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The equivalence ratio is one of the significant factors affecting the propagation characteristics and performance of the rotating detonation engine (RDE). Using the compressible reacting flow solver based on the OpenFOAM open-source platform, the effect of different equivalence ratios of premixed H 2 /Air gases on the propagation characteristics and performance of RDE under different total inlet pressures (P 0) is investigated. The reactants are injected through the discrete inlet to mimic the spatial inhomogeneity of the reactants in the actual RDE combustor. The results indicate a Y-shaped flow field structure is formed behind the rotating detonation wave (RDW) using the discrete inlet. There is only one RDW in the flow field with the change of the equivalence ratio when the P 0 is 0.5 MPa, and the primary factors affecting the RDW propagation velocity differ under fuel-lean and fuel-rich conditions. The RDW propagation mode switches from single-wave to co-directional double-wave and double-wave collision with the change of the equivalence ratio when the P 0 is 1.2 MPa. The velocity deficit of RDW in the double-wave mode is larger than that in the single-wave mode. Additionally, in the single-wave mode, the specific impulse decreases as the equivalence ratio increases, but the specific thrust increases as the equivalence ratio increases. When the propagation mode changes, the specific impulse and specific thrust show different trends with the change of the equivalence ratio. It demonstrates that multiple co-directional RDWs have the effect of stabilizing the thrust. [Display omitted] • A Y-shaped flow field structure is formed behind the rotating detonation wave. • Effect of equivalence ratio on the rotating detonation wave propagation velocity. • Effect of the equivalence ratios on formation time of stable detonation wave. • Effect of equivalence ratio on the rotating detonation wave propagation mode. • The performance of RDE under different equivalence ratios is analyzed. [ABSTRACT FROM AUTHOR]

A three-dimensional numerical simulation of a combustion chamber with a continuous rotating detonation wave was studied. A nozzle-type main system feeding was located at the end part of the chamber, and an additional feeding from the inner and outer walls of the annular channel was considered. A hydrogen-oxygen mixture was considered as a fuel. A detailed kinetic mechanism by Z. Hong (2010) consisting of 20 reactions with 9 components was used to simulate chemical processes. Various relations of the fuel were considered: lean mixture [H 2 ]:[O 2 ] = 1:1, stoichiometric mixture [H 2 ]:[O 2 ] = 2:1, and rich mixture [H 2 ]:[O 2 ] = 3:1. Besides, either the same mixture, or additional oxygen, or air was supplied from the side-wall feeding system. For all the considered feeding options, a detonation mode was obtained, and the traction characteristics were calculated. • A simulation of an engine powered by a hydrogen-air mixture was carried out. • Hong's kinetic mechanism was considered. • The influence of the fuel concentration on the processes in the chamber was studied. • Traction characteristics were obtained and analyzed for all cases. [ABSTRACT FROM AUTHOR]

• Solar thermal propulsion system combined with thermal energy storage is proposed to overcome thrust failure in shadow area. • The whole process from light concentrating, heat storage, to trust generation is numerically investigated and verified by ground tests. • The time to complete heat storage is within the illumination time of low Earth orbit. • Thrust and specific impulse can achieve 2 N and 691 s respectively in solar eclipse region. Possessing relatively high specific impulse and moderate thrust levels, solar thermal propulsion (STP) is a promising candidate in spacecraft propulsion system. However, the traditional solar thermal propulsion system suffers from thrust failure in the shadow area, which seriously affects its applicability. In this paper, we investigate feasibility of regenerative solar thermal propulsion system (RSTP) incorporating thermal energy storage, which can effectively overcome unmatched synchronous working time and illumination time. A numerical model for RSTP considering the whole energy transfer process from light concentrating, heat storage, to thrust generation is built, which is verified by experiment measurements with relative errors less than 15 %. The result shows that the maximum time to complete heat storage is about 4000 s, which is within the illumination time for low Earth orbit. In the solar eclipse region, the thrust (F t) and the specific impulse (I sp) of the system increase with the propellant flow rate, which can reach about 2 N and 690 s, respectively. What's more, the system can operate for around 100 s continuously at the maximum thrust in the shadow area. This work provides alternative approaches for microsatellite propulsion with high specific impulse, high thrust, and continuous operation despite presence of solar eclipse. [ABSTRACT FROM AUTHOR]

Dual-combustion ramjets have the advantages of both ramjets and scramjets, and further research is required to understand the combustion flow state in supersonic combustors and improve the overall performance of engines. Based on the computational fluid dynamics method, the objective of this study was to explore whether the two distinct combustion flow modes, scramjet and ramjet modes, can emerge in a dual-combustion ramjet under the same operating conditions when it was burned with a constant area. If so, the combustion flow characteristics and comprehensive engine performance under the two modes would be compared. Kerosene C 12 H 23 was used as fuel and the equivalence ratio was 0.71. A 6-species 4-step chemical reaction mechanism was adopted, and the RNG k − ε turbulent flow model was selected. The altitude was 35 km and the Mach number was 7. The results show that, under the same operating conditions, the scramjet and ramjet modes can be achieved when the dual-combustion ramjet burns with a constant area. The generation of the two modes is related to the ignition efficiency, and the ramjet mode requires higher ignition efficiency. In the supersonic combustor, type X-shock waves with a first Mach stem length of 20 mm were generated under the scramjet mode, and a C-shock wave with a Mach stem length of 52 mm was generated under the ramjet mode. The core principle of mode transition is to control the rate of heat released in the supersonic combustor. The engine performance was better under the ramjet mode in this case. Specifically, the thrust, specific impulse, and specific thrust of the ramjet mode were all 1.4 times those of the scramjet mode. This is because, compared to the scramjet mode, the total pressure loss of the ramjet mode was slightly higher by 1.4%, but the combustion efficiency increased by 31.5%, significantly improving the engine performance. The ramjet mode can improve the combustion process and combustion efficiency. The reason is that, on one hand, the recirculation zone formed by the C-shock wave not only has a larger number of vortices, but also has a longer length, which is conducive to the efficient mixing of high temperature gas and secondary flow. On the other hand, the airflow velocity in the ramjet mode is lower, which is equivalent to an increase in the residence time of the airflow. [Display omitted] • The scramjet and ramjet modes can be achieved under the same operating conditions. • The generation of the two modes is related to the ignition efficiency. • Type C-shock and X-shock waves were formed. • The core principle of mode transition is to control the rate of heat released. • The performance of ramjet mode was better due to its high combustion efficiency. [ABSTRACT FROM AUTHOR]

• The oxyhydrogen-air explosion characteristics under water jetting propulsion scenes are revealed. • The propulsion performance of the explosion jetting is quantitatively analyzed. • The peak thrust of the aerial explosion jetting is dramatically reduced. • The asymmetrical water repulsion phenomenon is found. Hydrogen has attracted many researchers in the field of explosion jetting propulsion owing to its appealing advantages of zero carbon emissions, wide flammability range, and ease of preparation. However, the research gap in hydrogen explosion characteristics in water jetting scene has slowed down the advancement of the impulsive jetting propulsion field. In this work, static tests under three typical explosion jetting scenes (vertical submerged jetting, oblique submerged jetting, and aerial jetting) are conducted to reveal their explosion and thrust characteristics. The static tests results demonstrate that although the peak pressures of the aerial and vertical submerged explosion jettings are comparable, the peak thrust of the former is dramatically reduced by 40.6 %. We also find that the maximum pressure rise rate and peak pressure of the oblique submerged jetting are 22.9 % and 11.4 % lower than those of the vertical submerged jetting, respectively. Meanwhile, the peak thrust of the inclined jetting is also reduced by 14.1 % compared to the vertical jetting. High-speed photography applied to capture the evolution details of the oxyhydrogen-air explosion reveals that the preceding differences are attributed to the asymmetrical water repulsion phenomenon. Specific impulse is used to characterize the propulsion performance of the oxyhydrogen-air–water explosion jetting at varied volume fractions, and the results show that the maximum specific impulse is obtained for a given 0.45 kg thruster at 40 % water volume and 4:6 ratio of air to oxyhydrogen. The correctness of the explosion characterization metrics and thrust results are finally corroborated via the outdoor aquatic-aerial jetting propulsion tests. [ABSTRACT FROM AUTHOR]

