel study, Grump’* also performed T-burner testing on reduced smoke propellants, evaluating the effects of additives. He found that one-percent additives (various

Page 2 of 9

forms of carbon, ZrC and Al2033 generally, (but not always) caused a decrease in driving, and the expected increase in damping. Contrary to the Aerojet work, he observed that adding either ZrC or A&O3 to a propellant resulted in a significant increase in damping, but the effect of particle diameter was not completely consistent with the theory. His results also showed that damping increased with increasing concentration, consistent with theory.

Kubotalg has examined the effectiveness of both aluminum and Al203 as stability additives in a small motor. He observed that aluminum was apparently more effective than Al203 as an instability suppressant, and that finer additives were more effective than larger particles.

In recent papers, Blomshield, et a1.‘*2V3 also evaluated the effect of seven additives in reduced smoke propellants, replacing 1 percent AI’ with 1 percent: 3 pm ZrC, 8 and 90 ,um Als03, graphite flake, 4 pm and 0.7 pm ZrSiO, and plain HTPB binder. T-burner tests of all of these additive propel- lants had response values lower than a baseline propellant for frequencies up to > 1000 Hz, but the damping data were inconsistent. The ZrC propellant and the two A1203 propel- lants were fired in motors and pulsed. All 3 motors with I- percent additive were stable to ahpulses while the baseline propellant without an additive was dramatically unstable to the third pulse, see Fig. 1. Again, these tests provide dramatic evidence that stability additives work in motors. However, it is not clear if the additives work due to a reduc- tion in the propellant response, increased particle damping or a combination of both. Unfortunately, the data from the motors were not sufficiently quantitative to discern which additive was more effective.

The effect of stability additives in motors has been quite consistent and effective. Additives do increase the

Page 3 of 9

margin of stability in a motor, and can totally stabilize a motor with oscillatory characteristics. However, it is not clear what the mechanism is that causes the stabilizing influence, nor is it clear which additive or what size of additive is optimum or most effective. Also, based on

Horton’s and Kubota’s data, and simple logic, using aluminum as an additive appears particularly attractive. Substituting 1 percent ZrC for AP reduces I,, by -0.8 set-‘, while substituting 1 percent Al increases I,, by -1.1 se&, thus there is a net difference of almost 2 sec.’ between the two additives. For this difference, it would seem worthwhile to evaluate using aluminum as an additive. Some will argue that the aluminum causes smoke. That is true, but so does ZrC. The virtues of one additive versus the other will depend on the damping generated, which in turn depends on the size of the smoke generated. This can only be evaluated by experimental data.

It seems appropriate that the effectiveness of ZrC as a stability additive should be re-evaluated using more quantitative means to compare the results with other potential additives such as aluminum or aluminum oxide, In a past study, four propellants were cast with differing degrees of z.(-N21 Four the present study, two more propellants were cast, one containing three percent HMX and one containing three percent Ultra Fine Aluminum (UFAL). Because of the relevance to the current work, the results of the past work examining ZrC will be presented as well. Table 1 shows the formulations and specifics of all the propellants discussed in this paper, both past and present. As can be seen in Table 1, all propellant had roughly the same burning rate, exponent and propellant combustion properties. The base line pro- pellant, IRO, contained no stability additive. For all the add- itive propellants, the additive was substituted for ammonium perchlorate. The relative size ratio of the t&modal blend

of AP was kept constant. One ,mconsistency in the formu- lations has to do with the presence of carbon black. In the baseline and in the propellants containing ZrC, 0.5 percent carbon black was in the formulation. Due to an error during the mixing of the HMX and UFAL propellants, the carbon black was left out. An additional 0.5 percent of the n-i-modal ammonium perchlorate was in its place. Although this obviously adds some uncertainty in the data that follows, there is little evidence that the carbon black affects the combustion response. Its original purpose was to prevent “worm-holing” in double base propellant rocket motors. With the short burn times and thin web of the composite propellant testing described in this paper, worm holing was not an important issue to be concerned with.

T-BURNER RESPONSE MEASUREMENTS

The pressure-coupled response is the amplification or attenuation of acoustic pressure waves by the combustion zone of a burning solid propellant. The higher the response at a particular frequency the more a propellant will drive and couple with acoustic oscillations in a motor. It is desirable to know the response so that predictions can be made on the stability of potential solid propulsion systems. In this study, the response was measured for each of the six propellants at 1OOOpsibetween300and2000Hz.

