Bibliografie tematiche / Solid propellant

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Letteratura scientifica selezionata sul tema "Solid propellant"

Autore: Grafiati

Pubblicato: 4 giugno 2021

Ultima modifica: 7 dicembre 2024

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Articoli di riviste sul tema "Solid propellant"

Jain, Prakhar, Vineet Kumar Rathi e Shelly Biswas. "Study of Aging Characteristics for Metalized HTPB Based Composite Solid Propellants Stored in Ambient Conditions". Defence Science Journal 74, n. 5 (29 agosto 2024): 615–26. http://dx.doi.org/10.14429/dsj.74.19786.

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The aging of any propellant is defined as the change in the physical, chemical, and performance parameters of solid rocket propellants. The propellant’s service life and aging properties are important parameters of the study, especially for missiles and other defense applications. Hydroxyl-terminated polybutadiene (HTPB) based composite solid propellants with ammonium perchlorate (AP) are the most prominently used propellants in the operations of solid rocket motors in the defense and space sectors. Thus, studying this composite solid propellant is of essential when determining ambient service life. Performance parameters studied in this research are burn rate under high-pressure conditions in Crawford bomb setup, Thermogravimetric Analysis, and Fourier Transform Infrared Spectroscopy (FTIR). SEM and X-ray diffraction (XRD) analysis of the aged sample were also conducted to ascertain the chemical composition and morphological changes in the samples. Naturally aged propellant strands manufactured in different years have been compared with freshly prepared ones to establish a trend for deriving conclusions. The results from different analysis techniques, FTIR, XRD, and FESEM, depicted that oxidation of metals happens while aging of propellant due to atmospheric moisture, and the metal oxides prominently affect the propellant chemical composition and decomposition process of the propellant samples. The ballistic properties of the aluminium added samples showed an increment in burn rate. In contrast, the bimetal addition of aluminium and magnesium combined as an additive decreased the ballistic burn rate.

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Aziz, Amir, Rizalman Mamat, Wan Khairuddin Wan Ali e Mohd Rozi Mohd Perang. "Review on Typical Ingredients for Ammonium Perchlorate Based Solid Propellant". Applied Mechanics and Materials 773-774 (luglio 2015): 470–75. http://dx.doi.org/10.4028/www.scientific.net/amm.773-774.470.

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Ammonium perchlorate (AP) based solid propellant is a modern solid rocket propellant used in various applications. The combustion characteristics of AP based composite propellants were extensively studied by many research scholars to gain higher thrust. The amount of thrust and the thrust profile, which may be obtained from a specific grain design, is mainly determined by the propellant composition and the manufacturing process that produces the solid propellant. This article is intended to review and discuss several aspects of the composition and preparation of the solid rocket propellant. The analysis covers the main ingredients of AP based propellants such as the binder, oxidizer, metal fuel, and plasticizers. The main conclusions are derived from each of its components with specific methods of good manufacturing practices. In conclusion, the AP based solid propellant, like other composite propellants is highly influenced by its composition. However, the quality of the finished grain is mainly due to the manufacturing process.

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Poryazov, V. A., K. M. Moiseeva e A. Yu Krainov. "NUMERICAL SIMULATION OF COMBUSTION OF THE COMPOSITE SOLID PROPELLANT CONTAINING BIDISPERSED BORON POWDER". Vestnik Tomskogo gosudarstvennogo universiteta. Matematika i mekhanika, n. 72 (2021): 131–39. http://dx.doi.org/10.17223/19988621/72/11.

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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.

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S A, Reshmitha Shree, Saif Ahmed Ansari, Steven Raj e Sukh Arora. "EVALUATION OF SOLID ROCKET PROPELLANTS FOR LOW EARTH ORBIT". INTERANTIONAL JOURNAL OF SCIENTIFIC RESEARCH IN ENGINEERING AND MANAGEMENT 08, n. 008 (31 agosto 2024): 1–3. http://dx.doi.org/10.55041/ijsrem37252.

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In contemporary aerospace technology, solid rocket propellants are commonly employed in the thriller and boost of space cargoes to LEO. These are reliable, uncomplicated, and high in thrust; and thus, can be used in both military as well as civil space ventures. This paper provides a comprehensive evaluation of five prominent solid rocket propellants: Five categories namely Ammonium Perchlorate Composite Propellant (APCP), Double-Base Propellant, Composite Modified Double Base (CMDB) Propellant, Hydroxyl-terminated polybutadiene (HTPB)-Based Propellant, and Ammonium Nitrate Composite Propellant (ANCP) have been identified. This also falls under the evaluation and includes aspects such as historical bases to see how the formulation of the propellant has changed over time regarding the development of propellant technology. In the case of each propellant, the chemical properties of the mix and reactions that its fundamental are analyzed. It also shares details regarding the changes and advancements made in the little while to enhance capability, safety, and environmental compatibility of these fuels, such as synthesizing fuels for liquid bipropellants. Some of the modern applications of technologies within the active spectrum of aerospace operations are explained in more detail to provide clients with the most versatile and essential information. Among the essentials of this work, the complex calculation of thrusts developed by the specific propellant with corresponding graphics has significant importance. These estimations afford chances of comparing the efficiency of the propellant in terms of producing thrusts as well as achieving the indicated optimal performance. Thus, the main objective of this study is to identify which of the solid rocket propellants to use in LEO missions and if any, what modifications could be made to improve them. Thus, the purpose of this paper is to contribute to the development of the constantly evolving branch of rocket propulsion by discussing the benefits and shortcomings of each propellant. These findings and recommendations could be useful in perfecting the recipes for the propellants and improving the efficiency of the subsequent missions to planets. Keywords: Aerospace, Solid rocket propellants, Thrust, Low earth orbit.

