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Energetic materials are a kind of materials involving a huge amount of chemical energy and they can release this energy suddenly or slower. In this review study, types, usage areas and compositions of energetic materials are explored. Especially, energetic materials for solid propellant rocket engines are evaluated more extensively. The advantages and synthesis techniques of metal borides as an alternative to fuel in solid propellant rocket engines are criticized in the light of the relevant literature. It is concluded that, metal borides, particularly aluminum and magnesium borides, are promising materials as solid rocket engine fuel. However, new methods should be developed for the synthesis of higher amount and purity of metal boride. [ABSTRACT FROM AUTHOR]

Abstract Solid propellant burning rate, gas phase temperature, and condensed phase thickness depend on combustion chamber pressure, laser intensity, and propellant compositions. The ignition and combustion of solid propellants occur in three phases namely solid phase, condensed phase, and gas phase. In this study, moving boundary modeling was applied to each of the phases by coordinate transformation. This research includes modeling and dynamic simulation of the ignition and combustion of HMX/GAP, a high-energy material in the ratio of 8:2, with the gas phase of the combustion model consisting of 50 species and 234 reactions. The mathematical modeling used mass, energy, and momentum conservation equations, as well as constitutive equations for the moving boundary. Parametric studies were run under different operating conditions, with an initial temperature of 300 K, a pressure of 10–100 atm and a laser intensity of 100 W/cm2. A burning rate of 2.2 cm/s and a gas phase temperature of 2700 K were obtained under an operating pressure of 100 atm. Extinguishment of the solid propellant was rigorously analyzed in terms of dynamic simulation with various depressurization rates during combustion. This was carried out by depressurizing the solid propellant from 70 atm to 40 atm. Important factors of the extinguishment were discussed based on the mathematical model and various depressurization rates with parametric studies. At a depressurization rate of −8000 atm/s, the solid propellant was fully extinguished. From this study, one can identify the phenomenon for the extinguishment of HMX/GAP propellant using fast depressurization with rigorous mathematical model used for ignition and combustion. [ABSTRACT FROM AUTHOR]

In order to make a vacuum pouring cylinder cover have an improved function, where two large solid rocket engines rather than one can be poured through simultaneously in a vacuum environment, the structure of the original vacuum pouring cylinder cover should be reformed. In addition to the original one, two other pouring ports are required. Therefore, to ensure its strength and stiffness properties, some stiffeners (or stiffen ribs) should be welded on its inner surface. Based on the response surface method and finite element analysis method, a multi-objective optimization model is obtained to optimize the structure of the two-engine pouring cylinder cover. Three sizes related to the stiffeners are defined as design variables, and the structural stresses and deformation values are calculated by using ANSYS software. Box-Behnken design experiments are conducted after narrowing the range of design variables by using a combination test. Design-Expert, a kind of common experiment design software, is used to optimize and solve the response surface. The optimal values of sizes of stiffeners for the two-engine pouring cylinder cover structure are obtained. The research results build a solid foundation for the practical application of the project. [ABSTRACT FROM AUTHOR]

The article, which is of a review nature, discusses the problems of designing and calculating the hulls of rocket engines of solid propellant with a high degree of weight perfection and made of composite materials by the method of automatic continuous winding. The history of the creation of such structures is briefly described and the problems of optimal reinforcement and optimization of the structural forms of composite membrane-free shells of revolution are considered. The results obtained in this direction by the scientific school of academician V. V. Vasiliev are discussed. [ABSTRACT FROM AUTHOR]

