Strand Burner Research Articles

Investigations into the copper chromite particle size effect on the combustion characteristics of poly(vinyl-chloride) plastisol propellants

The investigation of the possibility to optimize ballistic behavior and consequently to obtain better performance without affecting the cost and safety aspect of plastisol propellants has been presented in this paper. Addition of specific additives and changes in particle size can be one of these solutions, which are continuously in research. The use of burn rate modifier and metallic fuel, with finer particle sizes, becomes mandatory, to enhance the catalytic effect on ammonium perchlorate and to offer additional high heat release. Moreover, the knowledge of the combustion characteristics and especially the burning rate depending on the combustion chamber pressure of the propellant are important conditions in the validation and successful design of a solid rocket motor. The first part of this work deals with the selection of the higher energetic catalyst among three of the most common of them, namely copper chromite, ferric oxide and ferrocene. The selection results in choosing copper chromite as the best burning catalyst. Then, we have studied the effect of its particle size on the thermoanalytical properties of poly(vinyl-chloride) plasticized propellants, by the determination of energetic and kinetic characteristics through the use of an adiabatic bomb calorimeter and a differential scanning calorimeter respectively. The kinetic parameters were determined by Ozawa and Kissinger methods and compared. Besides, the effect of particle size on the combustion properties has been also studied. The plot of the burning rate-pressure curves is used to determine the combustion laws using a strand burner in the typical Crawford bomb. Copper chromite and its particle size have been found to influence the decomposition and to enhance the burning rate of the plastisol propellants. In the second part, we have investigated the energetic effect of fine aluminum powder on the thermal decomposition of ammonium perchlorate and of plastisol propellant as a function of its concentration, both in the presence and the absence of 1 mass-% copper chromite in the basic propellant composition. Aluminum is chosen because of its effect to get high flame temperatures and increased performance.

Aging Effects on the Burning Rates of Composite Solid Propellants with Nano-Additives

The aging behavior of composite propellant is a key parameter in determining the usable service life of an energetic system. The effect of aging on composite propellant burning rates is an understudied topic in the open literature. Recent research in the authors’ laboratory has led to the development of high-burning rate propellants containing titania nanoparticles and fluorescing propellants containing quantum dots. The current study assessed the aging behavior of composite ammonium perchlorate (AP)/hydroxyl terminated polybutadiene propellants, with specific emphasis on fluorescent properties and burning rates. Propellant formulations were based on either 80% monomodal or 85% bimodal AP distributions with and without titania nanoparticles and quantum dots. Propellant samples were thermally aged at temperatures of 63 and 85°C and humidity levels of 25 and 75% for up to 195 days to yield approximate service lives in the range of 0–25 years. Fluorescent imaging experiments were conducted with an ultraviolet light source to measure the uniformity of propellant samples. Ballistic testing was conducted in a constant-volume strand burner over the pressure range of 500–2250 psia (3.5–15.5 MPa). Spatial uniformity of the propellant fluorescence was not significantly altered with propellant age but did decrease in intensity to some degree. In general, the burning rate and pressure index of propellants decreased with increasing service life. Propellant formulations containing 85% bimodal AP were found to be more stable than formulations containing 80% monomodal AP. Inclusion of quantum dots or titania nanoparticles in the propellant formulation increased aging stability.

Burning Rates of Nanoaluminum–Water Solid Propellants at Various Pressures

The burning rates of solid propellants comprising different nanoaluminum powders and distilled water were examined at pressurized conditions up to . Several aluminum powders were tested using a strand burner technique: uncoated nanopowder (ALEX), stearic-acid coated nanopowder (L-ALEX), Viton-coated nanopowder (V-ALEX) with two different particle sizes, and bimodal mixture of V-ALEX and micrometer-size powders. Propellants containing ALEX, L-ALEX, and bimodal compositions followed the power law relation throughout the pressure range, with pressure exponents of , 0.5, and 0.28, respectively. Propellants containing V-ALEX powder demonstrated two different regimes: at pressures lower than 1.8 MPa, the burning rates were constant with respect to pressure; whereas for higher pressures, the burning rates followed the power law relation with and for powder diameters of 60 and 100 nm, respectively. The combustion efficiency, defined as the fraction of aluminum that oxidized during combustion, showed high values (90–95%) for all tested compositions.