The standard way to measure the combustion response is by the T-burner.” The T-burner shown in Figure 4 requires two disks of propellant of equal thickness mounted as shown in Figure 5. One sample is placed on each end of a l.S<mch diameter pipe combustor and they are ignited simultaneously. Ideally, they will also burn out simultaneously. Data is obtained by pulsing the burner during the bum and after bum out. For some propellants, pulsing is not required and the T-burner acoustic oscillations grow spontaneously. In either case, the difference between the alpha during sample burn, aI, and the decay alpha after burnout, az, is known as the combustion alpha, &. The pressure amplitude rate of change of each alpha is measured by a piezoelectric quartz pressure transducer.

/-- PROPELLANT SAMPLES

-\ -/ . ‘\

/‘ HIGH PREQUENCY (-PRESSURE TRANSDUCER

\

I I

The combustion alpha is directly related to the pressure coupled response. The combustion driving alpha, Al, and the damping alpha, CQ, can be determined from:

and ad =a2(h) (435)

where: al pressure decay rate constant during bum a2@b pressure decay constant after burning

with correction to frequency of aI Sds, propellant burning surface area to

channel area ratio

From the computed G burner length and propellant properties, the combustion response can be computed with:

(6)

Where: jj mean pressure

G measured burning rate

f frequency a theoretical speed of sound of the gases

PP propellant density am measured speed of sound a,,, = 2fL, L = burner length

The test frequency depends on the burner length and the combustion gas temperature. Currently at NAWCWPNS, burners of lengths 50 to 4 inches, corresponding to a fieq- uency range of 300 to 4000 Hz are used. The NAWCWPNS High Pressure T-burner is capable of evaluating the response at pressures up to 4200 psi. The measurement of a2 requires the triggering of the second pulse at burnout of the two propellant samples. In order to determine burnout for proper timing of the second ulse, a phototransistor is mounted behind each sample. 2f See Fig. 5. The detector sees a burst of light as the sample bums out. It is desired that outputs of each phototransistor mounted on each sample occur at the

SCALE k-1 IN@H

\ ,-PROPELLANT SASLE

I+ \ 7 w’-

EXHAUST

Fig 4. T-Burner

fit$&--/ I+ WY

L’“l’k,k, INa I EPOXY FILLER

Fig. 5. Propellant Sample and holder

Page 4 of 9

same time. Ideally, the two samples should burn out simultaneously. However, if the samples do not bum out simultaneously then the analysis of the second pulse must be delayed until after the slowest sample has burned out.

Figs. 6 and 7 shows sample data traces for the a test of the HMX additive propellant and the UFAL propellant, respectively. From top to bottom, the traces shown are the AC pressure, DC pressure and the two phototransistor out- puts. The pulse and oscillations to determine a, can be seen in the top AC pressure trace around 0.3 seconds. The decay pulse and oscillation to determine a~ can be seen around 0.7 seconds. The exact pulse times are different for each propel- lant due to the burning rate variations between the HMX and UPAL propellants, see Table 1. Phototransistor outputs often have different characteristics as shown. For the HMX pro- pellant, which was the color of peanut butter and was slightly translucent, the phototransistor actually sees ignition through the propellant. The top burn out trace shows a spike where the second pulse occurred, while the bottom burn out trace barely shows the first pulse. The UFAL propellant was more opaque and the phototransistor shows this. Only the second pulse is observed by the bum out detectors. In both cases, sample burnout is evident and that is where the second pulse

ot...t f 0.0' 0.2 0.4 0.6 0.8 1.0

Time - seconds 400 800 1200 1600. 2000

Frequency-Hz

Fig. 6. Sample Data for 3HMX Propellant Fig. 8. Decay Alphas During Combustion for Propellants Containing ZrC Additive

0.4 0.6 0.8 1.0 Time - seconds 400 800 1200 1600 2000

Frequency-Hz

Fig. 7. Sample Data for 3UF4AL Propellant

Page 5 of 9

Fig. 9. Decay Alphas During Combustion for Propellants Containing 3% Additive

was fried. Although somewhat subiective, the point where the burnout trace rolls over is where burnout occurs. In both these test cases, burnout occurred simultaneously for both samples. If the traces show uneven burnout, that test is suspect and usually not used. For both tests shown in Figs. 6 and 7, the burner length was 15 inches, which corresponds to a nominal frequency of 1200 HL The test pressure was 1000 psi. For the HMX example in Fig. 6 the alpha 1, alpha 2 and combustion alphas were computed to be -9.1, -26.8 and 17.4 se&, respectively, and the computed response function for this test was 1.63. For the UFAL. example in Fig. 7 the alpha 1, alpha 2 and combustion alphas were computed to be -23.7, -5 1.5 and 29.1 sed’, respectively, and the computed response function for this test was 2.30.