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Zhang, Jing, Zhen Wang, Shixiong Sun e Yunjun Luo. "Preparation and Properties of a Novel High-Toughness Solid Propellant Adhesive System Based on Glycidyl Azide Polymer–Energetic Thermoplastic Elastomer/Nitrocellulose/Butyl Nitrate Ethyl Nitramine". Polymers 15, n. 18 (5 settembre 2023): 3656. http://dx.doi.org/10.3390/polym15183656.

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Glycidyl azide polymer (GAP)–energetic thermoplastic elastomer (GAP-ETPE) propellants have high development prospects as green solid propellants, but the preparation of GAP-ETPEs with excellent performance is still a challenge. Improving the performance of the adhesive system in a propellant by introducing a plasticizer is an effective approach to increasing the energy and toughness of the propellant. Herein, a novel high-strength solid propellant adhesive system was proposed with GAP-ETPEs as the adhesive skeleton, butyl nitrate ethyl nitramine (Bu-NENA) as the energetic plasticizer, and nitrocellulose (NC) as the reinforcing agent. The effects of the structural factors on its properties were studied. The results showed that the binder system would give the propellant better mechanical and safety properties. The results can provide a reference for the structure design, forming process, and parameter selection of high-performance GAP-based green solid propellants.

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Zhang, Jing, Zhen Wang, Shixiong Sun e Yunjun Luo. "Influence of Solid Filler on the Rheological Properties of Propellants Based on Energetic Thermoplastic Elastomer". Materials 16, n. 2 (13 gennaio 2023): 808. http://dx.doi.org/10.3390/ma16020808.

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Glycidyl azide polymer-energetic thermoplastic elastomer propellant (GAP-ETPE) has high development prospects as a green solid propellant, although the preparation of GAP-ETPE with excellent performance is still a challenge. Focusing on the demand of high-strength solid propellants for free-loading rocket motors, a GAP-ETPE model propellant with excellent overall performance was prepared in this work, and the influence of adhesive structure characteristics on its fluidity was studied. Furthermore, the influence of filler on the rheological properties of the model propellant was investigated by introducing hexogen (RDX) and Al, and a corresponding two-phase model was established. The results may provide a reference for the structural design, molding process, and parameter selection of high-performance GAP-based green solid propellants.

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Abdullah, Mohamed, F. Gholamian e A. R. Zarei. "Noncrystalline Binder Based Composite Propellant". ISRN Aerospace Engineering 2013 (24 settembre 2013): 1–6. http://dx.doi.org/10.1155/2013/679710.

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This study reports on propellants based on cross-linked HTPE binder plasticized with butyl nitroxyethylnitramine (BuNENA) as energetic material and HP 4000D as noncrystalline prepolymer. This binder was conducted with solid loading in the 85%. The results showed an improvement in processability, mechanical properties and burning rate. In addition, its propellant delivers (about 6 seconds) higher performance (specific impulse) than the best existing composite solid rocket propellant. Thermal analyses have performed by (DSC, TGA). The thermal curves have showed a low glass transition temperature () of propellant samples, and there was no sign of binder polymer crystallization at low temperatures (−50°C). Due to its high molecular weight and unsymmetrical or random molecule distributions, the polyether (HP 4000D) has been enhanced the mechanical properties of propellants binder polymer over a large range of temperatures [−50, 50°C]. The propellants described in this paper have presented high volumetric specific impulse (>500 s·gr·cc−1). These factors combined make BuNENA based composite propellant a potentially attractive alternative for a number of missions demanding composite solid propellants.

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Glascock, Matthew S., Joshua L. Rovey e Kurt A. Polzin. "Impulse and Performance Measurements of Electric Solid Propellant in a Laboratory Electrothermal Ablation-Fed Pulsed Plasma Thruster". Aerospace 7, n. 6 (30 maggio 2020): 70. http://dx.doi.org/10.3390/aerospace7060070.

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Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. This behavior may enable the propellant to be used in multimode propulsion systems utilizing the ablative pulsed plasma thruster. The performance of an electric solid propellant operating in an electrothermal ablation-fed pulsed plasma thruster was investigated using an inverted pendulum micro-newton thrust stand. The impulse bit and specific impulse of the device using the electric solid propellant were measured for short-duration test runs of 100 pulses and longer-duration runs to end-of-life, at energy levels of 5, 10, 15 and 20 J. Also, the device was operated using the current state-of-the-art ablation-fed pulsed plasma thruster propellant, polytetrafluoroethylene (PTFE). Impulse bit measurements for PTFE indicate 100 ± 20 µN-s at an initial energy level of 5 J, which increases linearly with energy by approximately 30 µN-s/J. Within the error of the experiment, measurements of the impulse bit for the electric solid propellant are identical to PTFE. Specific impulse when operating on PTFE is calculated to be about 450 s. It is demonstrated that a surface layer in the hygroscopic electric solid propellant is rapidly ablated over the first few discharges of the device, which decreases the average specific impulse relative to the traditional polytetrafluoroethylene propellant. Correcting these data by subtracting the early discharge ablation mass loss measurements yields a corrected electric solid propellant specific impulse of approximately 300 s.