The effects of SiO2 and TiO2 on the two-phase burning behavior of aqueous HAN propellant

Silica and titania nanoparticles were included at mass loadings of 1% and 3% in aqueous HAN propellants to evaluate their effects on liquid- and gas-phase decomposition and combustion. Both the liquid-phase and overall burning rates of propellant formulations were indirectly measured in a constant-volume strand burner filled with Argon from pressures of 3–22 MPa using a novel, pressure-based method developed by the authors in recent work. This approach provides overall burn times for propellants such as aqueous HAN which continue to burn beyond the disappearance of the liquid, making it superior to methods based solely on visual observation which only monitor the liquid surface regression. The presence of silica nanoparticles increased the liquid-phase burning rate in the low- and medium-pressure regimes (<10 MPa) and increased the overall burning rate at all pressures evaluated. The maximum amount of burning rate enhancement was realized at the lowest evaluated pressure (3 MPa) which corresponded to 80% and 670% increases in the liquid-phase and overall burning rates, respectively, for a silica loading of 1%, and 160% and 830% increases in the liquid-phase and overall burning rates, respectively, for a silica loading of 3%. The presence of titania did not measurably affect the liquid-phase burning rate, but it did increase the overall burning rate in the low-pressure regime (<5.7 MPa). This low-pressure overall burning rate enhancement was not amplified by further titania loading from 1% to 3% and was maximized at the lowest evaluated pressure (3 MPa) which corresponded to a 500% increase in the overall burning rate. The observed enhancements of the propellant's liquid-phase and overall burning rates were attributed to the presence of catalytic processes which diminish at higher pressures. This work represents the first time nanoparticle additives have been utilized to tailor the combustion of liquid HAN-based monopropellants.

Effects of particle size and morphology of NQ on thermal and combustion properties of triple-base propellants

Particle size and morphology of nitroguanidine (NQ) are two of the dominant factors that influence the thermal and combustion properties of triple-base propellants. However, the investigations of these effects and the corresponding mechanisms are insufficient. In this work, we examined the thermal properties and nonisothermal decomposition kinetics using differential scanning calorimetry (DSC) and thermogravimetry (TG). Combustion properties i.e. flame structure, melting layer and burning rate were determined on the bases of windowed strand burner (3 MPa), high-speed video camera, interrupted closed vessel, scanning electron microscope (SEM) and closed vessel, respectively. It was found that the decomposition process of triple-base propellant could be divided into two steps. The first step was mainly related to the decomposition process of nitroglycerin (NG) with higher activation energy and lower peak temperature, which was enhanced due to reactions between the added NQ and the decomposition product of NG (NO2), while the second step mainly concerned the decomposition process of nitrocellulose (NC) with lower activation energy and higher peak temperature. The flame structure of triple-base propellants was heterogeneous and closely attached to the condensed phase without dark zone. When the NQ particle size was decreased, the active site number of NQ was increased, enhancing the intensity of the reaction between NO2 and NQ; the decomposition peak temperature was decreased; the flame structure became more heterogeneous and brighter, resulting in more heat feedback from gas phase to the condensed phase. Thus, the burning rate was increased. On the other hand, the apparent activation energy and thickness of the melting layer of triple-base propellants were also increased with the decrease of NQ particle size, which meant that the sensitivities of burning rate to the conditions of gas phase were reduced. Therefore, the burning rate pressure exponent was decreased.

APPROPRIATENESS OF BURNING RATE MEASURING TECHNIQUE WITH THE MEASURING CONDITIONS AND REQUESTED DATA

Burning rate data are usually obtained in several ways. Small-scale ballistic evaluation motors, constant volume vessels, standard strand burners, and full-scale motors with good instrumentation are the most common methods used to measure the ballistic properties of solid propellants. The objective of this work is to describe the facilities developed by each measuring technique and validate the newly developed SV-2 for strand burning measurements at low and high pressures. Modifications were made to use the SV-2 closed bomb as a Crawford bomb. The sample holder and the closing tip were totally changed to enable burn rate measurements at the beginning of strand burn measuring. The burning rate of different commercial extruded double bases was measured using the modified SV-2 strand burning and compared with the burning rate resulting from subscale rocket motors. The results showed that the burning rates generated from both methods are comparable and close to each other. One firing from the subscale motor or strand burning is equal to one point on the P–u curve. The effect of the initial grain temperature on the combustion process was investigated using the SV-2 closed vessel method compared to the subscale rocket motor method. Changes in the ignition process were reported through the variations in the durations for different ignition zones. The burning criteria were studied through the P–t curves and the actual burning time. The burning parameters (pressure exponent n and temperature coefficient a) were calculated. The results showed that raising the initial grain temperature increases the ignitability, burning rate, and operating pressure. The suitability of the burning rate measuring technique is discussed. The subscale rocket motor method can be considered the method that most closely simulates reality despite the fact that it requires more money and effort and is time consuming. Strand burning measurements are favorable when measuring the burning rates at lower operating pressures (2–10 MPa), while the method is not suitable for investigating the initial grain temperature effect on the combustion process. The SV-2 closed vessel can be considered a more convenient method for measuring the burning rate from an economical point of view.