The process of starting the rotating detonation engine initially filled with air in low outer pressure conditions was studied. Such engines can be used for bringing satellites into the Earth orbit. A three-dimensional numerical simulation of an engine combustion chamber with a rotating detonation wave of a cylindrical type with an internal body fed by a hydrogen-oxygen mixture or an acetylene-oxygen mixture was performed. The influence of the lengths of the combustion chamber channel on the stability of the detonation wave and thrust characteristics was considered. An author's computer code was used in simulations, the code was based on a model of multicomponent gas dynamics with chemical transformations and turbulence. The code was tested by comparing it with experiments and analytical solutions for simple cases. To describe the kinetics of acetylene combustion, a simplified chemical scheme of interaction of 10 components was used, without "heavy" radical species: [C 2 H 2 , CO, CO 2 , H 2 , O 2 , H 2 O, OH, O, H, N 2 ]. The influence of the chosen kinetic mechanism in case of hydrogen-oxygen mixture on the simulation results was studied. Various modes of the process in the detonation engine were obtained, including those with a stable detonation wave. • Under low external pressure conditions, the detonation engine starting was investigated. • A simulation of an engine powered by acetylene oxygen was carried out. • The influence of the fuel concentration on the processes in the chamber is studied. • Combustion chambers of two lengths are considered. • Traction characteristics were obtained and analyzed for all cases. [ABSTRACT FROM AUTHOR]

[Display omitted] The relative efficacy of real energetic plasticizers and polymers for model solid composite propellants comprising 25% aluminum hydride, 50% dinitramide ammonium salt and 25% binder (20% a plasticizer and 5% a polymer) has been estimated. The quantitative dependence of the efficiency of plasticizers on the value of their enthalpy of formation ΔH t 0, the oxygen coefficient α, percentage of hydrogen %H and density d has been revealed. 3,4-Dinitrofurazan tested as a plasticizer for the binder provides effective impulse values at the 3rd stage up to ∼2 s higher than those for other plasticizers. [ABSTRACT FROM AUTHOR]

The oblique detonation engine (ODE) has established a clear superiority for hypersonic flight because of its high thermal efficiency and compact structure. It has become the research hot spot all over the world in recent years. The aim of this study is to derive a criterion on the propulsive balance of ODE, from which we can find the key parameters governing the propulsive performance explicitly. A physical model of ODE is put forth, which consists of the inlet, the constant cross-section combustor and the divergent nozzle. The mathematical equations to calculate the thrust generated by the divergent nozzle and the pressure drag produced by the inlet are deduced. The net thrust of ODE is then obtained. The criterion shows clearly that the static temperature at the engine inlet exit is a very important parameter. The lower the inlet exit temperature is, the higher the specific impulse will be. The specific impulse of ODE with stoichiometric H 2 /air mixture and hydrocarbon/air mixture are calculated by using these equations. The results show that ODE can obtain positive net thrust from Ma 8 to Ma 15. • The formula to calculate specific impulse of oblique detonation engine is put forth. • The physical laws to generate thrust of oblique detonation engine is revealed. • Oblique detonation engine can generate net thrust in flight Mach number range of 8–15. [ABSTRACT FROM AUTHOR]

This study focused on the development of a vacuum cathode arc thruster (VAT), particularly on its design, manufacturing process, and demonstration. Characteristically, the proposed thruster does not require any additional propellant feeding system as the cathode electrode is simultaneously used as a propellant. In the ignition system, tiny spots are coated on the cathode surface to induce plasma flow. Such a setup has the advantages of simplicity, low price, small size, and low weight and is suitable for microsatellites. Moreover, a "trigger-less" method with an inductor storage power system was used for generating the pulsed plasma. This discharge method can significantly reduce input power. Thrust is mainly caused by high exhaust velocities of metal ions in the plasma flow, making ion density, ion velocity, and ion charge influential parameters. A battery was used instead of a power supply system to reduce the energy consumption of the entire VAT processing unit. The energy required for a single pulse was estimated to be 0.266 J, by measuring the change between the discharge current and the voltage. The ion current was measured using an ion detector and was 3.55 A, and the ion velocity was 23150 m/s. In the theoretical analysis, the VAT prototype proposed in this study achieved a single impulse of 4.31 μNs, a specific impulse of 1571 s, and a thrust efficiency of approximately 12.5%. • An inductor storage power system was used for generating the pulsed plasma. • A micro VACT with inductive energy storage circuit was designed and tested. • A battery was used to reduce the energy consumption of the entire VAT unit. • VAT prototype proposed in this study achieved a single impulse of 2.3 μNs. • A specific impulse is 2360 s, and a thrust efficiency is approximately 10%. [ABSTRACT FROM AUTHOR]

Boron based compounds have been looked upon as possible replacements to conventional toxic hypergolic fuels in propellant system. In an attempt to add to the pool of boron based hypergols, a series of new amine borane/ cyanoborane zwitterionic compounds were developed. Their hypergolic reactivity, performance parameters and physico-chemical properties were examined, in detail. The unsymmetrical dimethyl hydrazine (UDMH) based zwitterions, UDMH-borane and UDMH-bisborane displayed strikingly lower ignition delay times of 2.2 and 2 ms, respectively and high specific impulse 241 and 245 s, respectively. The pyridinium-cyanoborohydride exhibited high heat of formation (146 kJ/mol) than UDMH. While all the developed compounds exhibited higher density than the UDMH. Chemical stability of UDMH-borane was investigated at different environmental conditions; it was found to be stable for longer duration. Low ignition delay, good thermal and chemical stability, positive heat of formation, high specific impulse, ease of synthesis, and commercially available starting materials are the promising features of these compounds which make them attractive prospects in revolutionizing the area of green hypergolic fuel in liquid as well as hybrid rocket propulsion. [ABSTRACT FROM AUTHOR]

In this study, a series of full-scale high-pressure natural gas pipeline physical explosion experiments were performed, in which the overpressure of the shock waves was measured and analyzed. The concept of "physical explosion linear energy (PEXLE)" was proposed and a "PEXLE model" was established to evaluate shock waves in a cylindrical pipeline with high length–diameter ratio. Equations were fitted to describe the characteristics of peak overpressure, positive pressure duration and specific impulse. Results showed that peak overpressure of shock waves is not influenced by the length of pipeline. With the increase of the propagation distance, specific impulse first increases and then decreases. This indicates that the position of the maximum value of overpressure does not correspond with that of the specific impulse. • Full-scale high-pressure pipeline explosion experiments are carried out. • The "PEXLE model" is proposed for calculating shock waves. • The shock wave peak overpressure is not influenced by the length of pipeline. • The formula of overpressure, positive pressure duration, specific impulse are fitted. • Specific impulse rise first and then fall with the increase of propagation distance. [ABSTRACT FROM AUTHOR]