RESULTS

Fig. 8 shows the decay alpha measured during sample bum, at, for the propellants looking at the effect of ZrC stability additive. In this figure are four curves representing the propellants containing 0, 1,3 and 5 percent ZrC in their formulation (IRO, IRl, IR3 and IIU). Fig. 9 shows the measured decay alpha for propellants containing three- percent stability additive. Two of these propellants are the

same as in Fig. 8. They are the baseline propellant containing no additive (IRO) and the propellant containing three percent ZrC (W). The other two curves in Fig. 9 show the decay during bum alpha for the propellant containing three percent HMX and the one containing three percent ultra-fine alu- minum. Figs. 10 and 11 are similar to Fig. 8 and 9 except the measured decay alpha after sample burnout is presented.

Frequency - Hz

Fig. 10. Decay Alphas After Combustion for Propellants Containing ZrC Additive

-+- 3% HMX ---#-- 3% UF Aluminum

-80 I 0 400 800 1200 1600 2000

Frequency - Hz

Fig. 11. Decay Alphas After Combustion for Propellants Containing 3% Additive

Examining the measured alphas during combustion showed that the magnitude of the decay increased both as the concentration of additive increased and as frequency in- creased. The same can be said for decay alphas after burnout. The exception to this behavior is the three-percent HMX propellant, 3HMX, which showed very similar decays to the zero percent baseline propellant, IRO. This is expected since the HMX propellant, like the baseline, have no condensed phase products of combustion and hence no contribution to particle damping. The difference in decay after combustion, ad, between the additive propellants and IRO containing no additive is a direct measurement of the damping produced by the combustion products for the additive propellants. As expected, damping increased as the concentration of ZrC increased. In addition, the efficiency of the particle damping increased as frequency increased.’ This is expected referring

back to Fig. 2 showing the relative damping versus particle size for 10 different frequencies.

Figs. 12 and 13 shows the computed combustion alpha, a, for both propellant families. Recalling Equations (4) and (5), this is the difference between the alphas pre- sented in Figs. 6 and 8 for the ZrC propellants and in Figs. 9 and 11 for the three-percent propellants, respectively. This difference is adjusted for frequency shifts by extrapolating the alpha after burnout curve to the same frequency of the alpha during bum curves. The combustion alpha, o&, is used to determine the pressure-coupled response, Rr, of the propellants by using Equation (6). In Figs. 12 and 13, dif- ferences among the propellants become more pronounced as the frequency increases. It is also interesting to note that the three-percent ZrC propellant, IR3, and the three-percent ultra- fine aluminum propellant, 3UFAL, have very different combustion alphas. This can be seen in Fig. 13. For all the curves presented above, a first, second, third or fourth order polynomial curve tit was used to fit the data.

The final pressure coupled response, determined by Equation (6), is shown in Figs. 14 and 15. For Fig. 14, showing the ZrC propellants, the propellant with the highest

400 800 1200 1600 2000 Frequency - Hz

Fig. 12. Combustion Alphas for Propellants Containing ZrC Additive

Frequency - Hz

Fig. 13. Combustion Alphas for Propellants Containing 3% Additive

Page 6 of 9

propellant, 3HMX, showed only a slight reduction in the response function over the baseline propellant containing no additive.

0.0; 1 ( o I ’ 400 800 1200 1600 2000

Frequency - Hz

Fig. 14. Pressure Coupled Response Function versus Frequency for ZrC Propellants

3*o*

Frequency - Hz

Fig. 15. Pressure Coupled Response Function versus Frequency for 3% Additive Propellants

response was IRO containing no stability additive. The IRl and IR3 propellants were next. It was unusual to find that the. propellant with the highest concentration of ZrC, IR5, actually produced a higher response than some of the others, especially at higher frequencies. One possible reason for this could be the presence of distributed combustion response produced by the burning ZrC particles. This would have increased the measured response and reduced the particle damping, both of which happened comparing the IR5 propellant to the others. The T-burner cannot distinguish between distributed combustion and pressure coupled combustion response.