Energetic Nanocomposites as Burn Rate Catalyst for Composite Solid Propellants

AbstractIn the present study the effect of addition of Al/Fe2O3 based energetic nanocomposites (ENC) to non‐aluminized and aluminized propellants was investigated. Comparative study was carried out using n‐Fe2O3 as burn rate catalyst. Ferric oxide xerogel and Al/Fe2O3 ENC were synthesized by sol‐gel method. The dried xerogel was characterised by Scanning electron microscopy (SEM), Transmission electron microscopy (TEM), X‐ray diffraction (XRD) and X‐ray photoelectron spectroscopy (XPS). TEM and SEM images reveal presence of porous interconnected ferric oxide clusters of size 3–5 nm and intimate mixing of the ingredients at nanoscale in case of ENC. XRD analysis shows amorphous nature of the ferric oxide xerogel. Fe2O3 and FeO(OH) are the major constituents (&gt;84 %) of the xerogel as confirmed by XPS analysis. Thermal analysis of the ENC indicates that the composition is energetic in nature. Sensitivity analysis was carried out to assess the safety of the ENC during propellant processing. The differential scanning calorimetry (DSC) results for propellant samples show that the Al/Fe2O3 ENC is a better catalyst for AP decomposition than n−Fe2O3. Burn rate studies of ENC catalyzed propellants with μ‐Al and with n‐Al were carried out using acoustic strand burner. The results indicate that ENCs can be promising catalysts for composite propellants.

Studies on the Internal Ballistics of Composite Solid Rocket Propellants Incorporating Nano-Structured Catalysts

This paper deals with the analysis of burn rate using various catalysts of Iron Oxide and determining which gives the higher burn rate with low pressure variation. The Ammonium Perchlorate (AP) was obtained and ground into fine powder with the particle size ranging from 63 to 125 μm. The propellant strands were prepared with proportions by mixing AP with the binder (Hydroxyl Terminated Polybutadiene), the catalyst (Iron Oxide), curing agent (Isophorone diisocyanate) and the plasticizer (Dioctyladipate). The prepared propellant mixture was cured at around 63 deg C to get various propellant strands. The first strand was prepared with the absence of a catalyst to set an initial base of comparison with other Iron Oxide catalysts, namely, Flower Shaped, Micro and Nano, based on the size of the particles. The combustion process was carried out in a strand burner, which was in turn connected to a data acquisition system. The obtained output was analysed in the form of graphs. The burn rate was achieved by calculating the slope of the graph i.e . by calculating the difference between the highest and the lowest peak of the graph and dividing the total time by the answer. The experiment was repeated with the different catalyst types, as mentioned above, at different pressures. It was observed that the Nano shaped Iron Oxide exhibits better burning characteristics when compared to the rest with the pressure index of 0.792. In this paper, the various experiments carried out along with their procedures are explained in detail. The results obtained and the techniques used are also elaborately described in this paper.

Hydroxylammonium nitrate (HAN)-based green propellant as alternative energy resource for potential hydrazine substitution: From lab scale to pilot plant scale-up

Since 2000, Japan Aerospace Exploration Agency (JAXA) is involved in the research and development (R&D) of hydroxylammonium nitrate (HAN)-based liquid monopropellant, SHP163 which is composed of HAN, ammonium nitrate (AN), methanol and water for its possible utilization as an alternative of hydrazine toxic component. The freezing point of SHP163 is much lower, the density and the density-Isp are higher than those of hydrazine. The toxicity of SHP163 is also lower than hydrazine, therefore easy handling. Moreover, SHP163 monopropellant showed stable burning rate measured by strand burner and its thermal decomposition of HAN-based occurred in two steps. However, the catalytic process showed a violent exothermic reaction of HAN-based in one step at low temperature onset around 65 °C. The application of SHP163 on thruster firing tests showed that pressure slopes have been increased in the case of honeycomb catalysts and burning reactions have been improved by new synthesized Ir–CuO honeycomb catalyst.The obtained results on thruster 20 N type demonstrate that SHP163 is a good candidate as green propellant and proved the reason why it has been selected to be used as hydrazine alternative monopropellant for ``Innovative Satellite Technology'' project launched in 2016 by JAXA, Japan.