To investigate the effects of oxidizer injection on the performance of hybrid rocket motors (HRMs), we have designed, manufactured and tested four types of injector: showerhead (SH), hollow-cone (HC), pressure-swirl (PSW) and vortex (VOR). This study is motivated by the fact that the experimental measurements of N 2 O/paraffin firings are poorly presented in the open literature. Besides few publications are dedicated to the characterization of novel types of injectors in hybrid rocket propulsion application such as HC, PSW, and VOR. It is advantageous that the study was conducted in the same motor configuration, with the advantage that it allowed to compare the performance of different types of injectors. This paper analyzes the influence of the oxidizer injector design on the main performance parameters, such as fuel regression rate, specific impulse and combustion efficiency. First, in order to observe injector spray qualities, a series of cold tests using liquid water and liquid nitrous oxide are carried out, providing a good understanding of the spray profiles. Then, the motor performance data is obtained by a series of firing tests using N 2 O as oxidizer and paraffin as fuel. Comparison of the various injectors data is made with the same average oxidizer mass flux and feeding pressure. The showerhead is used as a benchmark in this study. During this experimental analysis, the VOR injector exhibits the highest regression rate, followed by HC and SH. Because the assessment of the regression rate was not enough to explain all the effects of the injectors on the motor performance, firing tests of small-scale hybrid motor have been carried out. In terms of spray properties, the PSW exhibits significant differences compared with the other injectors; it generates the smallest Sauter Mean Diameter (SMD) in the formed spray and achieves good atomization. In spite of the fact that the PSW injector leads to the lowest regression rate, it provides good specific impulse, increases the oxidizer-to-fuel ratio (O / F), as well as a uniform and smooth consumption of the paraffin fuel grain. VOR leads to the highest specific impulse. In terms of stability, VOR, HC and SH injectors exhibit lower oscillations in the chamber pressure. Some observations are made on exhaust plume intensity developed during combustion, and in firing tests with SH a blow-out phenomenon occur often. Similarly to liquid engines, it is possible in hybrid motors to increase the global propulsive performance using alternative designs of the injection system. [ABSTRACT FROM AUTHOR]

On the basis of the framework of cubic gauche nitrogen (cg-N), six one-eighth methanetriyl groups (>CH-) substitutes and fifteen one-fourth >CH- substitutes were optimized using the first-principle calculations based on density functional theory (DFT). Both one-eighth and one-fourth substitutes still keep the gauche structures with the simple formula CHN 7 and CHN 3 , respectively. The most thermodynamic stable gauche CHN 7 and CHN 3 are P 2 1 qtg-C 2 H 2 N 14 I and P 2 1 qtg-C 4 H 4 N 12 III, respectively. No probability density of C-C single bonds and high probability densities of C-N-C structures were found in the two substitutes. Although gauche CHN 7 and CHN 3 lose energy density in contrast to cg-N, they win kinetic stability and combustion temperature (T c ). Thus, they are more feasible than cg-N, and more effective than the traditional rocket fuels. [ABSTRACT FROM AUTHOR]

Based on density functional theory and volume-based thermodynamics methods, the crystal densities ( ρ ), heats of formation (HOFs), detonation performance, specific impulse ( I sp ), impact sensitivities ( H 50 ) and Gibbs free energies of formation of eight series novel energetic salts composed of 4-nitro-7-(4-nitro-1,2,3-triazol-1-olate)-furazano[3,4- d ]pyridazine-based anions and ammonium-based cations were studied. Results show that all title salts possess high ρ and positive HOFs. Therein, ammonium and hydroxylammonium salts exhibit the highest ρ and detonation performance in each series and several even surpass those of HMX and RDX. For guanidinium-based salts in every series, when the number of NH 2 group in cations increases, the HOFs, I sp and H 50 of corresponding salts improve, but their ρ values decrease. Consequently, detonation performance of guanidinium-based salts are close to each other (H series are similar to RDX). Otherwise, introducing N → O oxidation bond to anions is an effective method to improve ρ , detonation performance and I sp of the corresponding salts compared A with B-H series, but it decreases H 50 . However, all guanidinium-based salts show lower impact sensitivities than RDX and HMX. Meanwhile, the position of N → O oxidation bond also has an effect on these properties. [ABSTRACT FROM AUTHOR]

Turbocharged Solid Propellant Ramjet is a new concept engine that combines the advantages of both solid rocket ramjet and Air Turbo Rocket, with a wide operation envelope and high performance. There are three streams of the air, turbine-driving gas and augment gas to mix and combust in the afterburner, and the coaxial intake mode of the afterburner is disadvantageous to the mixing and combustion. Therefore, it is necessary to carry out mixing and combustion enhancement research. In this study, the numerical model of Turbocharged Solid Propellant Ramjet three-dimensional combustion flow field is established, and the numerical simulation of the mixing and combustion enhancement scheme is conducted from the aspects of head region intake mode to injection method in afterburner. The results show that by driving the compressed air to deflect inward and the turbine-driving gas to maintain strong rotation, radial and tangential momentum exchange of the two streams can be enhanced, thereby improving the efficiency of mixing and combustion in the afterburner. The method of injecting augment gas in the transverse direction and making sure the injection location is as close as possible to the head region is beneficial to improve the combustion efficiency. The outer combustion flow field of the afterburner is an oxidizer-rich environment, while the inner is a fuel-rich environment. To improve the efficiency of mixing and combustion, it is necessary to control the injection velocity of the augment gas to keep it in the oxygen-rich zone of the outer region. The numerical simulation for different flight conditions shows that the optimal mixing and combustion enhancement scheme can obtain high combustion efficiency and have excellent applicability in a wide working range. [ABSTRACT FROM AUTHOR]

One of the main goals for the scientific/ military rocket industry is to increase the operational burning time, thus the specific impulse. New homogeneous and heterogeneous propellants were tested and metallic fuels were added in the mixture to obtain best performance. To study heterogeneous propellant, containing large amount of fine boron and ammonium perchlorate, it is appropriate to estimate the combustion products to evaluate/obtain the values of the specific impulse, density, Mach number and mass flow of the mixture. Several composite propellant mixtures, ammonium perchlorate, nitramides (RDX – Cyclotrimethylene trinitramide), were defined with or without addiction of small particles of Boron and modeled. The energetic properties of boron and progress of boron particles on the burning surface of the ammonium perchlorate based composite propellants was modeled used a numerical algorithm. This paper reports the analysis of the influence of boron in the performance parameters for ammonium perchlorate based composite propellants. [ABSTRACT FROM AUTHOR]

The demanding of a holistic view for the engines in the fuel direct precooled family was clarified. A unified cycle model was proposed so as to provide a consistent thread of insight for the engine family. The model indicates that the engines in the fuel direct precooled family are not collections of different cycles, but essentially the same cycle operating under different conditions. Theoretical analysis of the unified cycle was carried out to make clear the dependencies of the engine performance parameters. Numerical model was developed also to take into account the actual conditions including the real fluid properties, the mass addition effects, etc. Performance evaluation and optimization based on the numerical model were completed for three typical cycles under the flight condition of Mach 5.0 with the dynamic pressure q 0 = 45 kPa. The results show that the PC-RT cycle possesses superiority on both the geometry and performance sides among the cycles evaluated. The selection of precooling temperature is a tradeoff between the specific thrust and the specific impulse. Compared to the conventional ramjet engine, the precooled cycles have the potential to provide both high specific thrust and high specific impulse. [ABSTRACT FROM AUTHOR]

[Display omitted] • 4,4′-Bis(trinitromethyl)-3,3′-azo/azoxy-furazan was synthesized as high-energy dense oxidizers. • Two oxidizers exhibit high densities of 1.865 and 1.887 g cm−3, respectively. • Two oxidizers have positive oxygen balances (OB CO 2) of +6.9 % and +10.0 %. • The thermal decomposition temperature of oxidizer 8 is higher than 150 °C. Finding a chlorine-free, environmentally friendly, high-energy dense oxidizer with a high heat of formation as an alternative to ammonium perchlorate (AP) has remained an extremely challenging task in the field of energetic materials. In the present work, two organic oxidizers, 4,4′-bis(trinitromethyl)-3,3′-azofurazan (8) and 4,4′-bis(trinitromethyl)-3,3′-azoxyfurazan (10), were designed and synthesized via a two-step reaction from 3-amino-4-(carboxymethyl)furazan. Their structural features were investigated by single-crystal X-ray diffraction, scanning electron microscopy, nuclear magnetic resonance, infrared spectroscopy, elemental analysis, and several theoretical techniques. Both of the newly synthesized organic oxidizers (8 and 10) exhibit positive oxygen balances, high densities, poor solubility in water, decent thermal stabilities, acceptable mechanical sensitivities, extremely high heats of formation, and excellent detonation performances. The best specific impulses are obtained for the environmentally friendly oxidizers 8 and 10 with calculated values of 273.9 s and 275.9 s, respectively, in a composite propellant consisting of 77 % oxidizer, 9 % aluminium, and 14 % binder. Both of these exceed the specific impulse (263 s) of the currently widely used oxidizer AP in the similar composition. Oxidizers 8 and 10 are expected to have promising applications as high-energy dense oxidizers for various scientific, space, and military projects. [ABSTRACT FROM AUTHOR]

