Propellant Burning Rate Research Articles

Burning rate analysis of laser controlled 5-aminotetrazole propellant

As an innovative propulsion technique, laser augmented chemical propulsion (LACP) seems superior to the traditional ones. However, the corresponding combustion theories have still to be ascertained for LACP. Burning rate of 5-aminotetrazole (5-ATZ) propellant has been studied by testing pressed samples under different combustor pressures and laser powers. Based on micro computed tomography (MicroCT), an advanced thickness-over-time (TOT) method to characterize the regression of the produced non-planar burning surface is established. Because of a shell structure covering the combustion surface, the burning rate of the implemented 5-ATZ propellant is not constant during laser ablation. Resorting to functional fitting, a new law of non-constant burning including the effect of the observed unique burning surface structures is proposed. Accordingly, applicable combustion conditions of 5-ATZ based propellants have been preliminarily speculated for future research activities.

Heat transfer enhancement in stagnation point flow of ferro-copper oxide/water hybrid nanofluid: A special case study

This present work is to investigate the behavior of Fe3O4−CuO/H2O hybrid-type nanofluid flow toward a magnetic field for enhancing the thermal transfer by horizontal stretchable surface. In the base fluid different tiny-sized nanoparticles collides and enhanced the heat transformation in the absence of magnetic field. Here the combinations of Fe3O4−CuO/H2O as hybrid nanofluids and CuO/H2O as nanofluid are implemented. The Ferro and Copper oxide are dispersed in water base fluid. Such Ferro nanomaterials are more fruitful in production technology, protein absorption, catalysis, medical technology and environment. The implementation of Fe3O4 NPs in medical science mainly involves targeted drug/gene delivery, biosensor, magnetic resonance imaging (MRI), contrast improvement and hyperthermia, biophotonics, detection of cancer cells, diagnosis and magnetic field-assisted radiotherapy and tissue engineering. Copper oxide used as a catalyst for increasing the rate of combustion in rocket propellant. It may significantly increase the homogenous propellant burning rate, lower the pressure index, and boost the effectiveness of the AP composite propellant as a catalyst. In additional, the governing coupled dimensional equations are reduced into the ordinary once with associated the suitable similarities. The resultant dimension-less expressions are programmed in the computational MATLAB software via bvp4c solver with shooting scheme. The important outcomes of this investigation are the behavior of numerous good parameters including stretching ratio parameterA∗, magnetic parameterβ, volumetric frictions for nanoparticlesφ, suction/injection parameterfw and thermal radiation parameterRd. The physical behavior of above discussed sundry parameters is described through figures. The dispersion of Fe3O4 and CuO in water-based fluid significantly improved the phenomenon of heat transfer. Results found that the velocity field is escalates when the stretching ratio parameter is enhanced. Furthermore, larger values of the suction/injection parameter causes a reduction in velocity profile for A<1 and A>1. The induced magnetic field is improved with larger estimations of the nanoparticle fraction. Thermal field is escalates by growing thermal radiation parameter and Eckert number. It is further observed that temperature profile is reduces by increasing stretching ratio parameter. From the analysis we noted that skin friction is raises via larger estimations of magnetic parameter.

Characteristics of the Reacting Flow for the Base‐Bleed Projectile under Dual Stress

AbstractWhile the base‐bleed projectile (BBP) is ejected from the muzzle, high‐speed rotation and rapid depressurization in the combustion chamber bring much initial disturbance to the BBP. In order to research the effect of propellant gas on the drag reduction performance for the BBP, a chemical non‐equilibrium flow model is established under the dual environmental stress of high‐speed rotation and rapid depressurization. This numerical model is coupled with a H2‐CO reaction mechanism and propellant burning rate model. Detailed numerical results of flow patterns, heat energy addition, and mass addition for various rotation speeds are discussed. The results show that the supersonic under‐expanded jet gradually transforms into a subsonic jet, and a low‐pressure recirculation zone appears at the wake of BBP. Moreover, the tangential velocity in the combustion chamber presents a typical distribution of the Rankie combined vortex. The high‐pressure jet rapidly expands to the outside low‐pressure environment, and the pressure in the combustion chamber (pc) drops sharply. After 1.15 ms, the outside gas flows back into the combustion chamber, and the rotation increases the fluctuation intensity of the inter flow field. pc reaches 0.718 p0 and remains stable after 4.0 ms. The rotation of BBP increases the pressure and Mach number near the nozzle at 5.0 ms.

