The regression rate of seven three-dimensionally printable materials have been investigated and compared to a traditional fuel [hydroxyl-terminated polybutadiene (HTPB)]. The materials investigated include acrylonitrile-butadiene-styrene (ABS), polycarbonate (PC), low-density polyethylene (LDPE), polypropylene (PP), polyethylene terephthalate glycol (PETG), polyamide 66 (Nylon 66), and a UV-curable resin (CLEAR). Although all fuels can be three-dimensionally printed, the former six fuel grains were cast to avoid porosity effects introduced from additive manufacturing. The latter fuel was printed using stereolithography. All the fuels were analyzed using elemental analysis and calorimetry methods to determine their respective compositions of carbon, hydrogen, oxygen, and nitrogen along with their heats of formation. A laboratory-scale hybrid rocket motor was used to perform a series of short- (2.5-3.5 s) and long-duration (10 s) burns to measure each fuel's regression rates using thickness-over-time and ballistic reconstruction techniques. HTPB demonstrated the highest regression rate, followed by ABS and CLEAR, while the remaining fuels showed similar, lower regression rates. The experimental data set provided in the current study also provides baseline data that can be used for predictive modeling of additively manufactured fuel systems.

The regression rate of seven three-dimensionally printable materials have been investigated and compared to a traditional fuel [hydroxyl-terminated polybutadiene (HTPB)]. The materials investigated include acrylonitrile-butadiene-styrene (ABS), polycarbonate (PC), low-density polyethylene (LDPE), polypropylene (PP), polyethylene terephthalate glycol (PETG), polyamide 66 (Nylon 66), and a UV-curable resin (CLEAR). Although all fuels can be three-dimensionally printed, the former six fuel grains were cast to avoid porosity effects introduced from additive manufacturing. The latter fuel was printed using stereolithography. All the fuels were analyzed using elemental analysis and calorimetry methods to determine their respective compositions of carbon, hydrogen, oxygen, and nitrogen along with their heats of formation. A laboratory-scale hybrid rocket motor was used to perform a series of short- (2.5–3.5 s) and long-duration (10 s) burns to measure each fuel’s regression rates using thickness-over-time and ballistic reconstruction techniques. HTPB demonstrated the highest regression rate, followed by ABS and CLEAR, while the remaining fuels showed similar, lower regression rates. The experimental data set provided in the current study also provides baseline data that can be used for predictive modeling of additively manufactured fuel systems.

The thermal runaway (TR) of lithium-ion batteries (LIB) presents various hazards to people in the vicinity of the event. Various experiments have been developed to study LIB TR safety. The current study details a state-of-the-art LIB TR experiment developed at Texas A&M University and results obtained from a representative LIB TR test. The experiment encompasses an optically accessible, constant-volume bomb with accompanying diagnostics and controls. A proportional-integral-derivative (PID) was integrated into the experiment's thermal subsystem to permit control over the heating rate that thermally induces TR. A gas conditioning system provides control over the atmosphere within the test vessel prior to TR and allows for gas capture after the event. Results obtained from a representative test are presented to illustrate the utility of the experiment. A LG INR18650 cell at 100% state-of-charge (SoC) was heated at a rate of similar to 5 K/min in an atmosphere of air under standard conditions. Analysis of the thermal and pressure data revealed a TR onset temperature of similar to 175 degrees C and the production of 0.18 +/- 0.004 moles of gas during TR. These results agreed well with values found in LIB TR literature for similar test conditions. Additional gas compositional analyses indicated the presence of hydrogen, hydrocarbons, carbon dioxide, and carbon monoxide in a gas sample taken just after the TR event.