FOX-7 was reacted with hydrazine, hydroxylamine, ammonia, aminoguanidine and metal hydroxides (MOH) in aqueous medium to form nitrogen-rich green energetic salts with high density, high positive enthalpy of formation (ΔH f °), and excellent propulsive and detonation properties. [Display omitted] • A green efficient synthesis strategy enhances energetic properties of FOX-7. • IR, NMR, elemental analysis, single crystal X-ray, DSC – support compounds. • Green energetic materials have high chemical and mechanical stabilities. • Mono- and composite propellant formulations are suitable replacements for AP. • Compounds 1 , 2 , 5 and 6 show excellent detonation performances. New 1,1-diamino-2,2-dinitroethene (FOX-7)-based energetic salts are reported in this study. A variety of bases including hydrazine, hydroxylamine, ammonia, amino guanidine and metal hydroxides (MOH) were found to react with FOX-7 to give the following salts: hydrazinium (E)-2-amino-2-hydrazinylidene-1,1-dinitroethan-1-ide (1) dihydroxylammonium mono((Z)-2-amino-2-(hydroxyimino)-1,1-dinitroethan-1-ide) (2) ammonium (1-amino-2,2-dinitrovinyl)amide (3), potassium/sodium/lithium(I)-[(Z)-2-amino-2-hydrazineylidene-1,1-dinitroethan-1-ide] (4 – 6), and aminoguanidine-(Z)-2-amino-2-hydrazinylidene-1,1-dinitroethan-1-ide (7). The title compounds were fully characterizedusing IR, NMR, and elemental analysis. The structures of the salts 2 and 7 were further confirmed by single crystal X-ray analysis. These salts have experimental densities in the range between 1.59 gcm−3 to 1.92 g cm−3. Thermal stabilities of these salts were measured by DSC and TGA analysis and their decomposition temperatures were found between 85 °C and 177 °C. Enthalpies of formation (ΔH f °) were calculated (Gaussian 03) and with their experimental densities, the propulsive and detonation properties of the nitrogen-rich green energetic salts were evaluated. These compounds exhibit high density, high positive enthalpy (ΔH f °), balanced nitrogen and oxygen contents with excellent propulsive (Isp) and good detonation properties (DP and DV) which are superior to that of RDX, ammonium perchlorate (AP), and comparable to TKX-50. The energetic salts (1 , 2, 5 and 6) show potential as highly promising secondary explosives. These energetic salts have potential as promising solid propellants. [ABSTRACT FROM AUTHOR]

Advanced liquid rocket engines (LREs) are being created by Yuzhnoye Design Office of Ukraine based on the fifty-year experience of rocket engines' and propulsion systems' development. These LREs use both hypergolic (NTO+UDMH) and cryogenic (liquid oxygen+kerosene) propellants. First stage engines have a range of thrust from 40 to 250 t, while the upper stage (used in space) engines - from several kilograms to 50 t and a re-ignition feature. The engines are intended for both Ukraine”s independent access to space and international market. [ABSTRACT FROM AUTHOR]

While polynitrogens are proposed to be the hyper rocket fuels, many of their properties in the solid states are unclear. Herein, we comprehensively investigated the thermodynamic and kinetic stabilities, polymorphs, and rocket performances of platonic and cubic gauche nitrogens, i.e. tetrahedral N 4 , cubic N 8 , platonic dodecahedral N 20 and cubic gauche nitrogen (cg-N), using the first-principles calculations. The results showed that the dissociation barrier of tetrahedral N 4 in the gas phase (>200 kJ mol −1 ) is high enough to keep stable. It would be a liquid monopropellant under slight external pressure or low temperature. The specific impulse ( I sp ) of tetrahedral N 4 is comparable to that of liquid H 2 /O 2 , but it provides a density of 1.459 g cm −3 with the body centered tetragonal packing. The conversion barrier of cubic N 8 in the gas phase (∼85 kJ mol −1 ) is marginal for stability. It may be a solid at ambient temperature, and its I sp exceeds that of liquid H 2 /O 2 . The crystal density of cubic N 8 is 1.964 g cm −3 with the rhombohedral centered hexagonal packing. The dissociation barrier for stability of platonic N 20 in the gas phase (∼25 kJ mol −1 ) is insufficient, however it possesses a lattice energy above 100 kJ mol −1 . Although the I sp of platonic N 20 is as powerful as that of the liquid H 2 /O 2 , it has a crystal density of 2.132 g cm −3 with the face centered cubic packing. For surface relaxation, cg-N is unstable at the ambient condition with a dissociation barrier of ∼17 kJ mol −1 . While the I sp of cg-N is less than those of platonic nitrogens, its effective impulse ( I ef ) is the best due to an extremely high density of 3.401 g cm −3 . In respect of the thermodynamics, the enthalpies per atom of platonic nitrogens are always higher than those of cg-N and ε-N 2 from 0 to 100 GPa. [ABSTRACT FROM AUTHOR]

This paper is the second paper on the study regarding the feasibility of low-enriched uranium (LEU) fuel in nuclear thermal rocket (NTR) applications. While the first paper dealt primarily with the issue of achieving and demonstrating criticality of the LEU-loaded NTR, this companion paper addresses other important reactor parameters which need to be verified in order to demonstrate the practicability of the LEU-NTR. Following the conclusions of the first paper, this second paper seeks to characterize various neutronic and performance parameters in order to definitely demonstrate the technical feasibility of the LEU-NTR. The parameters characterized and presented in this study are the reactor burn-up behavior including effects due to the build-up of Xe-135, the temperature coefficient of reactivity, the reactivity capability of the control drum system, and a rocket performance comparison with previous NTR systems. Impacts of the transient Xe reactivity change during operation have been characterized though various reactor depletion analyses using the MCNP5.1 and MCNPX codes in combination with the ENDF/B-VII.0 nuclear data library. The necessary level of excess reactivity of the LEU-loaded reactor has been quantified and the reactivity control drum, system has been assessed in term of the reactor shutdown capability. In this work, the reference core design presented in the first paper has been modified in order to improve the operational performance of the LEU-NTR to operate with a 900 s specific impulse and a 110 kN thrust. [ABSTRACT FROM AUTHOR]

This paper aims to study experimentally and numerically the effects of aluminum (Al) and aluminum hydride (AlH 3) additives on the performance of HRMs. A steady-state numerical model of HRM is established, which adopts the 95% hydrogen peroxide (HP) as the oxidizer. A firing test is completed for the certification of the numerical simulation model, the results illustrate that the regression rate of addition of 38% AlH 3 and 20% Al increases by 55.6% compared with pure hydroxyl-terminated polybutadiene (HTPB). A good agreement exists between the simulation regression rate and the experimental data. The deviations of thrust and pressure are less than 3.3% and 3.7%, respectively. Specifically, the error of the regression rate is only 0.3%, indicating that the accuracy of adopting numerical simulation method to predict the characteristics of the HRM is acceptable. Several steady-state simulations based on the different content of AlH 3 and Al are performed to investigate the effects on regression rate and combustion performance of the HRM. Due to the addition of H 2 , the thermal conductivity increases with the decrease of the mass fraction of AlH 3 , which leads to the increase of the regression rate. The regression rate of the fuel with 8% AlH 3 and 50% Al increases by 35.5% than the fuel with 58% AlH 3 , demonstrating that the effect of Al on the regression rate enhancement is more than that of AlH 3. The theoretical specific impulse increases with the mass fraction of AlH 3. However, it is owing that the part of H 2 does not participate in the reaction with O 2 and the oxidizer to fuel ratio (O / F) is far from the optimal O / F , the specific impulse is lower than the theoretical specific impulse. [ABSTRACT FROM AUTHOR]

