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Nuclear rockets operate using one of the several types of thermodynamic cycles which vary in complexity and efficiency. For nuclear thermal rockets, these thermodynamic cycles are “open,” in that during operation, the working fluid is discharged through the nozzle to produce thrust after circulating only once through the engine system. These engines typically use a turbopump to highly pressurize the propellant prior to being introduced into the reactor where the propellant is heated to high temperatures before being discharged to the nozzle. The pump is normally driven by an integrated turbine system which is powered by propellant that has been warmed somewhat using waste heat from the reactor. Nuclear systems, which are designed to produce electricity for some type of electric or ion rockets, generally incorporate “closed” thermodynamic cycles in which the working fluid always remains within the system and is circulated continuously during operation. Nuclear electric systems require, in addition to the turbopump assembly, radiators to reject the waste heat that results from the thermal-to-electric conversion process.

Due to the harsh environment present during nuclear engine operation, the materials used for the fuel, moderator, and other components must be quite resilient in order to survive these severe conditions for any extended length of time. Temperatures within the fuel, for example, may reach 3000 K or even higher. Besides the extremely high temperatures, the corrosive effects of hydrogen or other propellant gases along with the extremely high-radiation environment must be considered during the materials selection process. As a result of these factors, the number of options available with regard to the fabrication of the fuel elements is, not surprisingly, quite limited. The moderator and other structural components must also survive a similarly harsh environment, although, generally not quite as severe as that encountered by the fuel. Since these are material limitations which constrain the ultimate performance of the nuclear engine, considerable effort will be required in the future if the operational envelope over which the nuclear engine can function is to be significantly expanded over that which is currently envisioned.

The internal flow structures of detonation wave were experimentally analyzed in an optically accessible hollow rotating detonation combustor with multiple chamber lengths. The cylindrical RDC has a glass chamber wall, 20 mm in diameter, which allowed us to capture the combustion self-luminescence. A chamber 70 mm in length was first tested using C2H4single bondO2 and H2–O2 as propellants. Images with a strong self-luminescence region near the bottom were obtained, confirming the small extent of the region where most of the heat release occurs as found in our previous research. Based on the visualization experiments, we tested RDCs with shorter chamber walls of 40 and 20 mm. The detonation wave was also observed in the shorter chambers, and its velocity was not affected by the difference in chamber length. Thrust performance was also maintained compared to the longer chamber, and the short cylindrical RDC had the same specific impulse tendency as the cylindrical (hollow) or annular 70-mm chamber RDC. Finally, we calculated the pressure distributions of various chamber lengths, and found they were also consistent with the measured pressure at the bottom and exit. We concluded that the short-chamber cylindrical RDC with equal length and diameter maintained thrust performance similar to the longer annular RDC, further expanding the potential of compact RDCs., Available online 17 August 2020

Technological miniaturization has enabled the development of small satellites weighing as little as 1 kg. Unfortunately, there is still a lack of suitable efficient micropropulsion systems at these scales. The pulsed plasma thruster is a structurally simple form of electric propulsion. This simplicity also makes it ideally suited for miniaturization. Its history can be traced back to applications in satellites that are much larger than micro/nano-satellites. The vast majority of modern pulsed plasma thrusters use solid polytetrafluoroethylene (PTFE) as a propellant. Unfortunately, at lower discharge energy levels such as those necessitated by the power limitations of micro/nano-satellites, PTFE has a tendency to exhibit carbon deposition, which can ultimately lead to thruster failure. In this new era of small satellites, it is important to consider alternative propellants in the miniaturization of pulsed plasma thrusters. This brief review discusses the needs and limitations of small satellites and alternative propellants that may be able to meet these needs. Such propellants may be able to offer advantages such as a longer thruster lifetime, a higher specific impulse, or a higher thrust-to-power ratio. This would enable the development of different types of pulsed plasma thrusters that can be tailored towards specific mission requirements.

In order to improve the energy performance and mechanical properties of insensitive high-energy propellants, cyclotrimethylenetrinitramine (RDX)-based insensitive high-energy propellant with nitrocellulose (NC) and glycidyl azide polymer-energetic thermoplastic elastomer (GAP-ETPE) blends as the binder was prepared by solvent method. The effects of GAP-ETPE/NC ratio and RDX content on the energy properties of propellants were analyzed by numerical calculations. The effects of GAP-ETPE/NC ratio, RDX particle size and RDX content on the mechanical properties of propellants were studied by compression test, impact test and solid surface properties test. The impact sensitivity and friction sensitivity of all the propellant compositions were tested. The theoretical calculation results showed that the introduction of GAP-ETPE into propellant significantly decreased the explosion temperature and increased the specific volume. The density, compression strength, and surface energy of the propellant were gradually decreased with increasing the GAP-ETPE/NC ratio, and the impact strength was increased gradually as well. The mechanical properties and surface energy of the propellant were improved greatly when the RDX particles were smaller. The studied propellants showed a linear increase in the propellant force and explosion temperature with increasing the RDX content. However, increasing the RDX content significantly reduced the impact strength and surface energy. Mechanical reinforcement with solid fillers increased the compression strength and propellant density to a certain degree. Compared with the standard propellants SF-3, the prepared propellant composition is insensitive to the impact and friction.

The Pulsed Inductive Thruster (PIT) has the advantages of repeatable startup, no corruption and in-situ propellant feed. To study the flow expansion and circuit characteristics of PITs, the circuit-fluid model is developed, and the high temperature thermodynamic and transport models are combined with the circuit-fluid model to predict the critical plasma parameters. The flow fields of initial mass of 2–8 mg and charge voltages of 10–14 kV are simulated. Comparison of the flow fields of argon and helium propellants suggests that, the flow field structures are similar. Slight differences exist on the magnitude of the density and magnetic field, caused by larger velocity in lighter atom case and difference on the ionization gap between adjacent ionization levels. Analysis of the circuit characteristics by the two-dimensional results indicates that the ratio of coil inductance to circuit inductance affects both the rise rate and phase of the plasma current, the larger the ratio, the greater the rise rate and the better the following characteristic. The calculations show that the magnetic energy obtained within the decoupling distance determines the overall performance the thruster can be obtained; self-induced field maintained by the thermal motion after the main pulse leads to the long attenuation process and difference on the total impulse when the angle of conical pylon is varied under constant coil dimension.