Metal fuels, such as aluminum (Al) and iron (Fe), can be added to composite solid propellants to improve their performance, such as specific impulse, density, and burning rate. In comparison to aluminum, iron can theoretically provide improved density specific impulse and higher flame temperatures; reduce condensed combustion product (CCP) concentration and the associated two-phase flow losses; and eliminate hydrochloric acid (HCl) in the exhaust products. A fundamental and quantitative understanding of metal particle aggregation and agglomeration processes in solid propellants is required to understand the underlying combustion mechanisms in these systems. In the current study, composite strand and laminate AP/HTPB/AP propellant samples loaded with Fe microparticles (∼45 μm in diameter) were burned at elevated pressures in an optically accessible strand bomb. Combustion processes were monitored with transient pressure diagnostics and a high-speed camera fitted with a high-magnification lens system (3.83 μm/pixel resolution) for the laminate propellant experiments. An automated image processing algorithm was developed to measure burning rates and ejected particle/agglomerate sizes and velocities. Time-resolved statistical distributions of both particle size and velocity are presented at elevated pressure for multiple laminate propellant experiments with a high degree of repeatability and low measurement error estimated as < ±5% and < ±1.5% for particle size and velocity, respectively. The incorporation of iron microparticles into the composite strand propellants yielded over a 20% increase in the global burning rate over the range of pressures evaluated (3.45–13.8 MPa). Similarly, the addition of iron to the fuel lamina in laminate propellant samples led to an approximately 30% increase in the global burning rate at the evaluated pressure (3.45 MPa). Additive particles were observed to eject near the oxidizer/fuel interface, or to melt, aggregate, coalesce, and agglomerate on the fuel lamina surface prior to ejection. Particle velocities are controlled by a balance of gravitational forces, drag forces imparted by expanding combustion product gases, and particle inertia. The observed combustion enhancements are attributed to the combined effects of catalytic mechanisms, increased radiation heat transfer, and local energy release from reacting iron particles. In addition, discussions on the image processing methods developed in the current study, corresponding potential sources of error, and prospective areas of improvement are provided. The experimental approach developed enables high-speed and high-magnification visualization of propellant combustion at high pressures and can be utilized to better understand the fundamental combustion behavior of energetic systems.

Implementing metal fuels, such as aluminum (Al), boron (B), and iron (Fe), into composite solid propellants can improve their performance (e.g., burning rate, specific impulse, density specific impulse, etc.). In the current study, composite AP/HTPB propellant strands loaded with micro-aluminum, iron, boron, and nano-aluminum were burned in an optically accessible strand bomb over a pressure range of 500-2,175 psia (3.45-15.0 MPa). Metal loadings were optimized for maximum performance using NASA's Chemical Equilibrium with Applications (CEA) code. In addition, analogous laminate AP/HTPB/AP samples loaded with the same energetic additives were burned at an elevated pressure of 500 psia (3.45 MPa). Combustion processes were monitored with a high-speed, high-magnification system resulting in 3.83-mu m/pixel resolution. The addition of all additives yielded increases in the global burning rates and significant alterations to the combustion behavior. Propellant formulations containing micro-aluminum performed as expected, moderately increased burning rates (15-37%), and yielded the ejection of significantly agglomerated particles. The replacement of aluminum with iron provided similar burning rates (14-31% increase) and significantly reduced the size and incidence of condensed particle ejections. The inclusion of boron led to a periodic fuel shedding phenomenon which artificially increased burning rates (65-140%) but is unrealistic for propulsive applications due to accompanying two-phase flow loss issues. Finally, the partial replacement of micro-aluminum with nano-aluminum provided significant increases in burning rate (similar to 110%) which were accompanied by large fuel fragment ejections that might be overcome with further optimization.

Lithium-ion batteries (LIBs) are widely utilized for energy storage in a broad range of applications, such as handheld electronics, emobility, spaceflight vehicles, etc. Thermal runaway (TR) and subsequent combustion of LIBs represent several significant hazards to consumers, including substantial energy release, jet flames, toxic gases, airborne particulates, and secondary explosions. Numerous experimental investigations in the literature have measured the energy release and chemical composition of off-gasses during LIB TR, but little theoretical work has been completed to this end. The current study focuses on prediction of product chemical composition, flame temperature, and energy release for LIBs undergoing TR failure through implementation of chemical equilibrium analyses (CEA). Theoretical calculations for various LIB electrolyte decomposition and combustion scenarios were compared to data available in the literature. Excellent agreement was observed between predicted and experimental measured heats of combustion for plain LIB electrolytes, indicating the global thermodynamic properties are well captured by the modeling approach. Theoretical product gas production (i.e., moles of gas per mole of electrolyte) and composition were compared to experimentally measured values from accelerated rate calorimetry experiments. The results indicate that general trends are well captured by the CEA modeling approach developed here and that the standard experimental protocols documented within the literature can be improved. Similar theoretical predictions for battery failure experiments (product gas composition and energy release) are also presented. The novel modeling framework presented here can be used in future work to evaluate LIB failure hazards for existing systems and to evaluate the safety of future designs. In addition, this modeling approach provides unique insight into how adjusting global battery chemistry (cathode, electrolyte, etc.) changes the potential hazards produced battery thermal runaway and failure. Figure 1