The objective of work outlined in this paper was to prove experimentally the energy efficiency of Zel'dovich thermodynamic cycle (cycle with detonative combustion of fuel) by direct comparison of thrust performances of liquid-propellant rocket engine (LRE) prototypes operating in continuous-detonation and continuous-combustion modes using gaseous hydrogen as fuel and gaseous oxygen as oxidizer. For this purpose, a test rig and two LRE prototypes were designed and fabricated. It has been proved experimentally that Zel'dovich thermodynamic cycle with continuous-detonation combustion of hydrogen–oxygen mixture in two different LRE prototypes is more energy efficient than the cycle with continuous constant-pressure combustion of the same mixture, other conditions being equal. Thus, at predetermined limitations on supply pressures of fuel components and at a fixed mass flow rate of fuel mixture the specific impulse of the LRE prototypes operating in the continuous-detonation mode appeared to be 6–8% higher than the specific impulse of the same LRE prototypes operating in the continuous-combustion mode. [ABSTRACT FROM AUTHOR]

This paper deals with interior ballistics of hydro-reactive motors. The propellant consists of water as the main oxidizer and the fuel rich pyrotechnic as a solid fuel. The effects of water on the propulsive characteristics of this propellant are investigated. It is shown that raising the amount of water increases the pressure and thrust as well as specific impulse. An experimental study has been conducted in order to approve the feature of the hydro-reactive propulsion systems. A binary mixture of magnesium and sodium nitrate powders with mass fractions of 0.75 and 0.25, respectively, is selected as a solid fuel. Three tests have been performed at different levels of water mass flow rates. The data were reduced from the tests and average values of characteristics were extracted. The results confirm the predicted trend of the effects of water on the behavior of propulsive characteristics. The specific impulse of about 400 s has been achieved at chamber pressure of about 22 bar and water mass ratio of 0.74. It is expected that introducing higher values of water mass fractions increases the specific impulse. The results show that the motor behaves with more stability at this condition. [ABSTRACT FROM AUTHOR]

Hydrogen fueled scramjet is a candidate for use as the engine of the aerospace plane for its high specific impulse. To further improve the engine performance, analysis of combustion mode transition for a hydrogen fueled scramjet engine was investigated in this study. In order to identify the differences between Scram- and Ram-mode cycles in propulsive and economic performances for selection and optimization of operation modes, a thermodynamic cycle analysis was made with a dual-mode scramjet engine. It was found through comparative analysis that the specific impulse of the Ram-mode cycle was superior to that of the Scram-mode cycle in meeting the specific thrust requirement. From the viewpoint of performance optimization, the combustion mode transition in a dual-mode scramjet engine should occur in the range of Mach number 6–7. It was therefore concluded that a dual-mode scramjet engine should be operated as much at the Ram-mode cycle as possible when the flight Mach number is less than 6, and the combustion mode transition between Ram-mode cycle and Scram-mode cycle should take place in the range of flight Mach number 6–7. [ABSTRACT FROM AUTHOR]

For hydrocarbon-fueled TBCC engines, precooling is a crucial technology to solve "thrust gap" by expanding the flight Mach number of the turbojet engine, which can bring a sharp reduction of the specific impulse due to insufficient heat sink of hydrocarbon fuel. A low-temperature endothermic fuel precooling scheme is a great potential way to solve this problem. Heat sink experiments and thermodynamic cycle analysis are used to verify the effectiveness of the low-temperature endothermic fuel precooling scheme in improving specific impulse performance. The experiment results indicated that low-temperature endothermic fuel has a more considerable heat sink at lower temperatures than traditional endothermic hydrocarbon fuel. The results of thermodynamic cycle analysis show that the low-temperature endothermic fuel precooling scheme can effectively improve the specific impulse. Maximum specific impulse can reach above 800s with the low-temperature endothermic fuel precooling scheme, which is below 400s with hydrocarbon fuel precooling at Ma 0 = 3.0. When cracking rate equals 1.0, and precooling efficiency equals 0.85, maximum specific impulse can reach 600s, increased by approximately 1.4 times at Ma 0 = 4.0. The feasibility of precooled turbojet with the low-temperature endothermic fuel precooling scheme was preliminarily verified through this work. • Low-temperature endothermic fuel precooling scheme is proposed to improve specific impulse. • Low-temperature endothermic fuel can crack at a lower temperature and chemical heat sink can be used at Ma 3.0–4.0. • The heat sink of low-temperature endothermic fuel methanol is about twice that of hydrocarbon fuel. • Specific impulse can be improved by 1.0–1.4 times at Ma 3.0–4.0 by saving fuel. [ABSTRACT FROM AUTHOR]

Nanothermites possess much attentions relying on potential probability to the fields of the military and civilian. Whereas, the practical applications for nanothermites have been limited by the high electrostatic discharge (ESD) sensitivity and little gas generation, which is a pronounced problem in the microthruster particularly. In this work, the two typical nanothermites (Al/CuO and Al/MoO 3 , as Al/M X O Y) were assembled with various contents of CL-20 to be defined Al/M X O Y @CL-20. The characteristics of Al/M X O Y @CL-20 were examined by scanning electron microscopy (SEM), energy dispersive spectroscopy (EDS), powder X-ray diffraction (XRD) and differential scanning calorimetry (DSC). The ignition properties of Al/M X O Y @CL-20, such as burning rate, the unrestrained combustion, ignition temperature, delay time, pressurization and propulsion properties, were studied. In addition, the energy required for 50% ignition probability (E 50) was applied to evaluating the sensitivity of Al/M X O Y @CL-20 to ESD. Results revealed that the total heat release for Al/M X O Y @CL-20 was increased on account of CL-20, and CL-20 facilitated the transformation of CuO to Cu 2 O. The burning rates decreased with the increase of CL-20, whereas the burning rate of Al/MoO 3 @10%CL-20 was little higher than that of Al/MoO 3. Al/CuO@10%CL-20 and Al/MoO 3 @30%CL-20 exhibited superior performances respectively for the maximum flame area, the lowest ignition temperature, the highest pressurization rate and specific impulse. The particular significance was that CL-20 played a positive role for Al/M X O Y @CL-20 on the safety to ESD. For instance, the E 50 of Al/CuO@50%CL-20 was 163.84 mJ meanwhile Al/CuO of that was 0.24 mJ. In conclusion, in the light of the superior specific impulse and the excellent safety to ESD, Al/M X O Y @CL-20 (especially Al/CuO@10%CL-20 and Al/MoO 3 @30%CL-20) would possess the remarkable application prospects for practice in the microthruster. [ABSTRACT FROM AUTHOR]

Abstract: The present study proposed a unified framework to simulate multi-physical processes which are crucial for trade-off design of liquid rocket thrust chambers among propulsive performance, regenerative cooling, and pressure budget. As the first part, a turbulent combustion model based on the flamelet approach was developed to effectively incorporate detailed chemistry of high hydrocarbon fuel, turbulent mixing, enthalpy loss, and pressure variations within the nonadiabatic nozzle flow. In order to correctly capture the convective heat transfer and viscous friction in the turbulent boundary layer at the chamber wall, an advanced low-Reynolds number turbulence model is adopted in an axisymmetric compressible RANS solver, which is interactively coupled with a cooling analysis module for the conjugate heat transfer and hydraulics through the regenerative cooling channels. The present method has been applied to an actual regeneratively cooled thrust chamber and compared with measurement of hot-firing tests in terms of specific impulse, characteristic velocity, and thrust coefficient. Based on the numerical results, the effects of additional fuel cooling injection and wall friction on the propulsive parameters are discussed. [Copyright &y& Elsevier]