An investigation was performed to determine the pressure transient levels and time required to reach pressure equilibrium in a regulated liquid propulsion system due to the priming of a propellant tank during system initiation. An experimental fluid flow test setup was designed and fabricated using a pressure vessel to house distilled water as a propellant simulant for liquid hydrazine and high-pressure gaseous nitrogen as the pressurant. A snubber orifice was installed downstream of the high-pressure gaseous nitrogen bottles to protect a regulator from high-pressure transients, prior to the regulation of incoming gas to beginning of life propulsion system propellant tank conditions. Experiments were conducted to measure pressure transients and time to pressure equilibrium by varying initial system pressures, snubber orifice diameter, and system ullage volume. A mathematical model was developed and compared against the experimental regulated system priming results to predict the system pressures and time to reach system pressure equilibrium during the pressurization period for regulated propulsion systems given initial conditions.

This paper presents the combustion characteristics in hybrid rocket motors with multi-segmented grain through three-dimensional numerical simulations. Multi-segmented grain is composed of several thin grains with two or more ports. The numerical model consists of Navier-Stokes equations with turbulence, solid fuel pyrolysis, chemical reactions, a fluid–solid coupling model and a regression rate model. The simulations adopt 90% Hydrogen Peroxide (HP) and PolyEthylene (PE) as the propellant combination. The effects of the rotation, port number, fuel grain segment number and mid-chamber length on the flow field and combustion performances are analyzed. The results indicate that the multi-segmented grain configuration can strengthen the flow field, and the regression rate and combustion efficiency are enhanced. Take the cases with two grain segments and three ports for example, the regression rate is increased by 32.4%−45.1% and the combustion efficiency increases by 6%−8.6% in different rotation angles.

While increased pressure and temperature contribute to an overall efficiency gain in the mixing of propellants and oxidizers, characteristic of conditions in the combustion chambers of liquid rocket engines, they also propel mixtures to trans- and supercritical conditions. In these conditions, the engine flow exhibits a gas jet-like behavior that may be described using an approach developed for variable density incompressible flows. The present study focuses on an approach using the Reynolds-averaged Navier–Stokes equations to evaluate the jet topology for different injectors’ conditions. Based on the so-called ’thermal breakup mechanism concept’ proposed in the literature, the axial density decay in supercritical nitrogen jets is predicted for a wide range of conditions. The results show the influence of thermal breakup, providing a better insight of the available experimental data., Fundação para a Ciência e a Tecnologia

The modeling of the deformation and failure process of a cyclotetramethylene tetranitramine modified double-base (HMX-MDB) propellant based on the microstructure is of great importance for its manufacturing process and storage lifetime. In the present work, the debonding, nucleation, and crack propagation in an HMX-MDB propellant were investigated experimentally and numerically. Stress-strain responses and viscoelastic parameters were obtained through uniaxial tensile tests on the HMX-MDB propellant and stress relaxation tests on a nitrocellulose-nitroglycerine (NC-NG) propellant, respectively. A two-dimensional representative volume element (RVE) model was developed based on the molecular dynamics method by embedding HMX particles in the NC-NG matrix. A zero-thickness cohesive element was used to simulate the damage behavior of the matrix and the interface. A piecewise cohesive zone model (PCZM) was established to capture stress-strain responses resulted from the matrix rupture and the debonding process at the particle-matrix interface. The effects of matrix strength and interface strength on the mechanical properties of the HMX-MDB propellant were investigated by PCZM. The different ratio of matrix strength to interface strength revealed three kinds of HMX-MDB propellant’s damage patterns: matrix cracking, matrix cracking followed by particle-matrix interfacial debonding, and particle-matrix interfacial debonding followed by matrix cracking.

Propellants are the main energy source in the internal ballistic process. The use of 3D printing has promised to produce propellants with complex geometries. However, due to the degraded energy properties of propellants using an inert binder, there is a critical need to develop a printable energetic resin. In this paper, a novel energetic acrylate-terminated poly–3–nitratomethyl–3–methyloxetane (APNIMMO) oligomer was prepared and characterized. The performance of a new composite propellant composed of APNIMMO and CL-20 (Hexanitrohexaazaisowurtzitane Dodecane) was also demonstrated. The new energetic printable resin and its composites are suitable for stereolithography (SLA) 3D printing, offering not only an improved thermodynamic energy, but also a substantially improved burn rate. Compared with the inert binder, the energetic binder offers the possibility to improve the thermodynamic energy by 15% and the burn rate at 100 MPa by 480% for 3D printed propellants.

Thermal expansion coefficient (CTE) is a critical parameter of gun propellant because of its major role in fabrication, storage and combustion performance of the propellant. Further, controlling the CTE of the propellant is an effective solution to improve its loading density. Therefore, it is important to understand the thermal expansion of the propellant. To obtain the linear CTE of insensitive gun propellant, different weight percentages of NQ are added to the B# double-base absorbent propellant, the thermal mechanical analyzer (TMA) is employed to estimate their dimensional change over the temperature range of 213–323 K. The pure NQ flaky gun propellant exhibits a negative thermal expansion with a linear CTE of −2.006 × 10−4 mm/mm K−1. The results show that the linear CTE of the B# double-base absorbent propellant is decreased by 53.74% as the concentration of NQ is increased to 30%, whereas the glass transition temperature increases with increasing the NQ content. Keywords: Flaky gun propellant, Coefficient of thermal expansion (CTE), Negative thermal expansion, Nitroguanidine (NQ), Glass transition temperature

Metered dose inhalers (MDIs) are the mainstay of asthma and COPD therapy worldwide. MDIs are pocket-size hand-held drug delivery devices utilizing the energy of compressed propellant(s) for the aerosol generation. They have the drug in solution or suspension in the propellant(s), or a mixture of propellant(s) and co-solvent(s). MDIs deliver small metered drug doses either directly, or via add-on devices to the patients, like spacers or Valved Holding Chambers (VHCs) that can be used with a fast mask for children. Add-on devices hold the aerosol to eliminate the problem with correct synchronisation of dose release and inhalation. They also provide the possibility to inhale a single dose in multiple breaths. Alternatively, breath-actuated MDIs have been developed. MDIs have the disadvantage that the currently used hydrofluoroalkane (HFA) propellants contribute to the greenhouse effect and re-formulation of the drug formulations on short notice with more environment-friendly propellants may be necessary.