A common practice to deisgn composite propellants with high burning rates is to use several sizes of the oxidizer (e.g., ammonium perchlorate, AP) to increase the total amount of solids that can be packed into the composite propellant system. Small AP particles (<= 15 mu m) are difficult to obtain from commercial vendors due to restrictions derived from increased explosion hazards for superfine AP (SFAP). An in-house SFAP manufacturing process was devised herein using resonant acoustic mixing (RAM). A 90-mu m AP feedstock was utilized to produce SFAP batches with average particle sizes of approximately 2 mu m. Scanning electron microscopy (SEM) was used to characterize particle size distributions. Ballistic testing was conducted with an 80% SFAP composite propellant formulation was burned over a pressure range of 500 to 3,000 psi (3.45 to 20.68 MPa) and a burning rate of 1 in/s (25 mm/s) at 1,000 psi (6.89 MPa) was achieved without optimizing the propellant with additives.

Hybrid rocket engine (HRE) performance is dependent on fuel/oxidizer selection and fuel grain geometry. A literature review was performed to identify key trends and findings related to the application of the additive manufacturing (AM) of fuel systems for HREs. The effects of complex combustion port geometries, embedded structures, and end-burning systems, along with the use of metallic additives, turbulators, diaphragms, gel-like fuels, powdered fuels, liquid fuels, and liquifying fuels and their impact on regression rates, combustion efficiencies, and/or mechanical strength are thoroughly documented here. In general, the application of AM to HRE fuels can be implemented to increase regression rates and combustion efficiency, and tailor HRE designs. Chemical equilibrium analysis computations were completed to characterize the theoretical performance of HTPB and common AM fuels (ABS, PLA, PC, PMMA, Nylon 6, and a UV-based fuel) with common oxidizers (LOX and N2O). AM fuels exhibit a similar theoretical performance as the commonly used HTPB fuel, and proper selection of the fuel can yield improved performance and design metrics. Development of AM approaches for HRE fuel design have significantly expanded their design trade space and should enable the competitive application of HREs for future propulsion missions.

LMP-103S is a storable, green, ammonium dinitramide (ADN)-based liquid monopropellant with signifi-cant potential owing to its improved specific impulse ( >= 6%) and density ( -24%) in comparison to hy-drazine, and its superior safety and handling characteristics. LMP-103S already has a significant history of thruster testing and flight heritage. However, its baseline combustion behavior is a surprisingly under-studied topic. Constant-volume strand burner experiments with LMP-103S are reported herein for pres-sures between 0.69 and 34.5 MPa (100 - 5,000 psia). Linear burning rates were measured from transient pressure data and high-speed video. Visible and near-IR emission spectra were recorded as well. Phe-nomenological combustion behaviors are reported for the entire pressure range as deduced from these diagnostics, and two clear pressure regimes are observed. Hydrodynamic instabilities and supercritical combustion processes dominate the combustion behavior in the low-and high-pressure regimes, respec-tively. Supporting experiments include thermal cook-off of LMP-103S and impact sensitivity testing of ADN crystals extracted from LMP-103S solution. The data presented herein significantly expand upon those available in the literature for LMP-103S, and the phenomenological combustion behavior deduced from transient, high-speed video data improve upon the state-of-the-art knowledge regarding ADN-based liquid monopropellants.(c) 2023 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