Abstract: This chapter traverses through nearly 20 years of theoretical developments and applications devoted to investigating the geometric metastability of molecules that are speculative, but if they could indeed be made and if they were to scale up to practical applications, they could revolutionize the applications in defense, space travel, etc., where such materials are in high demand. The underlying theme of this research has been the development of new predictive level theory—theory that is accurate and reliable to use in the absence of experiment. Without such assurance from theory, application to “speculative molecules” that are discussed here would have been without merit. [Copyright &y& Elsevier]

Abstract: A two-dimensional axis symmetrical model for multi-phase turbulent combustion simulation has been developed in a magnesium-based fuel ramjet, the stochastic model is applied to track the motion and transportation of discrete phase in the whole flow field in Lagrangian frame. Firstly, the flow field parameters distribution including temperature, velocity, Mg particle burning rate, main components are obtained to describe the basic operating characteristics, and the thermoacoustic oscillation during the combustion process is predicted due to the observed vortex. Then combined with the developed thermodynamic calculation program, the influencing rules of water/fuel ratio and magnesium-based fuel composition ratio on ramjet operating characteristics are analyzed. The results obtained show that the distribution of two water/fuel ratios exerts a direct impact on the temperature fluctuations in combustor, further influences the combustion stability, as for theoretical specific impulse, a maximal value is observed along with the total water/fuel ratio, while the monotonous trends for thermal efficiency and propulsive efficiency are achieved respectively. Moreover, the magnesium-based fuel composition ratio is a dominant factor to influence the operating characteristics, and it is proved that the high-proportion Mg in fuel is beneficial to specific impulse augment through validation of both simulation results and thermodynamic calculation results. [Copyright &y& Elsevier]

Abstract: To predict cyclic performance and propulsive performance, and further instructs the integral engine design, an ideal thermodynamic cycle model for a magnesium-based fuel ramjet engine employed in underwater vehicles is established. Especially, the magnesium-based fuel grain consisting of most magnesium, small ammonium perchlorate, and some binder is carried on-board, while water as the main oxidizer and working medium is injected into the chamber through different pipelines. Thus, the ideal thermodynamic cycle model is regarded as a combination of Brayton cycle and quasi-Rankine cycle due to phase-transition characteristics of water. Based on the thermodynamic cycle model and thermodynamic calculation model, thermal efficiency, propulsive efficiency, as well as specific impulse are predicted theoretically in the present study. The results show that thermal efficiency is dominantly influenced by water/fuel ratio and fuel composition distribution when fixing navigational depth of ramjet, instead, the chamber pressure has the slight influence especially under a higher pressure environment. Otherwise, thermal efficiency relies on navigational depth rather than the fuel composition distributions while fixing the water/fuel ratio. As a heat engine and thruster, the rules of propulsive efficiency for this water ramjet also provide qualitative guidance for improving the energy conversion between kinetic energy and propulsive energy. Furthermore, toward rules of specific impulse versus water/fuel ratio under different chamber pressure, the occurrence of peak specific impulse of order 5000 N s kg−1 on each curve is observed. [Copyright &y& Elsevier]

Abstract: The chemical rocket propulsion has limitations while applied to space application, with respect to specific impulse, delta-v, volume and stored energy, which can be overcome by electric propulsion system. Among various type of Electric propulsion system, the electromagnetic propulsion has added advantages for many space applications. In this paper an electromagnetic propulsion system is considered known as Magnetoplasmadynamic thrusters (MPD). The MPD thrust works based on principle of Lorentz force.The thruster is modeled based on Paschen''s Law and snow plow model to operate it in pulsed mode to generate instantaneous peak thrust. This model includes the designing of high-current power supply, ignition circuit, and electrodes. Simulations have been performed by considering all the parameters involved in the model. The design parameters are analyzed for optimum value to get maximum thrust force. [Copyright &y& Elsevier]

Abstract: In view of the invalidity of suppression and isolation apparatus for gas explosion, a closed vacuum chamber structure for explosion suppression with a fragile plane was designed on the base of the suction of vacuum. Using methane as combustible gas, a series of experiments on gas explosion were carried out to check the feasibility of the vacuum chamber suppressing explosion by changing methane concentration and geometric structure of the vacuum chamber. When the vacuum chamber was not connected to the tunnel, detonation would happen in the tunnel at methane volume fraction from 9.3% to 11.5%, with flame propagation velocity exceeding 2000m/s, maximum peak value overpressure reaching 0.7MPa, and specific impulse of shock wave running up to 20kPas. When the vacuum chamber with 5/34 of the tunnel volume was connected to the flank of the tunnel, gas explosion of the same concentration would greatly weaken with flame propagation velocity declining to about 200m/s, the quenching distance decreasing to 3/4 of the tunnel length, maximum peak value overpressure running down to 0.1–0.15MPa and specific impulse of shock wave below 0.9kPas. The closer the position accessed to the ignition end, the greater explosion intensity weakened. There was no significant difference between larger section and smaller vacuum chambers in degree of maximum peak value overpressure and specific impulse declining, except that quenching fire effect of the former was superior to the latter. The distance of fire quenching could be improved by increasing the number of the vacuum chambers. [Copyright &y& Elsevier]

Abstract: The present work focuses on the determination of the in orbit performance of the Alsat-1 microsatellite propulsion system. The satellite mass is 90kg, of which 6.2kg is the propulsion system dry mass. The system is a butane propulsion system using low power resistojet thruster with 2.3kg of propellant. The liquefied butane gas was selected due to its higher storage density and safety compared to the other propellants used for microsatellites. The purpose of this paper is the analysis of the firings performed after the launch of the satellite and to evaluate the system specific impulse and thrust level during the system lifetime. A total of 273 firings were performed on the Alsat-1 propulsion system in the period between the end of 2002 and mid 2009, the cumulated firing time is more than 12h 49min. The analysis of all the propulsion telemetry data shows that the system provides a total mission delta V of 25.3m/s which is more than the 10m/s specified for this mission. Furthermore, the mission average specific impulse and thrust are respectively 99.9s and 48.8mN. [Copyright &y& Elsevier]

Abstract: The present study focuses on the combustion of liquid propellants used in engines for space propulsion applications (satellites and launch vehicles). The combustion of the following propellant combinations was studied: hydrogen/oxygen, hydrocarbon fuel/oxygen, hydrocarbon fuel/hydrogen peroxide, hydrazine/nitrogen tetroxide, Mono-methyl hydrazine/nitrogen tetroxide and Unsymmetrical dimethyl hydrazine/nitrogen tetroxide. The purpose of this paper is to determine the combustion flame temperature and the other thermochemical parameters (combustion products politropic parameter and molecular mass) as function of mixture ratio by mass of oxidizer to fuel (O/F). Furthermore, the vacuum specific impulse was also calculated at assumed pressure conditions of propulsion engines to show the effect of mixture ratio of the propellants on the performances. For the determination of the equilibrium composition at assumed temperature, nonlinear systems of equations were solved numerically using Lieberstein’s method. Some results were presented and compared with previous research in this area, the comparison shows good agreement between the results and the difference is less than 5%. As example, three cases of bi-propellant engines were studied using the programs of combustion and frozen fluid flow approximation: Space Shuttle Main Engine (SSME), Zenit second stage engine (RD-120) and Ariane-5 upper stage engine (Aestus). [Copyright &y& Elsevier]

Abstract: Thrust and specific impulse are two critical parameters to estimate the performance of rocket engine. Utilizing liquid kerosene as the fuel, oxygen as oxidizer and nitrogen as purge gas, a series of multi-cycle detonation experiments were conducted to systemically investigate the relationships among the operating frequency, fill fraction and performance parameters of the pulse detonation rocket engine (PDRE). The operating frequency of PDRE was up to 49Hz. The mass flow of liquid kerosene was measured by orifice meter and the mass flow of oxygen was measured based on the method of gas collection. The detonation chamber pressure traces were recorded by dynamic piezoelectric pressure transducers. A dynamic piezoelectric thrust transducer was used to record the instantaneous thrust produced by PDRE. Average thrust and detonative mixture-based specific impulse of PDRE with different operation frequency were obtained by experiments. The experimental results indicate that in the practical operation, the operating frequency caused the change of fill fraction, which resulted in a thrust enhancement. Due to the effect of fill fraction, average thrust did not linearly increase as the frequency rises. Fill fraction has a significant influence on the specific impulse of PDRE. The detonative mixture-based specific impulse presented a second order exponential decay with fill fraction, and was correspondingly increased with the raise of the operating frequency. With the strategy of partial filling in detonation tube, the specific impulse can be remarkably enhanced. [ABSTRACT FROM AUTHOR]