Powdered Mg and CO 2 bipropellant engine providing a practical demonstration of in situ resource utilization (ISRU) for Mars Sample Return (MSR) mission seems to be feasible by current investigations. However, essential functions of the engine to satisfy the complicated ballistics requirements such as thrust modulation and multiple pulse have not been established yet. The aim of this experimental study is to evaluate the engine's thrust modulation feasibility and to investigate its thrust modulation characteristics. A powdered Mg and CO 2 bipropellant engine construction aiming to achieve thrust modulation ability was proposed. A mass flow rate calibration experiment to evaluate the gas-solid mass flow rate regulating performance was conducted before fire tests. Fire test result shows that the engine achieved successful ignition as well as self-sustaining combustion; Thrust modulation of the engine is feasible, detail thrust estimating result of the test shows that maximum thrust is 135.91 N and the minimum is 5.65 N with a 22.11 thrust modulation ratio, moreover, the transportation period is quick and the thrust modulation ratio is adjustable. At the same time, the powder feed system reaches a two-step flow rate regulating with a modulation ratio of 4.5–5. What’ more, caused by the uneven engine working conditions, there is an obvious difference in combustion efficiency value, maximum combustion efficiency of the powdered Mg and CO 2 bipropellant engine is 80.20%. [ABSTRACT FROM AUTHOR]

Mathematical models that simulate the internal ballistics of solid propellant rocket motors are widely used in lieu of the expensive, hazardous, time-consuming static firing experiments. Since they rely on the input of various measured data, these models are vulnerable to uncertainties that may deteriorate their prediction accuracy. Improving the accuracy of internal ballistics prediction can be achieved by coupling the mathematical models with optimization algorithm. This paper discusses the issues of uncertainties and their impact on prediction accuracy of developed mathematical models. The model handles two types of dual thrust rocket motors, where the impact of uncertainties is more pronounced since they include more geometric and ballistic parameters. The model is developed based on the fundamental principles of internal ballistics and is coupled with a commercial genetic algorithm optimization tool. Two static firing tests are conducted to assess the proposed optimized prediction model. It is found that prediction optimization via tuning the uncertain parameters has led to significantly improving the prediction accuracy. More importantly, the tuned uncertain parameters give better understanding and clearer insight of the phenomena taking place inside dual thrust solid propellant rocket motors. [ABSTRACT FROM AUTHOR]

The velocity of spherical gas-dust cloud expansion in the situation when the stages of solid-propellant rocket separate in the upper atmosphere have been determined. The measured velocity vary from 2.5 to 7.5 km/s. The dispersed component accelerates at the front of a shock that develops at engine-thrust shutdown. The model calculations of the gas-dust cloud luminosity intensity qualitatively coincide with the photometric profiles of object images. Such formations can vary from almost homogeneous ball-shaped clouds to rather thin spherical shells depending on the gas-dust cloud mass and the matter distribution within this cloud. [ABSTRACT FROM AUTHOR]

In composite solid propellants, shortening particle residence time at the burning surface and inducing particle microexplosions could decrease aluminum agglomeration, thus reducing two-phase flow losses in a rocket motor. We explore this by using aluminum particles modified with low-density polyethylene (LDPE) inclusion to drive intraparticle outgassing, which could break apart composite particles, yielding smaller and faster burning fragments during composite solid propellant combustion. We find that use of these particles in a propellant results in more prompt particle ignition, and surface residence time is decreased. For composite propellant burning at 6.9 MPa, mean coarse agglomerate diameter is decreased from 75.8 μm (spherical aluminum) to 29.0 μm (Al/LDPE 90/10 wt.% particles). Thermal analysis with DSC/TGA shows that 10 wt.% LDPE inclusion in aluminum (1.5% of propellant weight) results in enhanced oxidation characteristics that are more similar to nanoaluminum than neat spherical aluminum. Thermochemical equilibrium calculations indicate LDPE inclusion decreases specific impulse by 1.0% from 262.7 to 260.0 s, but it is expected that in a motor, LDPE inclusion could produce a net increase in specific impulse due to a substantially reduced agglomerate size. This work shows that reduced agglomeration is possible using gas generating inclusion materials that are only weakly reactive with aluminum. [ABSTRACT FROM AUTHOR]

Purpose -- This study aims to present a method for conceptual design and simulation of hybrid propellant motors. Design/Methodology/Approach -- The design methodology is based on previous studies and thermo-physical governor relations, and also a computational code that was derived to simulate the performance of designed system. Findings -- Conceptual design algorithm for space hybrid propellant systems and method of performance simulation are findings of this study. Practical implications -- Results of this study are applicable for design process of space systems with hybrid propulsion. Also, simulation results can help the users to improve the performance weaknesses. Originality/value -- This study shows the implementation of present algorithm for a specified space mission and also study on variation of performance parameters. [ABSTRACT FROM AUTHOR]