For Fig. 15 showing all the three-percent additive propellants, the ultra-fine aluminum propellant, 3UFAL, had the highest response at all frequencies except the highest around 1800 Hz. It is postulated that the intense aluminum combustion at the surface or distributed combustion from the line aluminum particles enhances the response. As frequency is increased, this phenomenon decreases either due to a change in the combustion response at higher frequencies or simply due to more efficient particle damping at the higher frequencies. The propellant with the lowest response was the three-percent ZrC propellant, IR3. The three percent HMX

The nondimensional frequency, omega-R, is oflen used by theoreticians to understand the nature of the insta- bility. Many composite propellants whose pressure coupled response has been evaluated by the T-burner have a response peak with an omega value of between 5 and 30.” Omega is the ratio of the acoustic time to the thermal conduction time and it is computed by the following:

Where: f Frequency in cycles per second 71 Thermal diffusivity, common value to

use is 3 x 10”’ i&sec F Burning rate in units of inkec

In this expression q / F2 is the characteristic time of the thermal wave. Figs. 16 and 17 plot the pressure-coupled response versus Omega. As can be seen on the plot the response peak varies between 7 and 10, well within the norm for response function data. The exception to this is the three-

3 I

-.- 0 5 10 15 20 25 30 Non-Dimensional Frequency - Omega

Fig. 16. Pressure Coupled Response Function versus Omega for ZrC Propellants

I-O- Baseline rl

I O-O0

I 5 10 15 20 25 30 Non-Dimensional Frequency - Omega

Fig. 17. Pressure Coupled Response Function versus Omega for 3% Additive Propellants

Page 7 of 9

percent HMX propellant whose peak is around 20. This propellant has a burning rate about 20 percent lower than the others. This lower rate in the denominator increases Omega at a given f?equency explaining some of this trend. The total lack of any particle damping for the HMX propellant might also explain some of this behavior. As mentioned previously, the damping is more effective at higher frequencies so differ- ences due to damping may be more apparent as frequency is increased when comparing the propellants in this study.

CONCLUSIONS

In this study, the pressure coupled response of six propellants differing in amount and type of stability additive was investigated. The response of all propellants was determined by the T-burner at 1000 psi between 300 and 2000 Hz. In general, the higher the concentration of ZrC the lower the response and the greater the particle damping. This is consistent with past studies performed at China Lake.‘2”4*‘s The propellant containing 5 percent ZrC did not follow this trend at higher frequencies and this may be attributed to distributed combustion, although this conclusion is very tentative without further evidence. The ultra-fine aluminized propellant had the highest response, although it did come down at higher frequencies. This result contrasts past studies by Horton and McGie where a change in response was not observed when conventional aluminum was added to pro- pellants in small quantities.” The result that ultra fine

aluminum increased the response also differs from Kubota’s work where finer aluminum was more effective in reducing combustion instability than coarse altinum.‘g However, Kubota’s work only dealt with conventional aluminum particle sizes. The HMK propellant had a slightly lower response than the baseline zero additive propellant at lower frequencies. As frequency (or Omega) was increased, it showed higher responses. Uncertainty in the results concern the presence of, or lack of, carbon black in the formulation. Although it is well established and believed by the authors of this paper that carbon black has minimal effect on combus- tion instability, it is not certain. Future studies need to look at the effect of carbon black more carefully.

Work in the future may apply other diagnostic techniques to examine differences and understand the instability mechanisms that cause these propellants to behave so strikingly different in actual rocket motors. These techniques may include cinema photography, spectrographic, light emission and differential scanning calorimeter measure- ments and planer laser induced fluorescence measurements. In addition, propellants should be formulated with 3 percent of additional binder in place of AP. It also might be desirable to keep the burning rate constant by holding the tine AP/HTPB ratio constant. In this manner, rate will not cloud the results or effect the response by changing the oxidizer to fuel ratio at the surface of the propellant.

,

REFERENCES

1. F.S. Blomshield, J.E. Crump, H.B. Mathes and Merrill W. Be&stead, “Stability Testing and Pulsing of Full Scale Tactical Motors,” NAWCWPNS TP 8060, Parts 1 and 2, Naval Alr Warfare Center Weapons Division, China Lake, CA, February 1996.

2. Blomshield, F. S., Beiter, C.A., Mathes, H.B., Crump, J.E. and Beckstead, M.W., ‘Nonlinear Stability Testing and Pulsing of Full Scale Tactical Motors,” AL4A Propulsion Meeting, A&4-91-1953,1991 and AL4A Journal of Propulsion and Power, 1997.