Abstract: In the context of special relativity, we discuss the specific impulse of a rocket whose exhaust jet consists of massive and/or massless particles. This work generalizes previous results and corrects some errors of a recently published paper by U. Walter. (The errors stem from the omission of a Lorentz factor.) [ABSTRACT FROM AUTHOR]

[Display omitted] • The lattice energies for Hg and Ag fulminates were calculated. • The respective specific impulse were calculated, by using absolute chemical hardness. • A series of relationship between thermochemical data were established. In the present work are calculated, for the first time, the lattice energies (kJmol−1) for mercury and silver fulminates (by using the Glasser-Jenkins equations): Hg(CNO) 2 = 1399.1, Ag(CNO), orthorhombic = 389.5 and Ag(CNO), rhombohedral = 305.1 (assuming the following charges: Hg: +2, Ag: +1 and CNO: −1). By using SE-PM6 molecular modelling, the "real" charges were calculated, and the "corrected" U POT values (kJmol−1) are: Hg(CNO) 2 = 777.0, Ag(CNO), orthorhombic = 107.2 and Ag(CNO), rhombohedral = 83.97. By using absolute chemical hardness values, the specific impulse (I s) values are 217.05 s and 178.41 s, for Hg(CNO) 2 and Ag(CNO), respectively. The following empirical equation was derived: D = 10 I s n (where D is the detonation velocity (ms−1, I s is the specific impulse and n is the number of CNO– per formulae). The gas phase explosion enthalpies (kJmol−1) are: Hg(CNO) 2 = -420.6 and Ag(CNO) = -164.4. The sublimation enthalpy for mercury fulminate was calculated as 51.9 kJmol−1. Others relationship between thermochemical data to mercury and silver fulminates were also established. [ABSTRACT FROM AUTHOR]

• Cyclopropanated α- and β-pinane with high yields were synthesized through cyclopropanation. • Cyclopropanated α- and β-pinane show higher specific impulses and densities than reported rocket fuels. • Cyclopropanated α- and β-pinane with high-energy-density can be used as advanced rocket fuels. With the rapid development of aerospace industry, it is significantly necessary to develop advanced rocket fuels with high density, high specific impulse and good combustion performance. In this work, we report the synthesis of cyclopropanated pinanes through cyclopropanation of pinene. The reaction conditions including reaction time, solvent type, halohydrocarbon and molar ratio of zinc carbenoid to pinene were optimized. The basic physical properties of fuel, including density, kinematic viscosity, freezing point, calorific value, and combustion performance were tested, and the specific impulses of cyclopropanated pinanes were also calculated. Importantly, the cyclopropanated α- and β-pinane possess high density of 0.90 and 0.91 g/mL, specific impulse of 330.58 and 330.39 s, which results in higher volumetric specific impulse than reported rocket fuels. They also show good low-temperature performance and better ignition and combustion performance than JP-10. These excellent performances indicate that the cyclopropanated α- and β-pinane have great application prospect as high-energy rocket fuel. [ABSTRACT FROM AUTHOR]

Abstract: Pulsed plasma thruster (PPT) is one of the most promising electric propulsions system. Its main drawback is the low thrust efficiency. A theoretical analysis is given in this paper for exploring the reasons. The low electrothermal acceleration efficiency is concluded as an important factor. For improving the electrothermal acceleration, a new electrode structure associated with a nozzle was designed and adopted. The experimental result shows that the new electrode impaired the electromagnetic acceleration slightly. But the overall thrust was improved significantly. It means that the electrothermal acceleration was enhanced distinctly. The thrust efficiency was also improved remarkably. The simple electrode structure with a nozzle implicates an efficient way to improve the performance of PPT. [Copyright &y& Elsevier]

Abstract: Pulsed plasma thruster (PPT) is an attractive micro-thruster for many flight missions. Compared to other electric propulsion systems, PPT is more compact, lightweight and robust while lower in thrust efficiency. It is very difficult to solve this problem because of lacking of theoretically thorough understanding of the energy transformation. In this paper, an attempt was made to explain the process from a different viewpoint where the acceleration is divided into two stages based on the dominative status. At the beginning of discharge, the electrothermal acceleration is primarily near the propellant surface. In the second stage, ions and electrons are accelerated by the electromagnetic field induced by the discharge between the electrodes. Electromagnetic acceleration contributes substantially to the thrust. Some measures should be taken to ameliorate the electrothermal acceleration at the first stage and enhance the electromagnetic acceleration at the second stage. For the purpose of improving thrust efficiency, the relations between the acceleration and electrode gap, electrode width, discharge energy have been discussed. Some methods by modifying design parameters are presented to enhance the accelerating effect. [Copyright &y& Elsevier]

The micro laser plasma thruster (μLPT) is a new micropropulsion device which is competitive with the pulsed plasma thruster (μPPT) for steering or propelling 10–100 kg ‘microsatellites’. It is driven by a cluster of fiber-coupled infrared diode lasers, focused onto a two-layer ablating tape. Electrical to optical conversion efficiency is 50%. The tape consists of a transparent supporting layer through which the light passes without perforating it, and an absorbing fuel layer, which is heated by the laser to form a jet on the side of the tape opposite to the laser. The μLPT operates on spacecraft bus voltage, and weighs 0.85 kg. At the millisecond pulse durations required to enable the diodes to achieve ablation threshold in a 25-μm focal spot, target materials are limited to those of low thermal conductivity, e.g. polymers. Passive ablatants produced 7.5 dyn thrust and a coupling coefficient (Cm) of 8 dyn/W. Specially-designed, carbon-doped exothermic photopolymers produced up to 68 dyn thrust with 2.1 W average optical power, achieving an exhaust velocity of 2–3 km/s in continuous tests and up to 12 km/s in single-shot static tests. We discuss the design and performance of these photopolymers, recent diagnostics of μLPT plasma plume and latest performance results. [Copyright &y& Elsevier]

The closed-electron-drift Hall-effect accelerator is a promising thruster in space. The effects of magnetic field shape and strength on Hall thruster performance were studied using 1-kW-class THT-series Hall thrusters. The discharge current characteristics were sensitive to magnetic field shape and strength. The thrust for the THT-IIIA thruster with more intensive concentration of magnetic field near the acceleration channel exit, i.e., with a smaller magnetic field at the anode was slightly higher than that for the THT-III thruster. The THT-IIIA thruster could be stably operated at a higher magnetic field strength than that for the THT-III thruster. As a result, for the THT-IIIA thruster, a higher thrust efficiency was achieved with a lower discharge current and a higher thrust. Furthermore, Hall thruster flowfield analysis was made to understand inner physical phenomena. Both ionization and acceleration were found to efficiently occur in a thin region with a few mm thick near the acceleration channel exit. The calculated thruster performance roughly agreed with the experimental ones. [Copyright &y& Elsevier]