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This article is devoted to a new block-module device used in the diagnostics of condensed combustion products of rocket engines during research and development with liquid-propellant rocket engines (Glushko NPO Energomash; engines RD-171, RD-180, and RD-191) and solid-propellant rocket motors. Soot samplings from the supersonic high-temperature jet of a high-power liquid-propellant rocket engine were taken by the given device for the first time in practice for closed-exhaust lines. A large quantity of significant results was also obtained during a combustion investigation of solid propellants within solid-propellant rocket motors. [ABSTRACT FROM AUTHOR]

Solid propellant used for Multi Barrel Rocket Launcher (MBRL) application is a composite propellant consists of energetic metallic fuel Aluminum (Al), oxidizer Ammonium Perchlorate (AP) and polymeric binder Hydroxyl Terminated Polybutadiene (HTPB). Solid propellant is preferred because it is convenient and simple for use by the troops. In addition to high specific impulse capability and quick response time, the solid propellant retains its usefulness even on prolonged storage and deployment in field condition. Good mechanical strength makes transportation and handling is easy. However, due to presence of AP in the propellant, toxic HCl plume is generated as one of the combustion products. This may not cause acute problem as field deployment of the weapon system is in open area. Tactical requirement of MBRL rockets demand little time of flight to reach the target and balastically the propellant needs to generate constant thrust. This can be achieved if the grain geometry is designed for constant pressure burning. For the design of a solid propellant the aim is to obtain maximum burning area, since it is this area that determines the maximum chamber pressure that the rocket would generate. Moreover, internal surface of the propellant must be uniform throughout the entire length of the grain to obtain constant thrust. Since neutral burning surface produces constant pressure time curve, the same is required for the MBRL propellant. Form function relation and surface area relation are the grain design parameters responsible for type of burning namely progressive, regressive or neutral. It is seen that star shape offers neutral burning, however leaves sliver at the end of effective burning. Based upon the requirement, manufacturing of solid propellant is done either by casting or by extrusion process The purpose of the paper is to analyze as to why solid propellant is preferred in rockets to propel through MBRL in military application and the implication of grain geometry The paper also carries out mathematical modeling of star shape propellant grain analyses web burn with respect to pressure developed and time of burning. It is observed that star shape composite propellant is suitable for MBRL application. It can be seen in the plot of web burnt - pressure curve that the pressure is constant initially due to pure neutral burning, after that it dips a bit and some instability is noticed during transient phase. It is then linearly increased due to progressive burning up to 41 mm web burning during equilibrium phase. [ABSTRACT FROM AUTHOR]

The flow of an inviscid liquid containing condensed-phase particles around the recessed nozzle of the solidpropellant rocket motor has been considered. To take into account the complex geometry of the prenozzle space of the rocket motor, the equations describing the liquid flow were written in a curvilinear coordinate system. The paths of solid particles were calculated in the known liquid flow field. The quality criteria of the meshes constructed with the use of different approaches have been compared. The results of the calculations permit determining the limit path of particles dividing the flow region into two subregions, one of which is occupied by particles and the other is free from condensed-phase particles. [ABSTRACT FROM AUTHOR]

A probabilistic storage life prediction model of solid rocket motor (SRM) grain is proposed based on the viscoelastic stochastic finite element method (SFEM) and the temperature accelerated aging test of solid propellant. The relationship between the digital characteristics of solid propellant property parameters of solid propellant and storage time is derived from the statistical analysis of aging test data, and the mean and standard deviation of grain structural response are calculated using 3 D viscoelastic response surface SFEM. The structural reliability of a certain SRM grain at the different storage times is analyzed, and its probabilistic storage life is estimated. The methods are available for the designers and users of SRM. [ABSTRACT FROM AUTHOR]