Development of Strand Burner Test by Using Aluminized AP/HTPB

One of the key factors that play an important role in the performance of a solid propellant rocket is dependent on its burning rate. A strand burner is a type of apparatus used to measure the propellant burning rate at elevated pressure. This study investigates the relation between burning rates of aluminized ammonium perchlorate at low pressure. Chamber pressure was varied from 1 atm, 3 atm, 5 atm and 7 atm. This study shows that propellant burning rate is about 50%-62% higher when the burn rate test is conducted at atmospheric condition compared to when it is done in inert gas. The investigation’s results also revealed an increasing propellant burning rate when the chamber pressure is increased. In conclusion, the burning condition and chamber pressure influences the propellant burning rate.

Combustion and Decomposition of Monomethylhydrazinium Nitrate

ABSTRACTMonomethylhydrazinium nitrate was synthesized by mixing the hypergolic bipropellant monomethylhydrazine and nitric acid under nonignition condition. The burn rate of this nitrate-based monopropellant is studied by using a high-pressure strand burner with optical access, coupled with a high-speed camera for capturing and analyzing the combustion characteristics. In addition, pyrolysis of this energetic material was investigated by a confined rapid pyrolysis setup with heating rates on the order of 2000 K/s, coupled to a rapid-scan Fourier transform infrared spectroscope. The decomposition mechanism of monomethylhydrazinium nitrate is discussed.

Factors Affecting the Primary Combustion Products of Boron-Based Fuel-Rich Propellants

Boron-based fuel-rich propellants with different formulation were prepared, and the condensed-phase products under different chamber pressures were collected in this study. The composition of boride (elemental boron, boron oxide, boron carbide, and boron nitride) in the condensed-phase products was measured by chemical analysis and the particle size distribution was determined using a laser particle size analyzer. Moreover, the burning rate and combustion temperature of the propellants were measured using a strand burner. The experimental results show that both the propellant burning rate and combustion temperature played an important role in the production of boride. There might be large amounts of boron carbide in the condensed-phase products which appear to have a negative effect on the combustion efficiency of the propellants studied in this work. The production of boron carbide could be reduced and combustion efficiency was improved by lowering the combustion temperature while increasing the burning rate. This study also shows that reducing the chamber pressure also causes a decrease in production of boron carbide.

Development of a Design Methodology Against Fast Cook‐Off Threat for Insensitive Munitions

AbstractThis work provides a design methodology for a typical energetic system to get Type V insensitive munition (IM) response against fast cook‐off threat. Experimental and numerical results for a generic test item filled with PBXN‐109 explosive are presented. Two key aspects of IM design against fast cook‐off threat are investigated, which are “time‐to‐reaction” and “critical ventilation area”. By dividing the problem to pre‐ignition and post‐ignition, the complexity of the problem is reduced. The first part of the study covers two‐dimensional and three‐dimensional simulations of the problem, where the generic test item is modelled using commercial CFD software (ANSYS – Fluent). The second part of the study reveals low pressure (2–10 MPa) burning characteristics of a PBXN‐109 by strand burner tests. After obtaining the pressure dependent burning rate, a conservation of mass equation is used to determine the chamber pressure using MATLAB Simulink software. Calculations are in reasonable agreement with the experimental results. Calculations are compared with the tests performed. The results are in reasonable agreement with some discrepancies of 8.9 % for time to reaction prediction and of 10.9 % for ventilation characteristics analyses. Possible reasons of these differences are discussed.