Effect of Metal Nanopowders on the Performance of Solid Rocket Propellants: A Review.

The effects of different types of nano-sized metal particles, such as aluminum (nAl), zirconium (nZr), titanium (nTi), and nickel (nNi), on the properties of a variety of solid rocket propellants (composite, fuel-rich, and composite modified double base (CMDB)) were analyzed and compared with those of propellants loaded with micro-sized Al (mAl) powder. Emphasis was placed on the investigation of burning rate, pressure exponent (n), and hazardous properties, which control whether a propellant can be adopted in solid rocket motors. It was found that nano-sized additives can affect the combustion behavior and increase the burning rate of propellants. Compared with the corresponding micro-sized ones, the nano-sized particles promote higher impact sensitivity and friction sensitivity. In this paper, 101 references are enclosed.

Insight into graphene-salen metal nanocomposites on combustion performance and mechanism of HMX-CMDB propellant

A series of novel graphene-salen metal nanocomposites with different active metals and molecular structures were designed, synthesized and characterized using SEM, EDS, FTIR, XPS and RAMAN instruments. Effects of the as-synthesized graphene-salen metal nanocomposites on the combustion performance and mechanism of HMX-CMDB propellant were studied. Among the nanocomposites considered, S-792-Ni has the best effect on the combustion performance of HMX-CMDB propellant, which makes the propellant possess high burning rate, low pressure index and good combustion stability. The excellent combustion performance of HMX-CMDB propellant containing S-792-Ni is closely related to the formation of filamentous clusters, the decrease of dark zone thickness and the increase of gas phase temperature gradient. In addition, the formation of flocculent network carbon skeleton and the decrease of metal aggregations in the quenched sample also contribute to the improvement of combustion performance. Additionally, the interaction mechanisms of efficient catalyst S-792-Ni with energetic components HMX and NC/NG were revealed by the in-situ solid FTIR and gas-phase MS-FTIR methods. The results show that S-792-Ni can obviously promotes the fracture of C–C bonds and C-N bonds of NC/NG and HMX respectively, and significantly increases the decomposition heat of HMX and NC/NG. The increased decomposition heat are beneficial to the rapid reactions beneath the combustion surface and the heat feedback between combustion surface and gas region, so as to improve the burning rate of HMX-CMDB propellant.

Evaluation of graphene-ferrocene nanocomposite as multifunctional combustion catalyst in AP-HTPB propellant

A novel graphene-ferrocene nanocomposite (G-792-Fe) was designed, prepared and characterized systemically using SEM, EDS, FTIR, XPS and RAMAN methods. The as-synthesized G-792-Fe was used as a multifunctional combustion catalyst in ammonium perchlorate-hydroxyl terminated polybutadiene (AP-HTPB) propellant, and the combustion, anti-migration, safety and mechanical performances of AP-HTPB propellant were revealed. The results showed that anchoring on the surface of graphene can effectively inhibit the migration and volatilization of ferrocene compounds in AP-HTPB propellant. The excellent combustion catalytic performance of G-792-Fe on AP-HTPB propellant (burning rate of AP-HTPB propellant containing G-792-Fe increases from 13.87 mm·s−1 to 17.28 mm·s−1 at 15 MPa) is attributed to its positive effect on AP decomposition, which generates more gas-phase products beneficial to the combustion surface thermal feedback and the gaseous REDOX reaction. Additionally, the G-792-Fe presents positive effects on reducing the electrostatic sensitivity and improving the mechanical properties of AP-HTPB propellant. The ignition energy (E50) and the maximum tensile strength (σm) of AP-HTPB propellant containing G-792-Fe were increased by 24.4 mJ and 0.75 MPa respectively, compared with AP-HTPB propellant containing catocene (156.8 mJ and 1.26 MPa). The positive effects of G-792-Fe on safety and mechanical performances of AP-HTPB propellant can be due the outstanding electrical and mechanical properties of graphene-based material. Results of this study have implications concerning design and application of multifunctional combustion catalyst, which can improve the combustion performance, anti-migration performance, safety performance and mechanical properties of AP-HTPB propellant.