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Kohga, Makoto, Tomoki Naya e Kayoko Okamoto. "Burning Characteristics of Ammonium-Nitrate-Based Composite Propellants with a Hydroxyl-Terminated Polybutadiene/Polytetrahydrofuran Blend Binder". International Journal of Aerospace Engineering 2012 (2012): 1–9. http://dx.doi.org/10.1155/2012/378483.

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Ammonium-nitrate-(AN-) based composite propellants prepared with a hydroxyl-terminated polybutadiene (HTPB)/polytetrahydrofuran (PTHF) blend binder have unique thermal decomposition characteristics. In this study, the burning characteristics of AN/HTPB/PTHF propellants are investigated. The specific impulse and adiabatic flame temperature of an AN-based propellant theoretically increases with an increase in the proportion of PTHF in the HTPB/PTHF blend. With an AN/HTPB propellant, a solid residue is left on the burning surface of the propellant, and the shape of this residue is similar to that of the propellant. On the other hand, an AN/HTPB/PTHF propellant does not leave a solid residue. The burning rates of the AN/HTPB/PTHF propellant are not markedly different from those of the AN/HTPB propellant because some of the liquefied HTPB/PTHF binder cover the burning surface and impede decomposition and combustion. The burning rates of an AN/HTPB/PTHF propellant with a burning catalyst are higher than those of an AN/HTPB propellant supplemented with a catalyst. The beneficial effect of the blend binder on the burning characteristics is clarified upon the addition of a catalyst. The catalyst suppresses the negative influence of the liquefied binder that covers the burning surface. Thus, HTPB/PTHF blend binders are useful in improving the performance of AN-based propellants.

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He, Zhong Qi, Ke Zhou e Shu Pan Yin. "Security Analysis on Single-Screw Extrusion Process of Solid Propellant by Numerical Simulation". Advanced Materials Research 997 (agosto 2014): 605–9. http://dx.doi.org/10.4028/www.scientific.net/amr.997.605.

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Solid propellants are flammable, explosive, which maybe initiated unexpectedly under specific excitation conditions through high-speed screwing. Though single-screw extrusion process for solid propellant occurred a large proportion accidents, the rheological parameters of the propellant material were difficult to get yet, which affects the process safety greatly. In this paper, the CFD software, POLYFLOW, and a numerical simulation method were adopted to investigate the rheological parameters in single-screw extrusion process of solid propellant. By analyzing the rheological state of a solid propellant in the screw extrusion process, the applicable numerical model was established with a substitute material for the propellant. As a result, distributions of key parameters, such as material temperature, pressure, viscosity, were obtained. The simulation shows that the material has relatively higher pressure, temperature and smaller viscosity at the screw edge, where solid propellant components were mixing and plasticizing severely, also where needs pay much attention to for safety reasons.

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Più fonti

Tesi sul tema "Solid propellant"

Smyth, Daniel A. "Modeling Solid Propellant Ignition Events". BYU ScholarsArchive, 2011. https://scholarsarchive.byu.edu/etd/3125.

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This dissertation documents the building of computational propellant/ingredient models toward predicting AP/HTPB/Al cookoff events. Two computer codes were used to complete this work; a steady-state code and a transient ignition code Numerous levels of verification resulted in a robust set of codes to which several propellant/ingredient models were applied. To validate the final cookoff predictions, several levels of validation were completed, including the comparison of model predictions to experimental data for: AP steady-state combustion, fine-AP/HTPB steady-state combustion, AP laser ignition, fine-AP/HTPB laser ignition, AP/HTPB/Al ignition, and AP/HTPB/Al cookoff. A previous AP steady-state model was updated, and then a new AP steady-state model was developed, to predict steady-state combustion. Burning rate, temperature sensitivity, surface temperature, melt-layer thickness, surface species at low pressure and high initial temperature, final flame temperature, final species fractions, and laser-augmented burning rate were all predicted accurately by the new model. AP ignition predictions gave accurate times to ignition for the limited experimental data available. A previous fine-AP/HTPB steady-state model was improved to predict a melt layer consistent with observation and avoid numerical divergence in the ignition code. The current fine-AP/HTPB model predicts burning rate, surface temperature, final flame temperature, and final species fractions for several different propellant formulations with decent success. Results indicate that the modeled condensed-phase decomposition should be exothermic, instead of endothermic, as currently formulated. Changing the model in this way would allow for accurate predictions of temperature sensitivity, laser-augmented burning rate, and surface temperature trends. AP/HTPB ignition predictions bounded the data across a wide range of heat fluxes. The AP/HTPB/Al model was based upon the kinetics of the AP/HTPB model, with the inclusion of aluminum being inert in both the solid and gas phases. AP/HTPB/Al ignition predictions bound the data for all but one source. AP/HTPB/Al cookoff predictions were accurate when compared to the limited data, being slightly low (shorter time) in general. Comparisons of AP/HTPB/Al ignition and cookoff data showed that the experimental data might be igniting earlier than expected.

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Lowe, C. "CFD modelling of solid propellant ignition". Thesis, Cranfield University, 1996. http://hdl.handle.net/1826/3921.