Thermal decomposition and combustion studies of catalyzed AN/KDN based solid propellants

Effects of catalyst on the thermal decomposition were envisaged for AN and AN/KDN mixtures as an oxidizer and on the composite propellants based on them. AN/KDN based oxidizer samples were prepared in the ratios of 75/25 and 50/50 by a co-crystallization method. Propellants were formulated based on AN and AN/KDN mixtures as an oxidizer, and HTPB as a fuel binder with the addition of catalyst. The catalysts used for present studies are, (i) cupric (II) oxide (CuO) and (ii) copper–cobalt based metal oxide (Cu–Co*). For all samples, 2% by weight of catalyst was added to the total weight of the samples. Thermal decomposition and burning rate studies were carried out for prepared propellant samples. All prepared oxidizers and propellants samples are subjected to DSC/TG studies using Simultaneous Thermal Analyzer (STA), while strand burners are used for burn rate determination of propellants in pressure regimes 100–4760kPa and at sub-atmospheric pressures regimes (<100kPa). Evolved gas analysis-QMS (EGA) was also performed to understand the decomposition mechanism of propellant samples. Propellants based on AN/KDN (50/50) in the presence of Cu–Co* showed the highest burning rate and also sustain combustion up to 40kPa. Propellants based on CuO was also observed to enhance the burning rate but lesser as compared to the propellants based on Cu–Co*.

EFFECT OF CU-CO MIXED METAL OXIDES ON THE COMBUSTION OF PSAN-HTPB BASED SOLID PROPELLANTS

Composite propellants based on ammonium perchlorate (AP) as oxidizer are state-of-the-art. However, global environmental impact restricts AP-based propellants because of their high chlorine exhaust. Therefore, efforts are on to innovate propellants with clean exhaust. Ammonium nitrate (AN) has, thus, regained importance in the field of propellants and explosives, more specifically in insensitive formulations, because of its clean burning and low hazard. However, the stumbling block for the application of AN - as solid propellant oxidizer has been its dimensional instability caused by phase transformation which results in increase of volume and porosity of the propellant grain. This can be overcome by the use of phase stabilized ammonium nitrate (PSAN). In the present study AN phase stabilized by incorporating 10% KN to prepare composite solid propellant. Copper-Cobalt mixed metal oxides have been synthesized via citric acid complexing method. Three different catalysts were prepared with different Cu-Co molar ratios 1, 0.5 And 2 in the initial reactants. The synthesized catalysts were then added to the PSANHTPB(Phase Stabilized Ammonium Nitrate- Hydroxyl Terminated Poly-Butadiene) composite formulations. These propellant samples were subjected to burn rate measurement in a Crawford High Pressure Strand Burner and thermal degradation studies in Simultaneous Thermal Analyzer (STA). The thermal decomposition and burn rate of the propellant is observed maximum on the addition of Copper-Cobalt mixed metal oxide catalyst with molar ratio 2. A total of six propellant samples catalyzed and noncatalyzed and one preheated virgin sample were undertaken for the above analysis and results obtained are discussed. It has been observed that except the propellant sample with Cu-Co-III and virgin propellant where PSAN was prepared by solid mixing, other propellant samples were hard to burn.

New HAN-based mixtures for reaction control system and low toxic spacecraft propulsion subsystem: Thermal decomposition and possible thruster applications

HAN-based liquid monopropellant has been studied for its possible substitution of hydrazine toxic component. The onset temperature of decomposition was provided by DTA-TG thermal analysis. Moreover, the effect of 5M HNO3 and 5M NH2OH addition on catalytic decomposition was also examined. The major products measured by mass spectrometer were N2, NO, N2O, NO2, H2O and NH3. Otherwise, the burning rates of HAN-based monopropellants were measured from the strand burner videos taken by high speed camera. The effect of methanol addition; as fuel; on the burning rates was demonstrated. On the other hand, HAN-based liquid monopropellant was tested and decomposed in 20N thruster and different catalysts were compared. Honeycomb catalysts were tested instead Shell 405 catalysts, the obtained data demonstrate the improvement of burning reactions and pressure slopes after HAN-based decomposition.

COMBUSTION AND PERFORMANCE STUDIES OF GLYCIDYL AZIDE POLYMER AND ITS MIXTURES AS HYBRID ROCKET FUEL

The combustion of glycidyl azide polymer (GAP) was investigated. The GAP fuel tests were conducted in a strand burner with nitrogen gas purge and in a lab-scale hybrid rocket with gas oxygen as oxidizer. In the strand burner, at intermediate pressure range 3−6 MPa, burning rates of both cured and uncured GAP were nearly the same and strongly depend on the chamber pressure. At pressure larger than 6 MPa, uncured GAP shows less dependence on pressure, and the GAP ratio of mixtures begins to have dominant role on burning rate. Transition pressure region was observed, which was in between the pressure dependence to composition dependence. The fluctuating burning rate was the main appearance in the transition pressure region. The Zel'dovich formula shows good prediction for burning rate at intermediate pressure range, but has much deviation at high pressure due to the surface instability. A smaller sample tube can restrict the burning surface from twisting, and the resulting burning rate shows good agreement with Zel'dovich formula prediction. Uncured 100% GAP can reach its maximum burning rate at certain pressure less than 10 MPa. Based on the assumption that condensed phase reaction is the only heat source which dominates the burning rate, the prediction of maximum burning rate of GAP mixture is also possible. In a lab-scale hybrid rocket test, GAP/PEG solid fuel demonstrated a very good burning rate under mild chamber pressure and oxygen flux. Instantaneous burning rate was successfully measured by ultrasonic techniques, and the results implied complex internal ballistic behavior.