A Review on the Use of Burning Rate Suppressants in AP‐Based Composite Propellants

AbstractAmmonium perchlorate (AP) is the most widely used oxidant in composite propellants. Propellants based on AP usually have a high burning rate, but in some applications, it is necessary to have a slower burning rate. The functional additives, which can reduce the burning rate of propellants, were named coolants, burning rate suppressants (BRSs), burning rate depressants, or burning rate inhibitors. This paper introduced different types of BRSs, such as amide‐based compounds, cyclic azines, quaternary ammonium salts, metal salts, etc., their mechanism and development in AP‐based composite propellants. For organic amines, the burning rate trend is as follows: biuret&lt;melamine&lt;oxamide&lt;urea&lt;base. Cyclic azines are special organic amines which form large, thermally stable cyclic azines in addition to common low‐molecular‐weight gases NH3, CO2, HCN, and N2O when heated. With the smallest addition but the highest effect, quaternary ammonium salts will not reduce the specific impulse of the propellant, nor do they interfere with cure reaction or adversely affect the physical properties of the cured propellant. Metal salts reduce the burning rate by forming more stable substances than HClO4 with AP. The mechanism of BRSs is complicated, but all of them can be explained based on AP deflagration. At last, the development trend of efficient BRSs was discussed, the preparation of liquid BRSs and the combination of AP bonding agents with BRSs were both first proposed, which worth further research in the future.

Combustion Characteristics of Nanoaluminium-Based Composite Solid Propellants: An Overview

The nanosized powders have gained attention to produce materials exhibiting novel properties and for developing advanced technologies as well. Nanosized materials exhibit substantially favourable qualities such as improved catalytic activity, augmentation in reactivity, and reduction in melting temperature. Several researchers have pointed out the influence of ultrafine aluminium (∼100 nm) and nanoaluminium (&lt;100 nm) on burning rates of the composite solid propellants comprising AP as the oxidizer. The inclusion of ultrafine aluminium augments the burning rate of the composite propellants by means of aluminium particle’s ignition through the leading edge flames (LEFs) anchoring above the interfaces of coarse AP/binder and the binder/fine AP matrix flames as well. The sandwiches containing 15% of nanoaluminium solid loading in the binder lamina exhibit the burning rate increment of about 20–30%. It was noticed that the burning rate increment with nanoaluminium is around 1.6–2 times with respect to the propellant compositions without aluminium for various pressure ranges and also for different micron-sized aluminium particles in the composition. The addition of nano-Al in the composite propellants washes out the plateaus in burning rate trends that are perceived from non-Al and microaluminized propellants; however, the burning rates of nanoaluminized propellants demonstrate low-pressure exponents at the higher pressure level. The contribution of catalysts towards the burning rate in the nanoaluminized propellants is reduced and is apparent only with nanosized catalysts. The near-surface nanoaluminium ignition and diffusion-limited nano-Al particle combustion contribute heat to the propellant-regressing surface that dominates the burning rate. Quench-collected nanoaluminized propellant residues display notable agglomeration, although a minor percentage of the agglomerates are in the 1–3 µm range; however, these are within 5 µm in size. Percentage of elongation and initial modulus of the propellant are decreased when the coarse AP particles are replaced by aluminium in the propellant composition.

Very-High-Pressure Burning Rates of Aluminized and Nonaluminized AP/HTPB-Composite Propellants

To further characterize the roles of aluminum and ammonium perchlorate (AP) characteristics in ammonium perchlorate/R45-M hydroxyl-terminated polybutadiene (AP/HTPB)–composite propellants, aluminized and nonaluminized formulations were tested at very high pressures, up to 68.9 MPa (10,000 psi). A total of nine formulations were examined: seven nonaluminized with varying AP distributions, concentrations, and average particle sizes, and two aluminized with varying aluminum concentrations. All nine propellants showed an exponent or “slope” break above 20 MPa (2900 psi) and a postbreak pressure exponent greater than one. Decreasing the AP particle size decreased both the characteristic pressure where the exponent break occurred (or ) and the pressure exponent after the break. AP distribution also affected , whereas changes in AP concentration did not. The inclusion of aluminum lowered compared with the nonaluminized formulations. Additionally, increasing the aluminum concentration appeared to lower the postbreak pressure exponent. The present study provides one of the first fundamental, systematic studies on the exponent break feature of AP/HTPB-composite propellants and adds new data to the limited database available in the open literature on the AP-based, composite propellant burning rate exponent break at very high pressures, with emphasis on the effect of mixture variables and Al on the characteristic break pressure and the postbreak pressure exponent.