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Solid propellant is the highly energetic fuel burnt in the combustion chamber of ballistic weapons. It is manufactured, for this purpose, in either granular or stick form. Internal ballistics describes the behavior within the combustion chamber throughout the ballistic cycle upto projectile exit from the muzzle of the gun barrel. Over the last twenty years this has been achieved by modelling the process using two-phase flow equations. The solid granules or sticks constitute the first phase, which can be assumed to be incompressible over typical pressure ranges within the chamber. The gas-phase is composed of both the original ambient gas contained around the propellant and additional gas produced by the propellant gasifying on heating. Equations can be derived that describe the conservation of mass, momentum and energy in terms of average flow variables. The equations are a highly non-linear system of partial-differential- equations. High-speed flow features are observed in internal ballistics and ordinary fini te- difference methods are unsuitable numerical methods due to inaccurate prediction of discontinuous flow features. Modern shock-capturing methods are employed, which solve the system of equations in conservation form, with the ability to capture shocks and contact discontinuities. However, although the numerical solutions compare well with experiment over the bulk of the combustion chamber, the ignition models used in internal ballistics are unreliable. These are based on either gas or solid-surface temperature achieving some empirically measured 'ignition temperature' after which the propellant burns according to an empirical pressure dependent burning law. Observations indicate that this is not an adequate representation of ignition. Time differences between first solid gasification and ignition imply two distinct processes occurring. ]Further, ignition occurring in gas-only regions indicates that ignition is controlled by a gas-phase reaction. This thesis develops simple ideas to describe possible mechanisms for these physical observations. The aim is to provide an improved model of the ignition of solid propellant. A two stage reaction process is described involving endothermic gasification of the solid, to produce a source of reactant gas, followed by a very exothermic gas-phase ignition reaction. Firstly the gas-phase ignition is considered. A very simple reaction is suggested which is assumed to control the combustion of reactant gas, produced by solid gasification. Ignition is, by definition, the initiation of this exothermic reaction. Chemical kinetics are included in the gas-phase flow equations to explore the evolution of the reactant gas that is subject to changes in temperature and pressure. By assuming spatial uniformity, analytical solutions of the problem are deduced. The physical interpretation of the solution is discussed, in particular, the relationship between temperature, reactant concentration and ignition is explored. Numerical methods are required to solve the one-dimensional flow equations. Development of suitable CFD methods provides a method of solution. Finite-volume schemes, based on the original work by Godunov, are used to solve the conservation form of the equations. A simple test problem is considered whereby reactant gas is injected into a cylindrical combustion chamber. By examining the resulting flow histories, valuable information is gathered about the complicated coupling of chemistry and flow. Chemistry is included into a system of two-phase flow equations. By using standard averaging methods along with an equation for gas-phase species, equations are derived that describe the rate of change of average flo%v variables for both gas and particle phases. Numerical schemes are developed and some of the difficulties involved in two-phase flow systems, that are not an issue in single-phase flow, are presented. An internal ballistics application is considered as a test case and the solution discussed. The other important reaction involved in the combustion cycle, solid gasification, is explored. The model is based on detailed description of interphase mass and energy transfer at the solid-gas interface. This involves the solution of the heat conduction equation with a moving boundary that divides the solid and gas regions. Similar numerical schemes are constructed to solve the equations. Finally, this model is coupled with the equations of gas-phase reaction. This describes the complete cycle whereby increases in gas temperature cause the solid to increase in temperature and gasify. Subsequent gas-phase combustion of the reactant gases produces heat-transfer between the solid and gas and continues to accelerate gasification. Eventually this results in selfsustained combustion of the solid propellant.

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Butler, Albert George. "Holographic investigation of solid propellant combustion". Thesis, Monterey, California. Naval Postgraduate School, 1988. http://hdl.handle.net/10945/23252.

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Approved for public release; distribution is unlimited
An investigation into the behavior of aluminized solid propellant combustion in a two-dimensional windowed rocket motor was conducted using holographic techniques. Holograms were recorded in the motor port, aft of the propellant grain and at the entrance to the exhaust nozzle for two different propellant compositions at varying operating pressures. Quantitative particle size data for particles larger than 20 microns were obtained from the holograms. From these data, the mean diameters (D32) of the larger particles were calculated and utilized to compare what effects pressure, location in the motor and aluminum content had on the behavior of the aluminum/aluminum oxide particles. D 32 was found to decrease with increasing pressure, but was unaffected by variations in low values of propellant aluminum loading. D 32 at the grain exit was found to be significantly less than within the grain port.
http://archive.org/details/holographicinves00butl
Lieutenant, United States Navy

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Cekic, Ayca. "Experimental Study Of Solid Propellant Combustion Instability". Master's thesis, METU, 2005. http://etd.lib.metu.edu.tr/upload/2/12606947/index.pdf.

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In this study, experimental investigation of solid propellant combustion instability using an end burning T-Burner setup is performed. For this purpose, a T-Burner setup is designed, analyzed, constructed and tested with all its sub components. T-Burner setup constructed is mainly composed of a base part, a control panel and the T-Burner itself. Combustion chamber, pressure stabilization mechanism, pressurization system, measurement instruments and data acquisition systems form the T-Burner. Pressure stabilization mechanism is utilized in two different alternatives, first of which is by the use of nitrogen gas and a small surge tank with a cavitating venturi. This is a brand new approach for this kind of system. The second alternative is the use of a choked nozzle for pressure stabilization. Resonance frequencies of the system with the two different pressure stabilization mechanisms are experimentally evaluated. Helmholtz frequency of the T-burner constructed is calculated and no Helmholtz instability is observed in the system. Constructed T-Burner setup is operated for a specific solid propellant. System worked successfully and pressure data are obtained. Pressure data revealed oscillatory behaviour. Decay and growth rates of pressure oscillations are used for the calculation of pressure response of the propellant tested. By the use of this T-Burner comparison of the behavior of different propellants can be performed. It can be used as a test device for measuring quantitatively the response of a burning propellant to unsteady motions.