Non-Methane Volatile Organic Compounds (NMVOCs) from domestic aerosol sprays are\ud emerging pollutants and have substantial negative effects on human health and the\ud environment. This study, for the first time, carried out quantification of the NMVOC emissions\ud from off-the-shelf domestic aerosol sprays, at “source” in the UK. These aerosol sprays\ud contain harmful organic compounds as propellants and products. The results showed that the\ud cosmetic category (i.e. body sprays) have higher concentrations of NMVOCs with 93.7 wt.%\ud per can compared to households (i.e. air fresheners) with 62 wt.%. Also, water-based\ud products showed less NMVOCs in all analyses compared to solvent-based formulations. Direct\ud replacement of Liquefied Petroleum Gas (LPG) propellants from conventional products with\ud ‘clean air’ (i.e. nitrogen) showed the potential emission reduction of 50%. Hair spray products,\ud however, have the highest ozone forming potential with about 105.1 g of Ozone per litre of\ud the product compared to other domestic aerosol sprays. The level of global warming\ud contribution of the selected aerosol sprays in the UK was measured to be 129.8 ktCO2e in\ud 2018 and globally, this can be projected to be 3154.6 ktCO2e in 2020. Furthermore, NMVOC\ud emissions contribution from the domestic aerosol sprays in the UK was measured as 61.2 kt\ud in 2018 based on annual consumption of 520 million cans. Globally this can equate to 1437.6\ud kt based on the projected usage of 17.5 billion cans. Therefore, it is vital to expedite replacing\ud LPG propellant with nitrogen in a drive for a ‘near-zero’ emission in aerosol industry. The\ud results presented in this study can also be used to steer policy makers to the potentially\ud brewing danger from an otherwise passive emission source.

The established state-of-the-art for composite solid propellant grain manufacture consists in mix-cast-cure process using hazardous chemicals and specific molds for propellant forming. In most of the cases, polyaddition of oligomers involves isocyanate functional groups. Construction constraints limit the feasibility of propellant geometries, confining the pressure-time history of rocket motors to some established configurations. Composition pot-life becomes one of the most important parameters in the definition of correlated industrial processes. An additive manufacturing process for propellant grain production based on UV curing has been recently proposed for patenting. This technique enables more complex grain geometries, paving the way for new propulsive missions, thanks to customized thrust-time profiles or local composition fine tuning. The new curative method makes innovative use of pre-polymers, replacing isocyanates with UV-sensitive components characterized by lower chemical hazard for operators. The paper illustrates the experimental results obtained during the preliminary test campaign on propellant inert simulators, produced as lab-scale proof of concept. The activity targeted mono-layer samples, focusing on binder properties. Dynamic-mechanical thermal analysis, thermal-gravimetric analysis, and stress–strain tests have been carried out to measure mechanical and physical characteristics of different formulations. Both hydroxyl-terminated polybutadiene (HTPB) and polybutadiene diacrylate (PBDDA) binders have been considered, using ammonium sulfate as substitute of ammonium perchlorate. Aluminized and non-aluminized samples have been manufactured to evaluate the impact of opaque-reflective materials during the UV curing process. The reported analyses show comparable results for both the new UV-cured materials and demonstrate the feasibility of propellants based on isocyanate-free process.

Commercial spray products are commonly used in daily life and airborne particles generated by these products may cause adverse health effects. Our study was aimed to characterize the behaviors of airborne particles from spray products and to determine the deposition loss rate. Four categories of spray products with highly frequent use - air fresheners, fabric deodorants, window cleaners, and a bathroom cleaner - were selected for the study. The products were applied in a cleanroom according to the instructions for use. Airborne particles (10–10,000 nm) were measured within the breathing zone of a user with a scanning mobility particle sizer and an optical particle spectrometer. Additionally, filter sampling was performed to examine the morphological characteristics of the particles using a field emission-scanning electron microscope (FE-SEM). The initial concentration and particle size distribution varied among different spray types and products. Two propellant-type air fresheners that we tested showed a high initial concentration of smaller sized particles. However, one of these and all hand-pressure type propellants showed a low initial concentration in all size ranges. We observed that particles in nucleation mode (10–31.6 nm) decreased and aggregated particles shifted to accumulation mode (100–1,000 nm) over time. The FS-SEM analysis confirmed the aggregation of nano-sized particles for all products. The deposition loss rates of various particle sizes depended on the initial concentration and distribution of particle sizes. For two air fresheners with high initial concentrations, the loss rate of small-sized particles was higher than that of the other products whereas the particle loss rate of large-sized particles was higher, regardless of initial concentration. The results of this study can give us useful information in the behaviors of airborne particles in the consumer spray products and resulting exposure assessment especially in the application to the exposure modeling of spray products.