Nanodiamonds characterization and application as a burning rate modifier for solid propellants

A nanodiamond powder (n-D) synthesized by detonation method was characterized in detail, and the combustion of solid propellants contained 10 wt. % of n-D as a burning rate modifier was investigated. n-D characterization was implemented by XRD, SEM, DSC, and size distribution analyses. Specific surface area analysis was performed using BET-method. The size of primary nanodiamond particles has not changed after acid treatment and drying, but the particles' agglomeration decreased the value of a specific surface area. The effect of n-D modification on the burning rate of solid propellants with an active and inert binder was investigated. The addition of the n-D to propellants formulation instead of carbon black powder has significantly increased their burning rate (by at least 50 %) in the studied pressure range. The observed trend was associated with the high chemical reactivity of n-D and its extremely high value of specific surface area.

Thermal Decomposition of Ammonium Perchlorate in the Presence of Copper Tungsten Oxide (CuWO4)

AbstractThe catalytic effect of copper tungsten oxide (CuWO4) was investigated on the thermal decomposition of ammonium perchlorate (AP) using differential scanning calorimetry. A lowering of the high‐temperature decomposition peak temperature of AP was observed due to catalytic effect of copper tungsten oxide particles, now referred to as CTO in this paper henceforth. The kinetic parameters were evaluated using the Kissinger method. The decrease in the activation energy confirmed the catalytic activity of CTO. Further, CTO was evaluated in a standard composite propellant formulation having 86 % solid loading. Propellant performance and processing parameters such as the end of mix (EOM) viscosity, density, burning rate, pressure exponent (n‐value), etc. were measured for standard as well as a composition containing 1 % CTO. The results revealed that at 1 % level CTO causes more than 36 % increase in propellant burning rate compared to the standard propellant composition at 6.86 MPa pressure. The pressure exponent also increased to 0.48, whereas the standard composition was having its value of 0.41.

Statistical Evaluation of Burning Rate Data of Composite Propellants Obtained from Acoustic Emission Technique

The acoustic emission technique has been considered to be one of the most reliable and robust methods for the measurement of the steady burning rate of composite propellants. In this work, attempts were made to quantify the measurement variability of the burning rate of composite solid propellants by acoustic emission method using statistical tools. A total of 1100 individual measurements were subjected to statistical treatment. The combination of confidence interval and repeatability limit delineated the extent of natural dispersion in the burning rate measurement data. The very high coefficient of variation values for the propellant compositions, having a burning rate of more than 25 mm s–1 raised concerns about the suitability of the acoustic emission method for high burning rate compositions. The Reliability interval approach was employed to determine the statistically significant sample size for different composite propellants having a burning rate range of 5–31 mm s–1. The entire set of data was screened for identification of outlying observation using the Dixon Q test, and the extent of contamination was quantified. Moreover, the application of statistical techniques could have far-reaching implications for quality control perspectives of burning rate measurement by acoustic emission and could be implemented as reference tolerance limits and preventive measures for ensuring the good health of the instrument as well as propellant processing.

Enhanced the thermal conductivity of hydroxyl-terminated polybutadiene (HTPB) composites by graphene-silver hybrid

The heat conductivity affects the burning rate of propellant much, which would influence the properties of power propellant. In this work, hydroxyl-terminated polybutadiene (HTPB) was employed as a model propellant matrix, and series of composites were fabricated by introducing Graphene-Silver (GNs-Ag) hybrid as fillers to improve the heat conductivity of HTPB. GNs-Ag hybrids were prepared via chemical reduction of silver onto the graphene nanosheets, and the effect of preparation conditions on the fabrication of GNs-Ag were studied systematically. With the addition 0.5 wt% of GNs-Ag hybrid, the GNs-Ag/HTPB composites had a 52.6% increase in thermal conductivity compared with that of neat HTPB. The mechanical properties and thermal stability of the GNs-Ag/HTPB nanocomposites were also sightly improved.