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Lee, Sung-Taick. "Multidimensional effects in composite propellant combustion". Diss., Georgia Institute of Technology, 1991. http://hdl.handle.net/1853/12111.

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McDonald, Brian Anthony. "The Development of an Erosive Burning Model for Solid Rocket Motors Using Direct Numerical Simulation". Diss., Georgia Institute of Technology, 2004. http://hdl.handle.net/1853/4973.

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A method for developing an erosive burning model for use in solid propellant design-and-analysis interior ballistics codes is described and evaluated. Using Direct Numerical Simulation, the primary mechanisms controlling erosive burning (turbulent heat transfer, and finite rate reactions) have been studied independently through the development of models using finite rate chemistry, and infinite rate chemistry. Both approaches are calibrated to strand burn rate data by modeling the propellant burning in an environment with no cross-flow, and adjusting thermophysical properties until the predicted regression rate matches test data. Subsequent runs are conducted where the cross-flow is increased from M=0.0 up to M=0.8. The resulting relationship of burn rate increase versus Mach Number is used in an interior ballistics analysis to compute the chamber pressure of an existing solid rocket motor. The resulting predictions are compared to static test data. Both the infinite rate model and the finite rate model show good agreement when compared to test data. The propellant considered is an AP/HTPB with an average AP particle size of 37 microns. The finite rate model shows that as the cross-flow increases, near wall vorticity increases due to the lifting of the boundary caused by the side injection of gases from the burning propellant surface. The point of maximum vorticity corresponds to the outer edge of the APd-binder flame. As the cross-flow increases, the APd-binder flame thickness becomes thinner; however, the point of highest reaction rate moves only slightly closer to the propellant surface. As such, the net increase of heat transfer to the propellant surface due to finite rate chemistry affects is small. This leads to the conclusion that augmentation of thermal transport properties and the resulting heat transfer increase due to turbulence dominates over combustion chemistry in the erosive burning problem. This conclusion is advantageous in the development of future models that can be calibrated to heat transfer conditions without the necessity for finite rate chemistry. These results are considered applicable for propellants with small, evenly distributed AP particles where the assumption of premixed APd-binder gases is reasonable.

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Mouille, Hervé. "Influence of strain rate and temperature upon the mechanical and fracture behavior of a simulated solid propellant /". This resource online, 1992. http://scholar.lib.vt.edu/theses/available/etd-07212009-040252/.

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Foss, David T. "Development and modeling of a dual-frequency microwave burn rate measurement system for solid rocket propellant". Thesis, Virginia Tech, 1989. http://hdl.handle.net/10919/45962.

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A dual-frequency microwave bum rate measurement system for solid rocket motors has been developed and is described. The system operates in the X-band (8.2-12.4 Ghz) and uses two independent frequencies operating simultaneously to measure the instantaneous bum rate in a solid rocket motor. Modeling of the two frequency system was performed to determine its effectiveness in limiting errors caused by secondary reflections and errors in the estimates of certain material properties, particularly the microwave wavelength in the propellant. Computer simulations based upon the modeling were performed and are presented. Limited laboratory testing of the system was also conducted to determine its ability perform as modeled.

Simulations showed that the frequency ratio and the initial motor geometry (propellant thickness and combustion chamber diameter) determined the effectiveness of the system in reducing secondary reflections. Results presented show that higher frequency ratios provided better error reduction. Overall, the simulations showed that a dual frequency system can provide up to a 75% reduction in burn rate error over that returned by a single frequency system. The hardware and software for dual frequency measurements was developed and tested, however, further instrumentation work is required to increase the rate at which data is acquired using the methods presented here. The system presents some advantages over the single frequency method but further work needs to be done to realize its full potential.

Master of Science

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McCrorie, J. David. "Particle behavior in solid propellant rocket motors and plumes". Thesis, Monterey, California. Naval Postgraduate School, 1992. http://hdl.handle.net/10945/24002.

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Gomes, Marc Faria. "Internal ballistics simulation of a solid propellant rocket motor". Master's thesis, Universidade da Beira Interior, 2013. http://hdl.handle.net/10400.6/1980.

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In the design and development of solid propellant rocket motors, the use of numerical tools able to simulate, predict and reconstruct the behaviour of a given motor in all its operative conditions is particularly important in order to decrease all the planning and costs. This study is devoted to present an approach to the numerical simulation of a given SPRM internal ballistics, NAWC no. 13, during the quasi steady state by means of a commercial numerical tool, ANSYS FLUENT. The internal ballistics model constructed in this study is a 2-D axisymmetric model, based on several assumptions. Among them is the assumption that there is no contribution of the erosive burning and the dynamic burning in the burning rate model. The results of the internal ballistics simulation are compared with the results found in the bibliographical research, thus validating the model that has been set up. The validation of the results also allows us to conclude that the assumptions made in the construction of the model are reasonable. Suggestions and recommendations for further study are outlined.
Na concepção e desenvolvimento de motores foguete sólidos, o uso de ferramentas numéricas capazes de simular, prever e reconstruir o comportamento de um dado do motor em todas as condições operativas ´e particularmente importante, a fim de diminuir todos os custos e planeamento. Este estudo ´e dedicado a apresentar uma abordagem para a simulação numérica de balística interna de um determinado motor foguete de propelente sólido, Naval Air Warfare Center no. 13, durante a fase quasi steady state por meio de uma ferramenta numérica comercial, ANSYS FLUENT. O modelo de balística interna construído neste estudo é um modelo axissimétrico 2-D. Tem por base vários pressupostos. Entre eles, está o pressuposto de que não há contribuição da queima erosiva e da queima dinâmica no modelo da taxa de queima. Os resultados da simulação balística interna são comparados com os resultados encontrados na pesquisa bibliográfica, validando assim, o modelo que foi construído. A validação dos resultados também nos permite concluir que os pressupostos assumidos na construção do modelo são razoáveis. Sugestões e recomendações para um estudo mais aprofundado são delineadas.