The aging behavior of softening composite solid propellant was investigated by measuring its mechanical and ballistic prosperities during prolonged storage at elevated and room temperatures. Accelerated aging was conducted at 65 °C for 231 days while the normal aging was performed at 25 ± 3 °C and relative humidity less than 50% for 8 years. The mechanical properties were obtained from uniaxial tensile tests for the aged propellant specimens while the ballistic properties were determined from static firing tests of subscale motors aged for 112 days at 65 °C. The mechanical results show that the maximum tensile strength and Young’s modulus initially increase and subsequently decrease with increasing aging time, while the maximum tensile strain generally increases with increasing aging time. The ballistic properties like burning rate show a small change which cannot affect the ballistic performance. The experimental results show that the changes in the mechanical properties are significant during the aging period, but the burning rate does not undergo significant changes. From this study, it is observed that the propellant ages through a combination of reactions like post-cure, oxidative cross-linking, chain scission, and hydrolysis. The chain scission and the hydrolysis effect are the most significant process, which makes the propellant soft and extendible. The observed aging mechanism has been modeled using an exponential function with two terms which can describe the complex behavior of the aging. By applying Arrhenius equation, the activation energy values were obtained based on the propellant mechanical properties. The shelf life of this propellant formulation at 25 °C is predicted to be 13 years using the modulus as failure criteria and control parameter. Keywords: Aging, Hydrolysis, Service life, Softening behavior, Solid propellant

This paper presents a method of thermal state calculation of combustion chamber in small thrust liquid rocket engine. The goal is to predict the thermal state of chamber wall by using basic parameters of engine: thrust level, propellants, chamber pressure, injection pattern, film cooling parameters, material of wall and their coating, etc. The difficulties in modeling the startup and shutdown processes of thrusters lie in the fact that there are the conjugated physical processes occurring at various parameters for non-design conditions. A mathematical model to predict the thermal state of the combustion chamber for different engine operation modes is developed. To simulate the startup and shutdown processes, a quasi-steady approach is applied by replacing the transient process with time-variant operating parameters of steady-state processes. The mathematical model is based on several principles and data commonly used for heat transfer modeling: geometry of flow part, gas dynamics of flow, thermodynamics of propellants and combustion spices, convective and radiation heat flows, conjugated heat transfer between hot gas and wall, and transient approach for calculation of thermal state of construction. Calculations of the thermal state of the combustion chamber in single-turn-on mode show good convergence with the experimental results. The results of pulsed modes indicate a large temperature gradient on the internal wall surface of the chamber between pulses and the thermal state of the wall strongly depends on the pulse duration and the interval. Keywords: Combustion chamber, Film cooling, Mathematical model, Nonstationary thermal mode, Small thrust liquid rocket engine, Steady pulse mode, Thermal state

We describe the critical condition necessary for the inner cylinder radius of a rotating detonation engine (RDE) used for in-space rocket propulsion to sustain adequate thruster performance. Using gaseous C2H4 and O2 as the propellant, we measured thrust and impulse of the RDE experimentally, varying in the inner cylinder radius ri from 31 mm (typical annular configuration) to 0 (no-inner-cylinder configuration), while keeping the outer cylinder radius (ro = 39 mm) and propellant injector position (rinj = 35 mm) constant. In the experiments, we also performed high-speed imaging of self-luminescence in the combustion chamber and engine plume. In the case of relatively large inner cylinder radii (ri = 23 and 31 mm), rotating detonation waves in the combustion chamber attached to the inner cylinder surface, whereas for relatively small inner cylinder radii (ri = 0, 9, and 15 mm), rotating detonation waves were observed to detach from the inner cylinder surface. In these small inner radii cases, strong chemical luminescence was observed in the plume, probably due to the existence of soot. On the other hand, for cases where ri = 15, 23, and 31 mm, the specific impulses were greater than 80% of the ideal value at correct expansion. Meanwhile, for cases ri = 0 and 9 mm, the specific impulses were below 80% of the ideal expansion value. This was considered to be due to the imperfect detonation combustion (deflagration combustion) observed in small inner cylinder radius cases. Our results suggest that in our experimental conditions, ri = 15 mm was close to the critical condition for sustaining rotating detonation in a suitable state for efficient thrust generation. This condition in the inner cylinder radius corresponds to a condition in the reduced unburned layer height of 4.5–6.5., Available online: 11 August 2018. ファイル公開：2020-08-11

The thrust characteristics of electrostatic-magnetic hybrid acceleration have been obtained in the central–cathode electrostatic thruster (CC-EST), which has a diverging magnetic nozzle and injects propellant through an annular slit on the inner surface of a ring anode. The electrostatic and electromagnetic contributions are evaluated by fitting experimentally measured thrust characteristics to formulae derived for steady-state, quasi-neutral plasma flows. The thrust is composed of two terms: one corresponds to the electrostatic acceleration interfaced by the magnetized electrons, the other to electromagnetic acceleration yielding a kinetic energy of swirl motion that is converted to that of axial motion in the diverging magnetic nozzle. The thrust performances of three thrusters having an axisymmetric shape with a hollow cathode on the centre axis and a diverging, applied magnetic field are examined. While only an electromagnetic thrust component is obtained in the conventional applied-field magneto-plasma-dynamics thruster, the other two, CC–ESTs, exhibit comparable, electrostatic and electromagnetic thrust components, and thus improve thrust performance., ファイル公開：2020-11-01

An experimental study on the Klystron effect of periodic injection modulated by pressure drop fluctuations on subsequent atomization is conducted. Time-resolved atomization backlit images and atomization Mie scatter images are captured by using the high speed camera. It is found that periodicity of forced atomization relies on pressure drop fluctuation amplitude and phase differences between atomization and pressure drop fluctuations relate to fluctuation frequencies. This feature of periodic atomization induced by Klystron effect corresponds to periodicities and high amplitudes of pressure fluctuations in unstable combustion chambers and chaos and low amplitudes of pressure fluctuations in stable combustions chambers. Drastically periodic varying of gross surface area of droplets with time was shown in Mie scatter images. The importance of periodic impinging jet atomization modulated by pressure drop fluctuations for acoustic liquid propellant combustion instabilities is illustrated. Keywords: Atomization, Combustion stability, Impinging jet injector, Klystron effect, Liquid rocket engine, Mie scatter