Potential of ADN-based Ionic Liquid Propellant for Spacecraft Propulsion

As an alternative to hydrazine, ammonium dinitramide-based ionic liquid propellant (ADN-based ILP)was developed. It is mixture of the solid powders: ammonium dinitramide, monomethylamine nitrate, and Urea. It has a feature to be able to adjust the performance of monopropellant by change of the compositions because the constituents have different role: oxidizer, fuel, and freezing point depressant. By adjusting the components, it has high specific impulse and fuel rich and low freezing point comparing to hydrazine. The basic combustion characteristics of ADN-based ILP wereinvestigated in the strand burner tests forapplication to thruster. The self-sustainable combustion,the linear burning rates,and the flame temperatureswere shown in the results. The features of the ADN-based ILP were discussed on the basis of the results. As a problem of the thruster application, the ignition in vacuum was also discussed.

Iron Nanoparticle Additives as Burning Rate Enhancers in AP/HTPB Composite Propellants

AbstractBurning rate measurements were carried out for ammonium perchlorate/hydroxyl‐terminated polybutadiene (AP/HTPB) composite propellants with iron (Fe) nanoparticles as additives. Experiments were performed in a strand burner at pressures from 0.2 to 10 MPa for propellants containing approximately 80 % AP and Fe nanoparticles (60–80 nm) at concentration from 0 to 3 % by weight. It was found that the addition of 1 % Fe nanoparticles increased burning rate by factors of 1.2–1.6. Because Fe nanoparticles are oxidized on the surface and have high surface‐to‐volume ratio, they provide a large surface area of Fe2O3 for AP thermal decomposition catalysis at the burning propellant surface, while also providing added energy release due to the oxidation of nanoparticle sub‐shell Fe. The increase in burning rate due to Fe nanoparticle content is similar to the increase in burning rate caused by the addition of iron oxide (Fe2O3) particles observed in prior literature.

Pre-ignition confinement and deflagration violence in LX-10 and PBX 9501

In thermal explosions of the nitramine octahydro-1,3,5,7-tetranitro-1,3,5,7-tetrazocine (HMX)-based explosives LX-10 and PBX-9501, the pre-ignition spatial and temporal heating profile defines the ignition location. The ignition location then determines the extent of inertial confinement and the violence of the resulting deflagration. In this work, we present results of experiments in which ∼23 g cylinders of LX-10 and PBX 9501 in thin-walled aluminum confinement vessels were subjected to identical heating profiles but which presented starkly different energy release signatures. Post-explosion LX-10 containment vessels were completely fragmented, while the PBX 9501 vessels were merely ruptured. Flash x-ray radiography images show that the initiation location for the LX-10 is a few mm farther from the end caps of the vessel relative to the initiation location of PBX 9501. This difference increases deflagration confinement for LX-10 at the time of ignition and extends the pressurization time during which the deflagration front propagates in the explosive. The variation in the initiation location, in turn, is determined by the thermal boundary conditions, which differ for these two explosives because of the larger coefficient of thermal expansion and greater thermal stability of the Viton binder in LX-10 relative to the estane and bis(2,2-dinitropropyl) acetal/formal binder of the PBX 9501. The thermal profile and initiation location were modeled for LX-10 using the hydrodynamics and structures code ALE3D; results indicate temperatures in the vicinity of the ignition location in excess of 274 °C near the time of ignition. The conductive burn rates for these two explosives, as determined by flash x-ray radiography, are comparable in the range 0.1–0.2 mm/μs, somewhat faster than rates observed by strand burner experiments for explosives in the temperature range 150–180 °C and pressures up to 100 MPa. The thinnest-wall aluminum containment vessels presented here rupture at lower pressures, in the range 10 MPa, suggesting that moderately higher temperatures and pressures are present near the deflagration front. For these explosives, however the most important property for determining deflagration violence is the degree of inertial confinement.