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Più fonti

Libri sul tema "Solid propellant"

Xristin, Schad, e United States. National Aeronautics and Space Administration., a cura di. Propellant variability assessment. [Huntsville, Ala.]: Quality Engineering Research Laboratory, University of Alabama in Huntsville, 1991.

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Kishore, K. Solid propellant chemistry: Condensed phase behaviour of ammonium perchlorate-based solid propellants. New Delhi: Defence Research & Development Organisation, Ministry of Defence, 1999.

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North Atlantic Treaty Organization. Advisory Group for Aerospace Research and Development. Performance of rocket motors with metallized propellants: Report of the Propulsion and Energetics Panel Working Group 17. Neuilly sur Seine, France: AGARD, 1986.

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North Atlantic Treaty Organization. Advisory Group for Aerospace Research and Development. Performance of Rocket Motors with Metallized Propellants: Report of the Propulsion and Energetics Panel : Working Group 17. S.l: s.n, 1986.

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L, Boggs Thomas, Derr Ronald L e Advisory Group for Aerospace Research and Development. Propulsion and Energetics Panel., a cura di. Hazard studies for solid propellant motors. Neuilly sur Seine: Agard, 1990.

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Butler, Albert George. Holographic investigation of solid propellant combustion. Monterey, Calif: Naval Postgraduate School, 1988.

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G, Schirk P., e United States. National Aeronautics and Space Administration., a cura di. Facility design consideration for continuous mix production of class 1.3 propellant. [Washington, DC: National Aeronautics and Space Administration, 1994.

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North Atlantic Treaty Organization. Advisory Group for Aerospace Research and Development. Design methods in solid rocket motors. Neuilly sur Seine, France: AGARD, 1988.

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North Atlantic Treaty Organization. Advisory Group for Aerospace Research and Development. Design Methods in Solid Rocket Motors. S.l: s.n, 1987.

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United States. National Aeronautics and Space Administration., a cura di. NASA's advanced solid rocket motor. [Washington, DC: National Aeronautics and Space Administration, 1993.

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Capitoli di libri sul tema "Solid propellant"

Wu, Jianjun, Jian Li, Yuanzheng Zhao e Yu Zhang. "Numerical Simulation of the Nanosecond Laser Ablation of Al Propellant". In Numerical Simulation of Pulsed Plasma Thruster, 61–87. Singapore: Springer Nature Singapore, 2024. http://dx.doi.org/10.1007/978-981-97-7958-1_4.

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AbstractPulsed plasma thrusters (PPTs) using gaseous propellants (Ziemer and Choueiri in Is the gas-fed PPT an electromagnetic accelerator an investigation using measured performance. AIAA 99–2289, 1999; Ziemer et al. in Performance characterization of a high efficiency gas-fed pulsed plasma thruster. AIAA-97–2925, 1997; Ziemer in Performance scaling of gas-fed pulsed plasma thrusters. Princeton University, 2001; Ziemer and Petr in Performance of gas fed pulsed plasma thrusters using water vapour propellant. AIAA 2002–4273, 2002) typically far outperform those using solid propellants in terms of parameters such as specific impulse and propulsion efficiency.(Porneala and Willis in J Phys D Appl Phys 42:1–7, 2009) Therefore, during the PPT operation, first, the propellant is transformed from a solid state to a gaseous or plasma state to ensure that what is actually ionized in the discharge channel or discharge chamber of the thruster is not the solid propellant but rather the gaseous or plasma propellant. In general, the solid propellant cannot be completely converted to a gaseous or plasma state during the laser ablation process. To increase the gas and plasma components in the discharge channel, an intense laser with a nanosecond pulse width is used as the energy source for propellant ablation. This approach is important for improving the PPT propulsion performance.

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Greatrix, David R. "Solid-Propellant Rocket Motors". In Powered Flight, 323–79. London: Springer London, 2012. http://dx.doi.org/10.1007/978-1-4471-2485-6_10.

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Mishra, D. P. "Solid-Propellant Rocket Engines". In Fundamentals of Rocket Propulsion, 195–259. Boca Raton: CRC Press, 2017.: CRC Press, 2017. http://dx.doi.org/10.1201/9781315175997-7.

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Ranjan, Rajeev, e H. Murthy. "Compressive Behaviour of Composite Solid Propellant". In Advances in Applied Mechanics, 25–30. Singapore: Springer Nature Singapore, 2024. http://dx.doi.org/10.1007/978-981-97-0472-9_4.

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Cheng, S. I. "L*-Combustion Instability in Solid Propellant Rocket Combustion". In Recent Advances in the Aerospace Sciences, 257–78. Boston, MA: Springer US, 1985. http://dx.doi.org/10.1007/978-1-4684-4298-4_13.

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Akbar, Mohammed, e Prabhat Dattakumar Phondekar. "Design and Analysis of Optimized Solid Propellant Grain". In Lecture Notes in Mechanical Engineering, 743–58. Singapore: Springer Nature Singapore, 2024. http://dx.doi.org/10.1007/978-981-99-7827-4_58.