The competing flames model, also termed the Beckstead–Derr–Price model, for steady-state heterogeneous propellant combustion has been widely used but has not been sufficiently updated in decades or compared to modern propellant combustion databases. In the current study, historical competing flames modeling approaches were thoroughly documented; and an improved framework was outlined and updated to include several improvements, such as variable flame temperatures, specific heat capacities, and latent heat terms. Model parameters were initially taken from previous literature, but the fuel and diffusion flame parameters were optimized based on a compiled database of unimodal propellant burning rates from the literature spanning a wide range of ammonium perchlorate (AP) particle sizes ([Formula: see text]), AP mass concentrations (70–87.5%), and combustion pressures (0.7–20.7 MPa). The improved model was compared to AP monopropellant, unimodal, and multimodal propellant burning rate databases from the literature. General dependencies of the burning rate-to-oxidizer concentration and size were accurately captured. The predictive capability of the improved model for AP monopropellant burning rates and unimodal propellant formulations was excellent, where the only significant discrepancies were noted for very fine AP particles ([Formula: see text]). Model predictions for multimodal formulations were moderate and could be improved by alternative pseudopropellant apportionment and statistical accounting schemes.

The ignition of flammable liquids on hot surfaces is both a safety concern to industry and an important phenomenon for some combustion devices. Hot-surface ignition (HSI) is a function of fluid properties and environmental parameters, primarily surface temperature. In the current study, a laboratory-scale experiment was designed and characterized to determine the HSI characteristics of flammable liquids. A stainless-steel block is heated using embedded electrical resistance heaters and single drops of liquid fuel are dispensed onto removable test surfaces. The experimental apparatus is well-characterized and provides a uniform surface temperature within the range of 25-750 °C with high accuracy (±5 °C) and control. Experimental measurements include temperature-dependent ignition probability and time-to-ignition. In addition, high-speed video was utilized to further probe the underlying phenomena that govern HSI behavior. Representative HSI data for heptane, nitromethane, and lubricant oils are presented to illustrate the utility of the developed experiment in providing valuable HSI data and elucidating the fundamental parameters that influence HSI behavior.

Lithium-ion battery (LIB) electrolytes are generally composed of an organic hydrocarbon and lithium salt (e.g., LiPF6). The effect of salt addition to the electrolyte solvent increases its density. Experimental density data for basic and realistic, multicomponent mixtures are scarce in the literature, and a predictive model for electrolyte solution density is not yet available. This work resolves to create a predictive method that can be used to estimate the density of these mixtures. A predictive method which accounts for intramolecular forces in the electrolyte mixture was developed and fit to the sparse density data available in the literature. The model exhibited high accuracy for single component electrolyte mixtures with most predictions falling within 1% of measured values and all predictions falling within 5%. The model was further extended to more realistic, multi-solvent electrolyte mixtures which exhibited similar accuracy, seen in Fig. 1. It was noted that the multi-component predictions were moderately worse than for the single component mixtures. The source of this disparity is not clear but may be derived from uncertainty in the data reported by electrolyte vendors and/or the presence of solvent-to-solvent interactions in the solution which are not accounted for in the current modeling framework. In addition to the density model, an accurate method for computing absolute molar and mass fractions of multi-solvent mixtures with specified volumetric concentrations (e.g., 1.2 M LiPF6 in 1:1:1 %vol. EC/ DEC/DMC) is also described for when that is the only known relation for a solution. Computed mass concentration values yielded an average error of 1.4% with a maximum value of 9.6%. The molar and mass approaches were combined to yield a modeling framework capable of computing molar concentrations and densities of all LIB electrolyte solutions based on LiPF6 loaded in any combination of ethylene carbonate (EC), ethyl methyl carbonate (EMC), diethyl carbonate (DEC), dimethyl carbonate (DMC), and propylene carbonate (PC). The largest fault with the final modeling system is that it suffers from a lack of sufficient, high-quality experimental data for validation of its accuracy over a wide range of electrolyte formulations and conditions. This model’s accuracy is dependent on the experimental data utilized for correlation. Accordingly, more data from differing sources will help increase the accuracy and confidence of the model developed herein and allow it for expansive use in LIB fields. Future work will be required to extend the current approach to any novel electrolyte formulations developed in the research and development sector. Figure 1