Investigation of Electrically Controlled Ammonium Nitrate – Epoxy Solid Propellant at High Pressures

AbstractThe effect of electric voltage/power application on the burning rate of solid propellants composed of ammonium nitrate (AN) and epoxy at high pressures was investigated experimentally. Such an effect can yield an effective means for real‐time managing of the thrust of a solid rocket motor. A burning rate increase up to about 4 folds when increasing voltage application from 70 V to 200 V, was demonstrated. It was attributed to energy (heat) release due to an electric current passing through the conductive melt layer formed at the burning surface. The relative increase in burning rate due to electric effect was higher at the lower pressure range, diminishing for the higher‐pressure end because of the lower relative electric energy contribution. The common pressure dependence correlation of the burning rate could be applied only for limited ranges. However, because of the coupling between the pressure and specific electric power (power per unit surface area) effects, a more complex correlation which includes a variable pressure exponent, was revealed and applied successfully to fit the data over a range of specific electric power and pressure.

Numerical Study on Performance Effect of Burning-rate Ratio in Nozzleless Rocket Motor with Series Dual Burning Rate Grain

ABSTRACT The application of series dual burning rate grain can improve specific impulse of nozzleless solid rocket motor (NSRM). In this paper, burning-rate ratio is the ratio of main propellant burning rate and nozzle propellant burning rate in dual burning rate grain, which has significant influence on NSRM performance. Focusing on relationship between burning-rate ratio and motor performance, internal ballistics simulation of NSRM with different burning-rate ratios has been carried out. Results show that, with the increase of burning-rate ratio, total impulse grows monotonically, while maximum head pressure has two variation stages – the stable stage and the increasing stage. By analyzing the head pressure curves, it is known that maximum head pressure is almost equal to the higher one of initial stage head pressure (p1) and combustion ending stage head pressure (p2). Relationship between p2/p1 and burning-rate ratio has been established by linear fitting, and the critical burning-rate ratio, which is the maximum burning-rate ratio to ensure that maximum head pressure almost is kept invariant, has been obtained. Critical burning-rate ratio has no significant change when geometry size remains invariant. Utilizing critical burning-rate ratio in NSRM with series dual burning rate grain, specific impulse can be enhanced with nearly no sacrifice of maximum head pressure performance. These results can provide references for NSRM designing and optimizing.

ЧИСЛЕННОЕ МОДЕЛИРОВАНИЕ ГОРЕНИЯ СМЕСЕВОГО ТВЕРДОГО ТОПЛИВА, СОДЕРЖАЩЕГО БИДИСПЕРСНЫЙ ПОРОШОК БОРА

A problem of combustion of the composite solid propellants containing various powders of metals and non-metals is relevant in terms of studying the effect of various compositions of powders on the linear rate of propellant combustion. One of the lines of research is to determine the effect of the addition of a boron powder on the burning rate of a composite solid propellant. This work presents the results of numerical simulation of combustion of the composite solid propellant containing bidispersed boron powder. Physical and mathematical formulation of the problem is based on the approaches of the mechanics of two-phase reactive media. To determine the linear burning rate, the Hermance model of combustion of composite solid propellants is used, based on the assumption that the burning rate is determined by mass fluxes of the components outgoing from the propellant surface. The solution is performed numerically using the breakdown of an arbitrary discontinuity algorithm. The dependences of the linear burning rate of the composite solid propellant on the dispersion of the boron particles and gas pressure above the propellant surface are obtained. It is shown that the burning rate of the composite solid propellant with bidispersed boron powder changes in contrast to that of the composite solid propellant with monodispersed powder. This fact proves that the powder dispersion should be taken into account when solving the problems of combustion of the composite solid propellants containing reactive particles.

НЕСТАЦИОНАРНОЕ ПОВЕДЕНИЕ ЗАРЯДА ТТ БЕССОПЛОВОГО РДТТ ПОД ДЕЙСТВИЕМ ГАЗОДИНАМИЧЕСКОЙ НАГРУЗКИ

The numerical solution to a conjugate problem of an unsteady flow of combustion products in a flow path of the nozzleless solid rocket motor (SRM) and the oscillation of a solid propellant charge under the action of the forces directed from combustion products is considered. The Navier-Stokes equations for a compressible viscous gas are used to mathematically describe the flow of the combustion products. To model the charge oscillations, the equations of solid mechanics are applied, which take into account the propellant hyperelasticity. Pressure distributions and the propellant burning rate along the charge channel are presented for different models of the propellant burning rate. It is revealed that at the stage of SRM design, the use of the burning rate law, determined by pressure in the head of the combustion chamber, is more preferable in order to assess the internal ballistic characteristics. The solution to the conjugate problem shows that in the nozzleless SRM with the propellant having low Young's modulus, resonance can occur, which causes uncontrolled charge oscillations.