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Nagappa, Rajaram. "The First Steps Toward Self-reliance in Solid Propellant Rockets". In The Mind of an Engineer, 401–8. Singapore: Springer Singapore, 2015. http://dx.doi.org/10.1007/978-981-10-0119-2_51.

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Traissac, Y., J. Ninous, R. Neviere e J. Pouyet. "Mechanical Behavior of a Solid Composite Propellant During Motor Ignition". In Advances in Chemistry, 195–210. Washington, DC: American Chemical Society, 1996. http://dx.doi.org/10.1021/ba-1996-0252.ch014.

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Hawthorne, M. Frederick. "Moving on to New Concepts for Solid Propellant Rocket Fuel". In Boranes and Beyond, 179–80. New York, NY: Springer New York, 2023. http://dx.doi.org/10.1007/978-1-0716-2908-6_25.

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Zhang, Xue-Xue, Hao-Rui Zhang, Ming-Hui Yu e Qi-Long Yan. "Metal-Based Green Energetic Catalysts for Solid Propellant Combustion Control". In Space Technology Library, 283–332. Cham: Springer Nature Switzerland, 2024. http://dx.doi.org/10.1007/978-3-031-62574-9_10.

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Atti di convegni sul tema "Solid propellant"

Chen, Yang, Vahid Morovati e Roozbeh Dargazany. "A Directional Damage Constitutive Model for Stress-Softening in Solid Propellant". In ASME 2020 International Mechanical Engineering Congress and Exposition. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/imece2020-24285.

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Abstract Solid propellants are particulate composite with a light cross-linked elastomeric binder filled with a high concentration of energetic, solid aggregates. Solid propellants are often considered as highly nonlinear elastomeric materials, with elastic behavior resulted from its binder and plastic behavior from its energetic particles. The study of the micro-structure and mechanical properties of solid propellant is crucial for its design, safety evaluation, and lifetime prediction of solid fuel carriers. The constitutive model proposed for rubber-like material can often be generalized to predict the nonlinear behavior of solid propellant due to the dependency on the mechanical behavior of solid propellant on its elastomeric binder material. This paper focuses on developing a model that predicts the stress softening and strain-residual mechanism of the solid propellant. This micro-mechanical model for solid propellant was proposed based on the network evolution theory. The motivation of this study is the lack of a micro-mechanical model that can describe both the stress softening effect and strain residual in the quasi-static behavior of propellants. The simplified network-evolution model with only five parameters is a simple micro-mechanical model that captures both the stress softening effect and strain residual. Besides the simplicity and reduced fitting procedure, the model was validated against several experimental data and illustrated good agreement in small and large deformations, making the proposed model a suitable option for commercial and other applications.

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Jang, Jin-Sung, Hyung-Gun Sung, Seung-Young Yoo, Tae-Seong Roh e Dong-Whan Choi. "Numerical Study on Properties of Interior Ballistics According to Solid Propellant Position in Chamber". In ASME-JSME-KSME 2011 Joint Fluids Engineering Conference. ASMEDC, 2011. http://dx.doi.org/10.1115/ajk2011-12005.

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Analysis of the interior ballistics is essential for the development of gun or propellant configurations. The granular solid propellants with high energy and fast burning rate produce a large thrust in extremely short time intervals. For the study of these, therefore, it is necessary of a numerical code for the two-phase flow of the interior ballistics. Recently, an interior ballistics code (IBcode) for the two-phase flow using the Eulerian-Lagrangian approach has been developed. The SIMPLE algorithm and the SMART scheme have been used for the IBcode. The ghost-cell extrapolation method has been used for the moving boundary with the projectile movement. In this study, a performance of the interior ballistics according to the position of the solid propellant in the chamber has been investigated using the IBcode. In previous researches, propellants had been evenly distributed in the chamber. In this study, however, three cases of the existence of empty space in the chamber at which the propellants are not evenly distributed have been considered; Propellants are located in the region near the base, propellants in the region near the breech, and propellants in the center of the chamber, respectively. The 7-perforated configuration of the solid propellant has been used in this research. The results have shown the performance variations of the interior ballistics according to solid propellant position in the chamber. The cases of the propellants located in the region near the base and breech have shown that the value of the negative differential pressure and the difference between the breech pressure and the base pressure are much higher than those of the propellants located in the center of the chamber. The case of the propellants in the center of the chamber is, therefore, more profitable to improve the performance of the interior ballistics.

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Vanderhoff, John A. "Multichannel absorption spectroscopy applied to solid-propellant flames". In OSA Annual Meeting. Washington, D.C.: Optica Publishing Group, 1992. http://dx.doi.org/10.1364/oam.1992.tugg5.

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Absorption spectroscopy has been used to characterize the steady-state flame structure of solid propellants burning over a pressure range of 0.3 to 2.0 MPa. A high-intensity-xenon-arc-lamp light source and two spectrometer-intensified photodiode-array detectors formed the experimental setup from which NO and OH absorption spectra were acquired. Vibrationally resolved transitions in the A2Σ - X2Π (0,1) electronic system of NO from 230 to 250 nm comprise the absorption spectra from which temperatures and absolute NO concentrations are determined in the dark-zone region of the solid-propellant flames, and rotationally resolved transitions in the A2 Σ - X2 Π (0,0) vibrational band system of OH from306to311nm comprise the absorption spectra from which luminous-flame temperatures are determined. These spectra can be least-squares fitted with respect to a variety of parameters, which include an instrument response function, an absorption baseline, as well as the temperature and concentration. Several different propellants that exhibit a dark zone have been studied as a function of pressure. The dark-zone temperatures ranged from 1300 K to 1500 K, and NO concentrations varied from 13 to 30 mole percent, depending on the propellant type and pressure. Temperatures in the luminous flame region reached adiabatic values.