Ammonium perchlorate (AP) is a common oxidizer in composite propellants, and its combustion behavior can be tailored by micro- and nano-metal oxide catalysts. Ammonium perchlorate pellets were manufactured with micro- and nano-titanium oxide (TiO2) and iron oxide (Fe2O3) at several mass loadings (0-3%) to isolate their effects on AP. The samples were burned from 3.45 to 34.5 MPa (500-5000 psi) in a constant-volume strand bomb. Microscopy characterization was completed for both additives and representative pellets. The incorporation of 1% micro-Fe2O3 yielded the highest burning rate within the investigated pressure range. All micro-additive formulations increased the burning rate at pressures ranging from roughly 13.45 to 17.24 MPa (1950-2500 psi) and from 4.50 to 8.60 MPa (650-1250 psi) for micro-TiO2 and micro-Fe2O3, respectively, and these effects were dependent on catalyst mass loading. The burning rates were increased at pressures greater than 11.27 (1620 psi) and 8.27 MPa (1200 psi) for nano-TiO2 and nano-Fe2O3, respectively. Both nano-additives yielded burning rates that were independent of mass loading within the range of evaluated concentrations (0.25-1%). In all cases, the presence of the catalytic additives removed the negative slope generally observed in the burning rate curve of plain AP. Higher low-pressure deflagration limits were observed for nano-additive formulations, potentially due to radiative heat losses.

Hydroxyl-terminated polybutadiene (HTPB) is a common ingredient in rocket propellants, but its thermochemical properties (composition, density, and heat of formation) are not well defined. A survey of the literature and thermochemical databases indicated wide ranges for these properties, especially heat of formation. Six group additive schemes were used to estimate the heat of formation of HTPB and analyze the effects of hydroxyl functionalization and curing reactions. Good agreement is observed between the methods for HTPB isomer units, but cumulative differences result in significant disparities for larger, practical polymers. Increased hydroxyl functionality and the curing reaction are predicted to yield nonnegligible decreases in the heat of formation. The heat of formation of propellant-grade, isophorone-diisocayante-cured HTPB R-45M (C213.8H323.0O4.6N2.3) was computed as 342 kJ/mol or 114 kJ/kg. Chemical equilibrium analyses were completed for solid propellants composed of ammonium perchlorate and HTPB, and for hybrid rocket engines based on HTPB burning with liquid oxygen or nitrous oxide, where the heat of formation of HTPB was varied within a reasonable range. The chemical equilibrium analysis computations indicated that combustion gas properties and theoretical propellant performance can vary up to 5% within practical operating conditions for the range of HTPB heats of formation implemented.

The competing flames model, also termed the Beckstead-Derr-Price (BDP) model, for steady-state, heterogeneous propellant combustion has been widely utilized, but has not been sufficiently updated in decades or compared to modern propellant combustion databases. In the current study, a competing flames model framework was outlined and updated to include several improvements, such as variable flame temperatures, specific heat capacities, and latent heat terms. Model parameters were initially taken from previous literature, but the fuel and diffusion flame parameters were optimized based on a compiled database of unimodal propellant burning rates from the literature spanning a wide range of AP particle sizes (5-500 μm), AP mass concentrations (70-87.5%), and combustion pressures (0.7-20.7 MPa). The improved model was compared to AP monopropellant, unimodal, and multimodal propellant burning rate databases from the literature. General dependencies of the burning rate to oxidizer concentration and size were accurately captured. The predictive capability of the improved model for AP monopropellant burning rates and unimodal propellant formulations was excellent where the only significant discrepancies noted were for very fine AP particles (< 10 μm). Model predictions for multimodal formulations were moderate and could be improved by alternative pseudopropellant apportionment and statistical accounting schemes.

To assess the fire hazard associated with venting gases coming from a lithium-ion battery during a thermal runaway, a mixture representative of such venting gas was determined by averaging 40 gas compositions presented in the literature. The final mixture is composed of C3H8, C2H6, C2H4, CH4, H2, CO, and CO2. The combustion properties of this mixture were determined using various combustion devices: shock tubes for ignition delay time measurements in air and for H2O time histories in very dilute mixtures (99% Ar), as well as a closed bomb to measure the laminar flame speeds. Experiments were performed at around atmospheric pressure and for several equivalence ratios in all cases. Several detailed kinetics models from the literature were assessed against the data generated with this very complex mixture, and it was found that modern detailed kinetics mechanisms were capable of appropriately predicting the combustion properties of thermal runaway gases from a battery in most cases, with the NUIGMech 1.1 model being the most accurate. A numerical analysis was conducted with the two most modern models to explain the results and highlight the most important reactions.