Thermophysical characteristics of phenolic carbon-filled plastic P-5-13P and graphite V-1 have been investigated on an experimental setup simulating an actual solid-propellant rocket engine. Temperature and heat-flux density on the thermally stressed boundaries of samples of the materials under study have been determined by solution, with a regularizing algorithm, of the boundary inverse problem. [ABSTRACT FROM AUTHOR]

The article presents a study which investigates the gas-dynamic instability of working process in the combustion chamber using solid-propellant rocket engines (SPREs). The study investigated the direct numerical simulation of the low frequency acoustic instability in the two-chamber SPRE by multiparameter optimization, along with the nonlinear problems of continuum mechanics. Results of the study indicate that the combustion chamber 2 operates in an established mode without pressure pulsations.

Abstract: Material and definition of dome-shaped ceramic port cover contour is presented. Two types of ceramic port covers with equal thickness and variable thickness structures are proposed. A group of cold-flow tests was carried out on a reusable dual pulse motor in which the performances of different port covers including sealing, first-pulse pressure bearing and second-pulse broken pressure were experimentally studied. A solid dual pulse motor for hot-firing test was designed, manufactured and statically tested. The feasibility of proposed bulkhead with ceramic port cover employed on actual solid dual pulse motor was proved. It is found that dome-shaped port cover made from machinable glass ceramic withstands a high pressure on its convex side, but is prone to fragmentation while imposing a reasonably low pressure on its concave side. The peripherally thickened port cover that is more robust in machining and assembly process is recommended. Grooves milled on the port cover induce fragmentation and are helpful to produce regular broken aperture shape. The bulkhead assembly consists of retaining ring, port cover and supporting ring can well fulfill the requirements of port cover in fixing, preloading and sealing. The broken fragments are less likely to damage the insulation structures of the combustion chamber and the nozzle unit which had been validated in the test. [Copyright &y& Elsevier]

Comprehensive numerical modeling of the processes occurring in the combustion chamber of a solid propellant rocket engine during the stabilization of the design operation mode is performed. The self-consistent problem considered includes nonstationary operation of an ignition device, warmup and ignition of a solid propellant charge followed by its nonstationary burning, nonstationary threephase homogeneous-heterogeneous flow of the combustion products in the combustion chamber, in the nozzle, and behind the engine nozzle unit, engine depressurization, and nozzle unit plug blowing-out. The results of the calculations are presented. [ABSTRACT FROM AUTHOR]

The ability to predict with some accuracy a given solid rocket motor's performance before undertaking one or several costly experimental test firings is important. On the numerical prediction side, as various component models evolve, their incorporation into an overall internal ballistics simulation program allows for new motor firing simulations to take place, which in turn allows for updated comparisons to experimental firing data. In the present investigation, utilizing an updated simulation program, the focus is on quasi-steady performance analysis and scale effects (influence of motor size). The predicted effects of negative/positive erosive burning and propellant/casing deflection, as tied to motor size, on a reference cylindrical-grain motor's internal ballistics, are included in this evaluation. Propellant deflection has only a minor influence on the reference motor's internal ballistics, regardless of motor size. Erosive burning, on the other hand, is distinctly affected by motor scale. [ABSTRACT FROM AUTHOR]

Pulse phase dynamic thermal tomography (PP-DTT) was introduced as a nondestructive inspection technique to detect the defects of the solid-propellant missile engine cladding layer. One-dimensional thermal wave mathematical model stimulated by pulse signal was developed and employed to investigate the thermal wave transmission characteristics. The pulse phase algorithm was used to extract the thermal wave characteristic of thermal radiation. Depth calibration curve was obtained by fuzzy c-means algorithm. Moreover, PP-DTT, a depth-resolved photothermal imaging modality, was employed to enable three-dimensional (3D) visualization of cladding layer defects. The comparison experiment between PP-DTT and classical dynamic thermal tomography was investigated. The results showed that PP-DTT can reconstruct the 3D topography of defects in a high quality. [ABSTRACT FROM AUTHOR]