In order to obtain the dynamic characteristics of a differential piston warm gas self-pressurization system for liquid attitude and divert propulsion system, a transient model is developed using the modular modeling method. The system includes the solid start cartridge, pressure-amplified tank with liquid monopropellant, liquid regulator, gas generator, and pipes. The one-dimensional finite-element state-variable model is applied to the pipes and the lumped parameter method is adopted for the other modules. The variations of the system operation parameters over time during the startup, steady-state, and pulsing operational processes are obtained from the transient model, and the characteristics of starting time changing with different system parameters are also analyzed. It is shown that the system startup process can be divided into three distinct processes. The starting time monotonically changes with variations of the liquid regulator parameters, first decreasing and then increasing with the mass change of the solid propellant charge of the start cartridge, initial gas cavity volume of the pressure amplified tank and initial gas cushion of the propellant tank. The starting time can be reduced to less than 1.0 s (0.68–0.75 s for the current system). For meeting the deviation requirements of ±10% of the steady-state propellant tank pressure, the positive deviation requirement is assured by the self-locking pressure and the negative deviation can be assured within an allowable maximum propellant tank volume flowrate (1.6 times the design value for the proposed system) for downstream thrusters for a designed system. The results from the simulation are useful as a guide for further system design and testing. Keywords: Differential piston, Dynamic characteristics, Dynamic models, Liquid attitude and divert propulsion system, Startup process, Warm gas self-pressurization system

By using the infrared thermographic method, a non-destructive testing technique was applied to detect the surface-temperature evolution of solid propellants during strain-control fatigue tests within finite cycles. When the applied strains were below the viscoelastic limit, two stages of temperature variation were observed before the initiation of macroscopic cracks: an initial temperature-increase stage, and a steady temperature state. Thermodynamic analysis was carried out and a method was developed to allow the acquisition of stored energy at different stages of cyclic loading, which can reflect the material damage on the macroscopic scale. In addition, temperature localization during fatigue was observed, which implied the occurrence of damage accumulation and crack propagation. The results reveal that the cyclic-loading induced temperature increase of solid propellants during the fatigue process has a significant effect on solid rocket motors in a transportation or storage state. Keywords: Fatigue, Heat dissipation, Infrared thermography, Solid propellants, Stored energy

The infrared (IR) irradiance signature from rocket motor exhaust plumes is closely related to motor type, propellant composition, burn time, rocket geometry, chamber parameters and flight conditions. In this paper, an infrared signature analysis tool (IRSAT) was developed to understand the spectral characteristics of exhaust plumes in detail. Through a finite volume technique, flow field properties were obtained through the solution of axisymmetric Navier-Stokes equations with the Reynolds-averaged approach. A refined 13-species, 30-reaction chemistry scheme was used for combustion effects and a k-e-Rt turbulence model for entrainment effects. Using flowfield properties as input data, the spectrum was integrated with a line of sight (LOS) method based on a single line group (SLG) model with Curtis-Godson approximation. The model correctly predicted spectral distribution in the wavelengths of 1.50–5.50 μm and had good agreement for its location with imaging spectrometer data. The IRSAT was then applied to discuss the effects of three operating conditions on IR signatures: (a) afterburning; (b) chamber pressure from ignition to cutoff; and (c) minor changes in the ratio of hydroxyl-terminated polybutadiene (HTPB) binder to ammonium perchlorate (AP) oxidizer in propellant. Results show that afterburning effects can increase the size and shape of radiance images with enhancement of radiation intensity up to 40%. Also, the total IR irradiance in different bands can be characterized by a non-dimensional chamber pressure trace in which the maximum discrepancy is less than 13% during ignition and engine cutoff. An increase of chamber pressure can lead to more distinct diamonds, whose distance intervals are extended, and the position of the first diamond moving backwards. In addition, an increase in HTPB/AP causes a significant jump in spectral intensity. The incremental rates of radiance intensity integrated in each band are linear with the increase of HTPB, and the growth rates of radiance intensities in some bands reach up to 50% as HTPB weight increases by 3%.

The mechanism of stabilization of double-base rocket propellants (DBPs) using inorganic stabilizer (nano- and micro-clinoptilolite) was investigated. The surface structures of the stabilizers, the double-base propellants containing the new stabilizers and the effect of the stabilizers on the surface behavior of propellants and vice versa were checked using XRD and Electron Microscope (AFM and TEM techniques). The results obtained from XRD suggested that the crystalline structure of the new inorganic stabilizers was completely changed when it was introduced into the propellants which may be attributed to the pressing processes of DBPs with stabilizers under very high pressure during their mixture preparation. The results obtained from Atomic Force Electron Microscope (AFM) and TEM indicated that nano-clinoptilolite particles become more regularly arranged on the propellant surface than micro-clinoptilolite which gives the stabilizer a higher ability to absorb more nitrogen oxide. The work aimed chiefly to use zeolite stabilizers for DBPs instead of classically used organic compounds; in order to avoid the harmful and carcinogenic organic products coming from the reaction of NOx gases with these organic stabilizers. This is achieved by studying the thermal behavior of these zeolites via investigation of their surface interaction with NOx gases obtained during stabilization process and suggesting possible interaction mechanism.