• The influences of droplet diameter and total equivalence ratio are studied. • Droplet distributions are significantly affected by their evaporation behaviors. • Mixed premixed and non-premixed combustion modes are seen in two-phase RDE. • Specific impulse varies non-monotonically with the droplet diameter. Eulerian - Lagrangian simulations are conducted for two-dimensional Rotating Detonative Combustion (RDC) fueled by partially prevaporized n -heptane dilute sprays by using coarse mesh resolution. The air is used as the oxidizer in the present simulations. The influences of droplet diameter and total equivalence ratio on detonation combustion and droplet dynamics are studied. It is found that small n -heptane droplets (e.g. 5 µm) are completely vaporized around the detonation wave, while intermediate n -heptane droplets (e.g. 20 µm) are consumed in or behind the detonation wave, with the escaped ones being continuously evaporated and deflagrated. The droplet distributions in the rotating detonation combustor are significantly affected by the droplet evaporation behaviors. Both premixed and non-premixed combustion modes are seen in two-phase RDC. The detonated fuel fraction is high when the droplet diameters are small or large, reaching its minimal value with diameter being 20 µm. The detonation propagation speed decreases with increased droplet diameter and is almost constant when the diameter is larger (>30 µm). The velocity deficits are 2 - 18% compared to the respective gaseous cases. Moreover, the propagation speed increases as the total equivalence ratio increases for the same droplet diameter. It is also found that the detonation propagation speed and detonated fuel fraction are considerably affected by the pre-vaporized gas equivalence ratio. The specific impulse first decreases for cases with initial diameter<5 μm, then increases with droplet diameter between 5 μm and 20 μm, and finally decreases with droplet diameter>20 μm. [ABSTRACT FROM AUTHOR]

Seven different high-performance fuels were synthesized from bio-based monoterpenes (3-carene, sabinene, α-pinene, β-pinene, limonene, 1,6-dimethyl-1,5-cyclooctadiene, and myrcene) by Simmons-Smith cyclopropanation. Key fuel properties of the resulting multicyclic fuel mixtures were measured including, density, net heat of combustion (NHOC), and low-temperature viscosity. The mixtures exhibited densities 10–16% higher and volumetric NHOCs 12–16% higher than conventional jet fuel, while maintaining low temperature (−20 °C) viscosities either below or close to the specification limit. To evaluate the fuels as potential rocket propellants, heats of formation were calculated using the G4 composite method. These values were then used to calculate specific impulse (I sp) and density-specific impulse (d-I sp) for each fuel mixture. The cyclopropanated monoterpenes exhibited sea level I sp values up to 1.2% higher than conventional rocket propellant (RP-1) and d-I sp values up to 4.1% higher than RP-1. The results of this study suggest that cyclopropanated monoterpenes may have utility as high-performance jet fuel blendstocks or specialized rocket propellants, offering the ability to increase the payload capacity and enhance the mission capability of space-delivery systems. This work further highlights the potential of biosynthetic fuels for high-performance applications. • Biosynthetic fuel mixtures were prepared by cyclopropanation of terpene substrates. • Fuel properties were evaluated and structure-property relationships determined. • Fuels exhibited volumetric heats of combustion up to 16% higher than Jet-A. • ΔH f and specific impulse (I sp) values were calculated for the fuel molecules. • Fuels had d-I sp values up to 4.1% higher compared to RP-1. [ABSTRACT FROM AUTHOR]

Nanothermites are considered promising energetic materials and a key to develop new technologies in the field of energetics and propulsion. However, relatively low peak pressure, smaller thrust output, and incomplete combustion represent some of the potential problems associated with nanothermites. In this study to improve the combustion performance of nanothermites, varying quantities of nitrocellulose (NC) were introduced to produce the quaternary NC/GO/Al/KClO 4 nanothermite using facile electrospinning. The morphology of nanothermites was characterized by SEM-EDX which confirmed that the nanoparticles were homogeneously dispersed without agglomeration. The combustion behavior of the composites was evaluated by igniting the samples at different packing densities (%TMD) in open acrylic tubes within a reaction chamber and using a 3.5 W continuous wave laser. High speed imaging captured the flame propagation and differential scanning calorimetry (DSC) was used to assess thermal behavior and energy output. Finally, for propulsion characterization, a small-scale test motor (STM) with a converging/diverging nozzle was used to evaluate the combustion performance of the prepared samples. In general, thrust output, specific and volumetric impulses (I SP and I SV), and total heat released exhibit remarkable enhancement with the addition of NC. Specifically, the total impulse (I FT) and I SP peaked at 5% NC/GO/Al/KClO 4 at ~ 50% TMD, with I FT of 19.9 mN.s and I SP of 203.2 s. These are improvements of over 50% compared to the sample without NC (13.4 mN.s and 137.4 s). The ignition delay time and the power required to ignite the NC-enriched mixture is increased, but a sufficiently fast response is maintained for practical applications. Thermal analysis implies that addition of NC introduced another step to the GO/Al/KClO 4 reaction: a gas− solid phase and liquid−liquid phase diffusion reaction, with the liquid−liquid phase reaction increasing with NC percentage. These results suggest that electrospinning is a simple technique for preparation of quaternary nanothermites that results in enhanced combustion performance and an improved suitability for small-scale propulsion applications. [ABSTRACT FROM AUTHOR]

To enhance the thrust performance of laser-pulsed plasma thruster (LPPT), we proposed to use chemical propellant as substitute for ordinary PTFE propellant. The discharge characteristics and thrust performance of chemical propellant were analyzed. Meanwhile, the contribution of different energy sources and the physiochemical process between laser and chemical propellant were investigated. The results showed that the chemical propellant had higher plasma peak current and shorter discharge delay time compared to PTFE propellant. For chemical propellant, the contribution of laser energy was about 3%, while the electric energy and chemical energy was less than 17% and more than 80%, respectively. Based on the self-made torsion pendulum micro-impulse test bench, we found the chemical propellant had the highest impulse bit of 27.9 μN·s at 0 V and 69.9 μN·s at 1500 V. Meanwhile, the PTFE propellant had the lowest laser ablation mass, which is 3.70 μg/shot. The chemical reaction between laser and chemical propellant enhances the impulse bit, and increases the mass consumption of the propellant. The chemical propellant had the highest specific impulse, which is 295.1 s at 0 V and 738.5 s at 1500 V. These findings can indicate that the chemical propellant is a better choice than PTFE propellant. • The technology of electric propulsion and laser propulsion were combined. • The chemical propellant was used to replace traditional PTFE propellant. • The contribution of different energy sources was analyzed. • The physiochemical process between laser and chemical propellant was investigated. • The chemical propellant was proven to be the better choice than PTFE propellant. [ABSTRACT FROM AUTHOR]

• Energetic composite with CL-20 and AP confined in carbonized butterfly wing scales (CWs) matrix has been designed to achieve laser ignited microcombusion. • High-efficiency, stable and self-sustaining microcombustion of the CWs/CL-20/AP energetic composite has been realized due to the enhanced heat and mass transfer. • Multi-channeled hierarchical porous structure and synergistic reactions are responsible for the excellent microcombustion performance. • Composite with oxygen balance close to zero has the best mircocombustion performance. To meet the requirement of miniaturization and integration of the micro thrusters based on the chemical fuels, high-energy density compound is an emerging material as the chemical fuel. However, it is still challenging to realize the high-efficiency, stable and self-sustaining microcombustion of the high-energy density chemical fuels. Herein, we report the high-performance microcombustion of CWs/AP/CL-20 composite with nanoscale 2,4,6,8,10,12-(hexanitrohexaaza)cyclododecane (CL-20) and oxidizer ammonium perchlorate (AP) confined in a self-standing carbonized butterfly wing scales (CWs) matrix. CWs with multi-channeled hierarchical porous structure provide a model architecture for loading of AP and CL-20. CL-20 is designed to serve as heat/pressure loss compensate agent due to its high-energy density, fast energy release rate and large amounts of gas release during combustion process. The combustion of CWs/AP/CL-20 is greatly enhanced due to the enhanced heat and mass transfer originating from the multi-channeled hierarchical porous structure and synergistic reactions between components at nanoscales, which is demonstrated as promising solid-propellant for the micro thruster. It is found that the sample with oxygen balance close to zero, which can be conveniently regulated by adjusting the mass ratio of CL-20 and AP, has the best microcombustion performance, resulting in a specific impulse of 383.5 N·s/kg. [Display omitted] [ABSTRACT FROM AUTHOR]