3. Blomshield, F. S., Grump, J.E., Mathes, H.B. and Be&stead, M.W., “Stability Testing of Full Scale Tactical Motors,” AL4A Propulsion Meeting, W-91- 19‘54,199 1 and AL4A Journal of Propulsion and Power, 1997.

4. Epstein, P. A. and Carhart, RR, “The Absorption of Sound in Suspensions and Emulsion: Water Fog in Air,” Journal of Acoustical Society of America, Vol. 25, No 3, pp. 553-565, 1953.

5. Ten&in, S. and Dobbhrs, RA.,” Attenuation and Dispersion of Sound by Particulate-Relaxation Processes, “Journal of Acoustical Society ofAmerica, Vol. 40, No 2, pp. 317-324, 1966.

6. Dobbins, R A. and Ten&in, S., “Propagation of Sound in a Gas-Particle Mixture and Acoustic Combustion

7.

8.

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Ten&in, S. and Dobbins, R. A., “Measurements of Attenuation and Dispersion of Sound by an Aerosol,” Journal of Acoustical Society of America, Vol. 40, NO. 5, pp. 1016-1024, 1966.

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Horton, M. D. and McGie, M. R, “Particulate Damping of Oscillatory Combustion,” AUA Journal, Vol. 1, No. 6;pp. 1319-1326, 1963.

Derr, R L., Kraeutle, K.J., Mathes, H.B. and Dehority, G.L., “Combustion Instability Studies Using Metalized Solid Propellants: Part I, Experimental Verifications of Particle Damping Theory,” 12th JANNAF Combustion Meeting, Vol. II, CPIANo 273, pp. 155-166, 1975.

Page 8 of 9

13. Mathes, H. B., Kraeutle, K.J., Dehority, G.L., and Derr, R.L., “Combustion Instability Studies Using Metalized Solid Propellants: Part III, Characteristics of Particulate Metal Oxide Residues in T-Burners,” 12th JANNAF Combustion Meeting, Vol. 11, CPIA No 273,. pp. 167- 180,1975.

14. Kraeutle, K. J., Derr, R.L., Mathes, H.B., and Dehority, G.L., “Combustion Instability Studies Using Metalized Solid Propellants: Additional Experimental Evidence for the Validity of Particle Damping Theory,” 13th JANNAF Combustion Meeting, Vol. II, CPIANo 281, pp. 155-166, 1976.

15. Micheli, P. L., Zimme&nan, GA. and Lovine, RL., “Smokeless Composite Propellant Combustion Stability Study (U), “I 0th JANNAF Combustion Meeting, Vol. IV, CPIA No 243, pp. l-26, 1973.

16. Micheli, P. L., “Stabilization &Smokeless Propellants with Additives,” 11 th JANNAF Combustion Meeting, Vol. III, CPIANo 261, pp. 123-136, 1974.

17. Lovine, R. L., Micheli, P.L. and Zhnmerman, G.A., “Smokeless Composite Propellants,” AFRPL-TR-74-26, Aerojet Propulsion Company, 1974.

18. Grump, J. E., “Combustion Instability Studies on Non Metalized AP-HTPB Propellants, “9th JANNAF Combustion Meeting, Vol. III, CPIA No 23 1, pp. 123- 134,1972.

19. Kubota, N. and Yano, Y., “Particula& Damping of Acoustic Instability in RDX/AP Composite Propellant Combustion, “18th Joint Propulsion Conference, AIAA- 82-1223, 1982.

20. F.S. Blomshield and Richard A. Stalnaker, “Solid Propellant Combustion Stability Additive Investigation,” Proceedings of the 1996 JANNAF 33’ Combustion Meeting, Naval Postgraduate School, Monterey, California, 4-8 November 1996.

21. F.S. Blomshield, Merrill W. Beckstead and Richard A. Stalnaker, “Solid PropelIant Combustion StabiIity Additive Mechanisms, “Proceedings of the 1997 28” International ICT Conference, Karlsruhe, Germany, June 1997.

22. “T-burner Manual,” CPIA Publication No. I91, November 1969. I

23. F.S. Blomshield, J. C. Finlmson and R Stalnaker, “Photoelectric Detection of T-Burner Grain Burnout,” 26th JANNAF Combustion Meeting, CPIA Publication No. 529, Vol. II, pp. 153-160, October 1989.

24. F.S. Blomshield, “Lessons Learned in Solid Rocket Combustion Instability,” 29th JANNAF Combustion Meeting, CPIA Publication No. 593, Vol. IV, pp. 177- 196, NASA Langley Research Center, Hampton, Virginia, 1992.

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