Accepted: 2017-01-13, 資料番号: SA1160256000

The article of record as published may be found at https://doi.org/10.1016/j.jsse.2020.07.026 Attention to the need to deorbit spacecraft at end-of-life will only intensify as space-faring nations grapple with the problems of increasingly crowded orbits. The use of drag devices for spacecraft self-disposal in low Earth orbit (LEO) merits review now. Drag-enhancing deorbit devices can be used to offset propellant reserved for deorbit operations and to extend mission life. Drag devices offer increased compliance with debris mitigation standards and can also offer other system-level advantages such as mass savings. We analyze the "trade-space" for spacecraft that are best suited to make use of this technology with specific examples. Our findings suggest that drag devices are well suited for both small and mid-sized LEO spacecraft. The suitability and feasibility of drag devices for small spacecraft (

This paper reports the design of some of the new ionic based high energy materials derived from the anion of picric acid (2,4,6-trinitrobenzene-1-ol), styphnic acid (2,4,6-trinitrobenzene-1,3-diol) and 2,4,6-trinitrophloroglucinol (2,4,6-trinitrobenzene-1,3,5-triol) and cation derived from the key synthon molecules such as 5-trifluoromethyl-1H-tetrazole, 5-dinitromethyl-1H-tetrazole and 5-azido-1H–tetrazole-1-carbonitrile. The detonation properties of these newly proposed compounds are predicted by using software such as EXPLO-5, EXTEC and LOTUSES and Keshvarz method. Moreover, other explosive parameters such as density, gurney velocity, and oxygen balance and decomposition products of the newly designed molecules have also been predicted and reported for the first time in this manuscript. The predicted detonation parameters of some of the newly designed compounds exhibit higher velocity of detonation (VOD) and detonation pressure in comparison to other well-known benchmark explosives such as 2,4,6-trinitrotoluene (TNT) and 1,3,5-trinitro-1,3,5-triazinane (RDX). Further, the peak over pressure (POP) and the blast impulse parameters of the newly designed compounds are predicted by using Shock physics explicit eulerian dynamics (SPEED) software, and the same is reported for the first time in this work. The work also reports the theoretical prediction of impact and electrostatic spark sensitivity parameters for the newly designed molecules. The ballistic performance parameters of the newly designed ionic energetic materials are also predicted by incorporating them into model composite rocket propellant formulations. The predicted ballistic parameters indicate that the proposed materials may find an application in the propellant formulation as an energetic additive., Organic chemistry; Theoretical chemistry, EXTEC; EXPLO-5; SPEED; Blast impulse; Peak over pressure

The cold spray technique was employed in this study to produce pure nickel (Ni) and nickel-Inconel 718 powder deposits on duplex stainless steel substrates. High quality coatings were manufactured using nitrogen as the propellant gas. The coatings exhibited satisfactory Inconel 718 (In718) retention within the Ni matrix due to the highly ductile binder phase necessary for fabrication. The characterization of the coating microstructures was implemented by means of X-ray diffraction, electron microscopy, energy-dispersive X-ray spectroscopy, and microhardness testing. In addition, the erosion resistance of both coatings was evaluated by performing cavitation erosion tests, with the analysis of the eroded surfaces revealing different erosion mechanisms for each coating. The results demonstrated the efficiency of the cold spray technique for use in the production of metal-metal matrix composite coatings and the potential of In718 use in applications which demand enhanced cavitation erosion resistance. This is the peer-reviewed version of the article: Surface and Coatings Technology, 2020, 382, 125195, [https://doi.org/10.1016/j.surfcoat.2019.125195] The published version: [http://cer.ihtm.bg.ac.rs/handle/123456789/3274]

Dimethyl ether has received great interest since it is considered as an environmentally-friendly chemical and can be utilized in a wide range of applications. DME is mainly used as aerosol propellant and intermediate in oil industry. It is also used as clean fuel in power generation and diesel engines. DME is produced directly from synthesis gas and indirectly via methanol dehydration over commercial γ-Al2O3 pellets. The methanol dehydration process has a low conversion which is limited by the thermodynamic equilibrium leading to a large recycle stream and size equipment. In addition, the presence of water in the feed with methanol reduces the equilibrium conversion that can be attained. An inter-stage membrane for water removal can overcome the thermodynamic limitations related to the conventional process. In this study, two adiabatic reactor stages with inter-stage hydrophilic ceramic membrane for water removal were considered for the synthesis of DME via methanol dehydration. A commercial conventional process was simulated using Aspen-Plus ™ for an annual plant capacity of 50,000 metric tons and a 99.5 wt. % purity of DME product. A modified process for DME synthesis using water-selective alumina-silica composite membrane was developed and simulated. A parametric study was conducted to show the effect of inlet temperature on the reactor conversion which can be increased to 93 % using the inter-stage membrane. The high temperature reactor effluent was used to preheat methanol feed. The modified process shows potentials to reduce both energy and equipment size which leads to a lower process cost for DME production.

Accepted: 2016-11-10, 資料番号: SA1160365000

Nanoenergetic materials became available during last three decades and offer a promise of much higher energy densities and faster rate of energy release than conventional explosives and solid propellants. The energy densities in nanoenergetic devices can reach 50 MJ/kg. Extremely fast burning rates exceeding 3 km/s can be realized in the combustion of porous silicon nanoenergetic composites. The chapter presents a brief review of novel achievements in studying the reaction mechanisms of nanoenergetic materials. This knowledge would allow the creation of molecularly manipulated energetic substances and formulations with well-tailored chemical and physical properties.

Aluminum powders of different particle sizes (from nano- [as = 0.01–0.5 μm] to micron-sized (as = 1–500 μm) and particle shapes (spherical and flakes) are widely used in energetics: from rocket propulsion to self-propagated high-temperature synthesis and water-splitting hydrogen-generating compositions. In this chapter the features of aluminum powder properties will be discussed along with their reaction kinetics and oxidation behavior in different oxidation media: from weak oxidizers (water and air) to strong oxidizers such as rocket propellants.

Composite propellant based on hydroxyl-terminated polybutadiene and ammonium perchlorate (AP) has become the workhorse propellant for modern-day missiles and space vehicles. AP is the main ingredient and is used as an oxidizer in composite propellant and accounts for approximately 70%–80% of the composition. AP plays a vital role in tailoring the burning rate of the propellant using multimodal particle size distributions and provides strength to the propellant as filler. AP is ground to different particle sizes for use in propellant formulations to achieve different burn rates and higher solid loading. Grinding of AP leads to generation of a large surface area with excess surface energy and there is a tendency for agglomeration, segregation, caking, bridging, and no flow in silo/bin/hopper, stockpile, feeder, chute, conveyor, etc. The propensity of the problems increases with increase in the time gap between grinding and mixing operations (with all propellant ingredients). Here, various grades (particle size fractions) of AP were used in propellant formulations and characterized with respect to physical and flow properties. Physical properties that were studied were particle size, shape, density, and moisture. Particle size distribution was determined using the sieve analysis and laser light scattering technique. Powder flowability was measured using shear strength, angle of repose, and tapped-to-bulk density measurements. The values of Hausner ratio and Carr index are highest for ultrafine AP, indicating that it is a highly compressible powder, whereas Hausner ratio and Carr index are lowest for the coarse AP, which indicates its free-flowing behavior.