A New Erosive Burning Model of Solid Propellant Based on Heat Transfer Equilibrium at Propellant Surface

Erosive burning refers to the augmentation of propellant burning rate appears when the velocity of combustion gas flowing parallel to the propellant surface is relatively high. Erosive burning can influence the total burning rate of propellant and performance of solid rocket motors dramatically. There have been many different models to evaluate erosive burning rate for now. Yet, due to the complication processes involving in propellant and solid rocket motor combustion, unknown constants often exist in these models. To use these models, trial-and-error procedure must be implemented to determine the unknown constants firstly. This makes many models difficult to estimate erosive burning before plenty of experiments. In this paper, a new erosive burning rate model is proposed based on the assumption that the erosive burning rate is proportional to the heat flux at the propellant surface. With entrance effect, roughness, and transpiration considered, convective heat transfer coefficient correlation proposed in recent years is used to compute the heat flux. This allows the release of unknown constants, making the model universal and easy to implement. The computational data of the model are compared with different experimental and computational data from different models. Results show that good accuracy (10%) with experiments can be achieved by this model. It is concluded that the present model could be used universally for erosive burning rate evaluation of propellant and performance prediction of solid rocket motor as well.

Acoustic pressure oscillation effects on mean burning rates of plateau propellants

• Influence of acoustic pressure oscillations on mean burning rates of plateau propellants. • Non-aluminized and aluminized propellants exhibit plateau burning trends are examined with corresponding matrixes. • Magnitudes of imposed acoustic amplitudes are varied from 0.1% to 1.4% with respect to mean chamber pressure conditions. • Both non - aluminized propellants and aluminized propellants show significant enhancement in the mean burning rate due to the fluctuations imposed by acoustic pressure amplitudes and frequencies on the propellant combustion. • Maximum burning rate augmentation factors of non-aluminized and aluminized are 1.27 and 1.47 respectively. Combustion instabilities constitute a well pronounced problem in all large rocket motors due to the inherent acoustic pressure oscillations established based on the port geometries. Suppression of the combustion instabilities in the solid rocket motor is required for controlling the mean burning rate variation which arises due to the interaction of acoustic pressure wave with propellant combustion. An experimental study has been carried out to investigate the effects of acoustic pressure oscillations on mean burning rates of non-aluminized and aluminized propellants which exhibit low pressure exponent index (n) in the burning rate trends. Steady and unsteady mean burning rates are determined from combustion photography method using a window bomb test facility over the pressure range of 17 MPa. A rotary valve is coupled with the window bomb setup to generate acoustic pressure oscillations inside the test chamber (cylindrical pressure vessel), which imposes the required frequencies of 140, 180 and 240 Hz respectively. The acoustic pressure amplitudes are varied from 0.04% to 1.4% of the mean chamber pressure. Both non-aluminized propellants and aluminized propellants have shown significant enhancement in the mean burning rate due to the fluctuations imposed by acoustic pressure amplitudes and frequencies on the propellant combustion. The enhancement in the mean burning rate also depends upon dynamic response of the flame to the excited frequencies and acoustic pressure amplitudes. The plateau burning behaviour of the non-aluminized propellant is completely distorted whereas it is retained in aluminized compositions. Conversely, it also shifts the mean pressure range of plateau burning rate trend. The maximum burning rate augmentation factors resulted from imposed acoustic pressure wave on non-aluminized and aluminized propellants are observed as 1.27 and 1.47 respectively.

Linear burning rate of double base propellant containing Azidodeoxycellulose

ABSTRACT Azidodeoxycellulose (AC) is evaluated in a closed vessel as part of a propellant formulation containing nitrocellulose (NC) and nitroglycerin (NG) as binders. It is found that the AC systematically generates greater pressure and a faster linear burning rate when compared to a formulation having a similar concentration of a mixture of cellulose and NC that is meant to replicate a comparable substitution degree. An additional finding is that a formulation containing 13%wt of AC generates higher pressure in a closed vessel than the mixture of NC and NG alone. These results are evidence that AC is an adequate solid filler for propellant formulations.