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Marshall, Tony, John Evans e Robert Frederick. "UAH Solid Propellant Characterization". In 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2007. http://dx.doi.org/10.2514/6.2007-5763.

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Wingborg, Niklas. "Solid ADN Propellant Development". In 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2013. http://dx.doi.org/10.2514/6.2013-3723.

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GODON, J., J. DUTERQUE e G. LENGELLE. "Solid propellant erosive burning". In 23rd Joint Propulsion Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1987. http://dx.doi.org/10.2514/6.1987-2031.

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Biggs, Gary. "Solid Propellant Aging Kinetics". In 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2009. http://dx.doi.org/10.2514/6.2009-5423.

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Parra, E. A., K. S. J. Pister e C. Fernandez-Pello. "A Practical Solid-Propellant Micro-Thruster". In ASME 2006 International Mechanical Engineering Congress and Exposition. ASMEDC, 2006. http://dx.doi.org/10.1115/imece2006-15061.

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Miniaturization of solid-propellant thrusters is an area of active research that has been motivated by the reduction in size of aerospace systems and the advancement of micromachining techniques. Though this micro-propulsion problem seems simplistic compared to the macro-scale counterpart, an efficient and reliable device has yet to be produced. A millimeter-scale novel composite solid-propellant thruster design that builds on pervious work [1] and increases efficiency is here presented. Current designs made primarily out of silicon suffer from high thermal losses and, in extreme cases, flame quenching due to the augmented surface area to volume ratio associated with miniaturization. Moreover, the reduced device dimensions drive the combustion reaction to complete outside of the thruster, misemploying the majority of the chemical energy. This occurs because the propellant mixing and chemical time do not scale with size, while the residence time does decrease as the size of the thruster decreases [2]. A novel thruster design that increases the propellant residence time is being characterized using ammonium perchlorate/binder composite propellant. The thruster geometry recycles thermal energy to the unburned propellant grain increasing its temperature and, therefore, burning rate and combustion efficiency. In addition, propellant formulation has been optimized for the thruster minimization.

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Price, E., R. Jeenu, S. Chakravarthy e J. Seitzman. "Solid propellant combustion - Surface disproportionation". In 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2000. http://dx.doi.org/10.2514/6.2000-3327.

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BLOMSHIELD, F., e J. OSBORN. "Nitramine composite solid propellant modeling". In 26th Joint Propulsion Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1990. http://dx.doi.org/10.2514/6.1990-2311.

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Rapporti di organizzazioni sul tema "Solid propellant"

Yang, Jiann C., e William L. Grosshandler. Solid propellant gas generators:. Gaithersburg, MD: National Institute of Standards and Technology, 1995. http://dx.doi.org/10.6028/nist.ir.5766.

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Fry, Ronald S. Solid Propellant Test Motor Scaling. Fort Belvoir, VA: Defense Technical Information Center, settembre 2001. http://dx.doi.org/10.21236/ada386366.

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Blomshield, F. S. Nitramine Composite Solid Propellant Modelling. Fort Belvoir, VA: Defense Technical Information Center, luglio 1989. http://dx.doi.org/10.21236/ada220198.

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Price, E. W., e G. A. Flandro. Combustion Instability in Solid Propellant Rockets. Fort Belvoir, VA: Defense Technical Information Center, febbraio 1987. http://dx.doi.org/10.21236/ada179701.

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Graves, V., G. Bader, M. Dolecki, S. Krupski e R. Zangrando. Crusader solid propellant best technical approach. Office of Scientific and Technical Information (OSTI), dicembre 1995. http://dx.doi.org/10.2172/179267.

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Baron, D. T., C. T. Liu e T. C. Miller. Subcritical Crack Growth in a Composite Solid Propellant. Fort Belvoir, VA: Defense Technical Information Center, maggio 1998. http://dx.doi.org/10.21236/ada409841.

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Stolovy, A., A. I. Namenson e J. M. Kidd. Solid Rocket Propellant Initiation Via Particle Beam Heating. Fort Belvoir, VA: Defense Technical Information Center, maggio 1990. http://dx.doi.org/10.21236/ada221900.

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Fry, R. S., L. DeLuca, R. Frederick, G. Gadiot, R. Strecker, H.-L. Besser, A. Whitehouse, J.-C. Traineau, D. Ribereau e J.-P. Reynaud. Evaluation of Methods for Solid Propellant Burning Rate Measurement. Fort Belvoir, VA: Defense Technical Information Center, gennaio 2002. http://dx.doi.org/10.21236/ada405711.

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Liu, C. T., Y. W. Kwon e T. L. Hendrickson. Predicting the Initial Crack Length in a Solid Propellant. Fort Belvoir, VA: Defense Technical Information Center, ottobre 2000. http://dx.doi.org/10.21236/ada408146.

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Liu, C. T., Y. G. Kwon e T. L. Hendrickson. Predicting the Initial Crack Length in a Solid Propellant. Fort Belvoir, VA: Defense Technical Information Center, gennaio 2001. http://dx.doi.org/10.21236/ada410143.

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