An indirect optimization procedure is presented to minimize the propellant consumption for finite-burn transfers with two practical thrust control models, namely, inertially fixed thrust and fixed-plane linearly varying thrust direction. The optimality equations are derived with theory of optimal control and the consequent boundary value problem is solved with a procedure based on Newton's method. A homotopic approach is used to find suitable tentative solutions and assure convergence. The method is applied to the optimization of Moon escape trajectories with accurate dynamic models and proves to be fast and accurate.

The use of different additives to modify the burning rate (BR) of composite solid propellants (CSPs) has been a topic of much study during decades within rocket motor technology. This perspective chapter initially describes the effects of the performance of ammonium perchlorate-based composite propellants by means of employing nanopowder transition metal oxides as BR catalysts in comparison with traditional microfillers, the most representative and active of which are iron(III) oxide and cupric oxide. In addition, titania represents a new generation of BR catalyst with great potential in this field. CSPs with microparticles as catalyst seem to be less stable due to oversensitivity to pressure variations, but the nanostructured propellants yield highly stable BRs over a broad pressure range. On the other hand, a brief review of surveys of recent progress on the synthesis, macrostructural characterization, and properties of metallo-polyurethanes, as well as their applications as advanced binders in the formulations of CSPs based on hydroxyl-terminated polybutadiene, is addressed. The incorporation of both metallic oxide nanoparticles and polymeric organocatalysts based on ferrocene in the formulations of these energetic materials could improve their ballistic and thermal combustion properties. The latter characterization was commonly done by thermal analytical techniques, such as thermogravimetry analysis and differential scanning calorimetry.

Several industrial and research-type composite solid propellants containing different nanosized metal particles (such as aluminum, zirconium, and titanium) with similar nominal composition were prepared and experimentally analyzed. The effects of different nanosized powders on the rheological properties of metal/hydroxyl-terminated polybutadiene (HTPB) slurries and ammonium perchlorate-based composite propellant slurries were investigated. The strand burning rate and the associated combustion flame structure of propellants were determined. The energetic properties and the hazardous properties of propellants prepared were analyzed and the properties mentioned earlier were compared to those of a conventional aluminized propellant. The results show that aluminum powder can be sufficiently dispersed in HTPB binder. The propellant formulations containing nanosized particles are sensitive to impact and friction except for the base propellant with common aluminum powder, which is less sensitive to friction as compared to the other compositions. Nanosized additives can affect the combustion behavior and increase the burning rate of propellants compared with the common reference propellant composition.

Superthermites have been regarded as a series of promising materials in defense science and technology. The preparation and characterization of superthermites and their applications in double-base propellants are systematically investigated in this chapter. The superthermites Al/PbO, Al/CuO, and Al/Bi2O3 were obtained using different methods, and their compositions, structures, and morphologies are characterized. The results confirm that they are composites instead of physical mixtures. To understand the formation mechanisms of superthermites, investigations were conducted on the decomposition reaction of the precursor for Al/CuO. The apparent activation energy of this reaction was calculated to be 164.1 kJ/mol and the Mampel power law with n = 1/4 ( d α / d t = 10 14.0 × 4 α 3 / 4 e − 2.0 × 10 4 / T ) was proposed to describe the decomposition reaction. The introduction of the three superthermites, especially Al/PbO and Al/Bi2O3, into double-base propellant results in enhanced burning rates and decreased pressure exponents, indicative of their high catalytic activity on propellant combustion.

Aerosol packaging in the food industry is limited to niche applications such as whipped cream that is very popular worldwide thanks to the high solubility of N2O that allows foaming restitution after the cream is expelled from the can. Flat product restitution is more diverse in culinary ingredients and needs only a nonsoluble propellant gas such as nitrogen. Only gaseous molecules described in the other chapters of this book are authorized by law for food aerosol use, taking into account that liquefied petroleum propellant molecules are dedicated to cosmetics and technical products.

Solid propellants used in microthrusters are the key energetic materials required to burn stably in microscale thruster chambers. Conventional propellants, such as hydroxyl-terminated polybutadiene/ammonium perchlorate, HMX-based propellant, RDX-based propellant, and glycidyl azide polymer propellant are not suitable for microthrusters since their critical burning diameters are too large to burn stably in microscale chambers. Nanothermite-based propellants are useful propellants for microthrusters. The propellants consist of a thermite and a gas-generating substance, in which nano-Al/CuO and nitrocellulose are the best substances for solid propellant. Thermite-based propellants can be synthesized or prepared by ultrasonic mixing, electrospray mixing, and the sol–gel process, in which the sol–gel process is the best to synthesize homogeneous thermite-based composites. The research results show that thermite-based propellant is suitable for the microscale chambers of microthrusters, but the energy efficiency of the thermite-based propellant is very low in microthrusters, whose specific impulse is lower than 30 s. How to improve the efficiency of energy is a key technology of nanothermite-based propellant and microthrusters for the future.

The introduction of nanosized energetic ingredients first occurred in Russia about 60 years ago and aroused great expectations in the rocket propulsion community, thanks to higher energy densities and faster energy release rates with respect to conventional ingredients. In spite of intense worldwide research programs, today mostly laboratory-level applications are reported and often for scientific purposes only. A number of practical reasons prevent applications at the industrial level: inert natural coating of the energetic particles, nonuniform dispersion, excessive viscosity of the slurry propellant, possible limitations in mechanical properties, more demanding safety issues, and so on. This chapter describes the main features in terms of performance of solid rocket propellants loaded with nanometals and intends to emphasize the unique properties made possible by the addition of nanosized energetic ingredients. Steady combustion regimes are examined in detail, including particle coating, chemical and mechanical activation, condensed combustion products, and the resulting gravimetric specific impulse. Unsteady combustion regimes, aggregation/agglomeration phenomena, hazards evaluation, and aging processes are treated in companion papers under preparation.