Hybrid rockets have distinct advantages over their pure solid or liquid propellant counterparts, and their performance can be improved by inclusion of metal additives. Several metallic additives (micro-Al, micro-Ti, micro-Mg, micro-Zr, nano-Al, nano-B, and Mg-coated nano-B) were selected as potential candidates for hybrid rocket applications and characterized by applicable microscopy techniques. The regression rates and combustion efficiencies of plain HTPB and HTPB loaded with each additive at various concentrations (10%, 20%, and 30% by mass) burning in GOX were evaluated at moderate oxidizer mass fluxes (10-150 kg/m(2)-s) and pressures (<= 0.86 MPa, 125 psia). In general, the inclusion of any of the metallic additives led to a reduction in the regression rate and did not significantly change the combustion efficiency. The only exceptions were fuel formulations containing micro-Zr, which yielded a moderate (10-20%) increase in the regression rate at a concentration of 10%. The observed trends were more prevalent at higher oxidizer mass fluxes and higher additive loadings. The reductions in regression rate were attributed to heat transfer blocking effects derived from accumulation of additive particles on the fuel surface layer. These phenomena were especially prevalent in highly loaded fuel formulations containing the nano-additives that exhibited unstable combustion and periodic surface-layer shedding. Zirconium appears to be the best metallic additive available since it can yield the highest theoretical density-specific impulse under the lowest O/F operation ratio without resulting in decrements to overall performance. Notably, combustion efficiency data for all fuel formulations were well correlated to the combustion residence time, and high combustion efficiencies (>95%) were achievable when a satisfactory residence time (similar to 75 ms) was realized.

Routine health information systems (RHIS) comprise data collected at regular intervals at public, private, and community-level health facilities and institutions and health programs. This article looks at how this data may be used for evaluations, and the reasons behind why some are optimistic about this and some have concerns.

Hybrid rockets have many advantages over pure solid or liquid propellant rockets, but low solid fuel regression rates and correspondingly low thrust have hindered their application to operational systems. Paraffin-based fuels regress significantly faster than traditional polymeric formulations, such as HTPB, and paraffin inclusion in HTPB represents a potential tool for performance augmentation in hybrid rockets. A survey of the available literature indicated disparities regarding the utility of this approach which are resolved herein. Fuel specimen consisting of plain HTPB; plain paraffin; and HTPB loaded with molten macrocrystalline paraffin wax (10-75%) or solid microcrystalline paraffin particles (10-60%) were manufactured and evaluated for their thermal decomposition and ballistic properties. Fuel samples were heated (10 IC/min) in an argon atmosphere in simultaneous TGA/DTA experiments. The inclusion of macrocrystalline paraffin enhanced the low-temperature decomposition of HTPB, while the inclusion of microcrystalline paraffin had the opposite effect. The prepared fuel grains were burned in gaseous oxygen on one of two lab-scale hybrid rockets over a range of oxidizer mass fluxes (5-430 kg/m(2)-s) and pressures (0.5- 1.0 MPa). The plain macrocrystalline paraffin fuel exhibited a 300% increase in regression rate over plain HTPB. However, none of the mixed-fuel formulations exhibited notable, if any, regression rate enhancement at the evaluated operating conditions. First principles modeling was completed for the combustion of plain HTPB, plain paraffin, and mixed-fuel systems comprised of HTPB containing molten liquid paraffin or solid paraffin particles. The combustion of mixed-fuel systems is dominated by the pyrolysis of HTPB which does not allow for the formation of a melt layer at the fuel surface, such that any enhancement is due to an increase in the vaporization rate of the fuel and not entrainment effects. This study was the first to concurrently evaluate the inclusion of both molten liquid paraffin and solid paraffin particles in HTPB and demonstrated a lack of performance augmentation with either strategy in two separate laboratories. The results presented herein resolve the disparities in the literature and indicate that paraffin inclusion in HTPB is not a viable means for tailoring the combustion behavior of hybrid rocket systems. (C) 2021 The Combustion Institute. Published by Elsevier Inc. All rights reserved.