International audience; A new mechanism of thermo-acoustic destabilization of combustors powered by burning droplets is investigated. Acoustic induced heat release disturbances are found to be driven by oscillations of droplet lifetime during the final stage of the droplet combustion before their extinction. The mechanism is the following. Synchronized acoustic disturbances alter the droplet evaporation rate and modify their diameter. Perturbations of the droplet diameter along the droplet trajectories lead to fluctuations of their lifetime. These fluctuations trigger in turn an oscillating motion of the burning droplet cloud boundary that is synchronized by the acoustic excitation. This leads to large heat release fluctuations before droplet extinction and constitutes a thermo-acoustic source. This mechanism is found to be particularly important in solid rocket motors in which aluminum droplets released from the propellant burn individually and are quenched as the droplet diameter falls below a critical residue diameter associated to an inert particle. Analytical models are derived in an idealized configuration where acoustic forcing takes place in the transverse direction to the droplet trajectories. Expressions are derived in the frequency space for the droplet diameter oscillations, the motion of the droplet cloud boundary and the resulting heat release disturbances, that take the form of a distributed Flame Describing Function. This model is used to reveal effects of the acoustic pressure level and of the residue diameter on the resulting motion of the combustion volume boundary and corresponding heat release disturbances. It is further extended to consider the peculiar flow in a solid rocket motor and the heat release disturbance model is compared to results from numerical flow simulations of the motor. While highlighted in the context of solid propulsion, the observations made are believed to be more general and the same mechanisms may persist in acoustically perturbed spray flames from hydrocarbon fuel droplets.

This paper focuses on the stability of capillary forced flow. In space, open capillary channels are widely used as the liquid and gas separation devices to manage liquid positioning and transportation. Surface collapse happens when the flow rate exceeds the critical value, leading to a failure of propellant management. Knowledge of flow rate limitation is of great significance in design and optimization of propellant management devices (PMDs). However, the capillary flow rate limitation in an asymmetry channel has not been studied yet in the literature. In this paper, by introducing an equivalent angle to convert the asymmetry corner to a symmetry one, the one-dimensional theoretical model is developed. The flow rate limitation can then be investigated as a function of the channel geometry as well as liquid property based on the model. Comparisons between the asymmetry and symmetry channels bring forth the characteristics of the two kinds of channels, and demonstrate good accordance between the new advanced model and the existing one in the literature. This theoretical model can provide valuable reference for PMD designers.

Electromagnetic formation flight (EMFF) leverages electromagnetic force to control the relative position of satellites. EMFF offers a promising alternative to traditional propellant-based spacecraft flight formation. This novel strategy is very attractive since it does not consume fuel. Due to the highly coupled nonlinearity of electromagnetic force, it is difficult to individually design a controller for one satellite without considering others, which poses challenges to communications. This paper is devoted to decoupling control of EMFF, including regulations, constraints and controller design. A learning-based adaptive sliding mode decoupling controller is analyzed to illustrate the problem of existing results, and input rate saturation is introduced to guarantee the validity of frequency division technique. Through transformation, the imposed input rate saturation is converted to state and input constraints. A linear matrix inequalities (LMI)-based robust optimal control method can then be used and improved to solve the transformed problem. Simulation results are presented to demonstrate the effectiveness of the proposed decoupling control.

Dealing with explosive wastes appropriately is a difficult problem recently. Many of them are just being buried and some of them are being treated. The main method to treat explosive wastes is to burn it with Rotary-kiln reactor. However this has a risk of explosion during the process because it is based on explosion mechanism. In addition, Rotary-kiln method can not deal with enough amount of explosive waste because Rotary-kiln method use batch reactor. Lastly, because of the limit of non-successive process, the efficiency of the purification process is low which results in an incomplete combustion. This incomplete combustion makes the exhaust gas more dirty which contains more harmful substances. To overcome the limits of Rotary-kiln method which are introduced above, development of combustion-based process is necessary. Therefore, a flow diagram which includes fluidized bed reactor was developed with Aspen Plus. A mixture of explosive waste slurry and water with the same ratio was fed in the fluidized bed reactor. Also, additional units were chosen and designed which clean exhaust gas to make the reactor more adaptable. The proposed process was thermodynamically analyzed and the efficiency evaluation was held. Also, the improvement possibility of the process was derived. The final exhaust gas from this process satisfied the environment regulation of Korea. This process can deal with 3000 ton/yr which is the total amount of waste propellant in Korea, while the exhaust gas fits the regulation of CO 25ppm/hr, NO2 0.10ppm/hr. This study is expected to contribute to the improvement of the efficiency of the explosive waste treatment process and the possession of domestic technology in Korea.

Recently, there have been various studies how to incinerate explosive waste safely due to environmental problems and safety problems. Explosive waste requires a new incineration mechanism, because of the pollution gas generated during the conventional treatment process, such as rotary kiln or outdoor exploration. This study focuses on the fluidized bed incinerator technology which can burn the target material using only a small amount of air at a relatively low temperature condition. In this study, we simulated the process of burning Double Based Propellant (DBP) mixed with water in the incinerator safely. The fluidized bed incinerator was modeled as a cylinder with a diameter of 0.5 m and a height of 2.0 m, and a case study was carried out by changing the mixing ratio of the injected slurry. As a result, we confirmed the optimal mixing ratio with water for burning DBP without explosion, and confirmed that DBP combustion and decomposition reaction occurring inside the incinerator can be safely simulated. Based on this, it is considered that the design of the actual incinerator will provide a new direction for research on the explosive waste treatment process in Korea in the future.