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Hydrazine and its derivatives have served as a conventional bipropellant fuel for several decades. However, their extremely acute toxicity and carcinogenicity, high volatility, and environmental impact have attracted significant attention. In order to synthesize green bipropellant fuels with high density, high specific impulse, and good thermal stability, three novel N -alkyl-1,2,4-triazole–borane complexes were successfully synthesized by reacting alkylated 1,2,4-triazole coordinated with sodium borohydride in the presence of ammonium sulfate. During the droplet test with white fuming nitric acid, there was a relatively short ignition delay time of 98 ms. Additionally, these hypergolic fuels possessed a high density exceeding 1.10 g cm–3 , and the specific impulse is ranging from 187 to 199 s, and the highest decomposition temperature reaches 153.4 °C. These results demonstrate their great potential as hypergolic fuels or hypergolic ionic liquid additives in the field of hypergolic materials. [ABSTRACT FROM AUTHOR]

Highly reactive Al/${\rm{B}}{{\rm{i}}_2}{{\rm{O}}_3}$Bi2O3 nanothermite reinforced with MLG (multi-layer graphene) is prepared using physical mixing technique, and its exothermic energy release and specific impulse are studied for small-scale propulsion applications using differential scanning calorimeter and force sensor, respectively. The chemical composition of nanothermite samples is examined using X-Ray Diffraction (XRD). Field emission scanning electron microscopy (FESEM) is employed to analyze the surface morphology of prepared samples. According to DSC data, the energy release in Al/Bi2O3 nanothermite increases with increasing content of MLG nanoplatelets up to three weight percent of MLG in nanothermite composition. The maximum value of observed energy release is 668 J/g for nanothermite with three weight percent of MLG. Similar outcomes are observed for the specific impulse values of Al/MLG/Bi2O3. The maximum value of specific impulse observed for Al/Bi2O3 nanothermite with two weight percent of MLG is 50.8 s. [ABSTRACT FROM AUTHOR]

To meet the demand for air-breathing power for wide-range vehicles at Mach 0–10, two thermal cycles with ammonia as the fuel and coolant were analyzed, namely the precooled rocket-turbine cycle (PC-RT) and the precooled gas-turbine cycle. Firstly, the operating modes of the precooled cycle engines were divided into turbine mode, precooling mode, and ramjet mode. Secondly, a fluid-structure coupling heat transfer program was used to evaluate the cooling effects of different fuels on the incoming high-temperature air. The result shows that the equivalent heat sink of ammonia is higher than that of other fuels and can meet the cooling requirement of at least Mach 4 in the precooling mode. Thirdly, the performance of the PC-RT in the turbine and precooling modes was compared at Mach 2.5. The result shows that air precooling alleviates the restriction of the pumping pressure on the minimum required β and improves the specific thrust within a reasonable range of β. Fourthly, the performance of the precooled cycle engines was compared when using different fuels. The result shows that the specific thrust of ammonia is greater than that of other fuels, and the performance advantages of ammonia are the most obvious in the precooling mode due to its highest equivalent heat sink. To sum up, the precooled cycle engines with ammonia as the fuel and coolant presented in this study have the advantages of no carbon emissions, low cost, high specific thrust, and no clogging of the cooling channels by cracking products. They are suitable for applications such as the first-stage power of the two-stage vehicle, and high Mach numbers air-breathing flight. [ABSTRACT FROM AUTHOR]

One of the ways to improve specific characteristics of the power plant of a launch vehicle for ladeploying payload to the near-Earth space is to provide the possibility of operation of a fixed nozzle in the design mode over the whole active leg of the flight trajectory. The nozzle should be compact, lightweight, well-cooled. For detailed testing of the possibility of introducing a nozzle into the rocket engine chamber it is necessary to be able to quickly assess the true value of the thrust and the specific impulse the chamber with such a nozzle can achieve. This article presents the results of comparison of the thrust and specific impulse, obtained during calculations using engineering methods, numerical modeling for the atmospheric section and high-altitude sections of the trajectory. The results of the calculation are compared with the experimental values of the specific impulse obtained on the rocket engine test-bed under atmospheric operating conditions. These results can be effectively applied both to evaluate new and to improve existing nozzle designs of wide-range rocket engines.

The energy potential of hypothetical zwitterionic nitrohydrazine H3N N NO2 as a component of solid composite propellants is estimated. Two-component formulations with a hydrocarbon or an active binder and three-component formulations with aluminum or aluminum hydride additives are considered. The highest calculated ballistic efficiency was obtained for formulations with an active binder and aluminum hydride. Their specific impulse exceeds 280 s, and with allowance for two-phase losses, the maximum effective impulse of the third stage rocket motor reaches 265.4 s, whereas for the optimized similar formulation based on ammonium dinitramide, this value is much lower (262.5 s). [ABSTRACT FROM AUTHOR]

To identify promising areas for the search for high-energy materials (HEMs), there is an urgent need for a comprehensive analysis of the energy potential of compounds of various classes. This paper studies the energy potential of some organic compounds containing the –N=N(O)–C(NO2)3 fragment in their structure as plasticizers of a polymeric binder in solid composite propellants. Nine trinitromethyl-ONN-azoxy-derivatives of furazan and one similar methane compound are studied, four of which are actually synthesized substances, the rest are still hypothetical structures. The ballistic efficiency of solid composite propellants of three different types (without metal, with aluminum, and with aluminum hydride (AH)) in which one of the studied compounds with a trinitromethyl-ONN-azoxy fragment acts as a plasticizer of the polymer binder is assessed. The values of its enthalpy of formation and density are determined by calculation. A comparative analysis of the ballistic efficiency of such propellants with similar compositions containing the most powerful of the currently considered energy-intensive components (nitroglycerin, tetranitromethane, or dinitrofurazan) as a plasticizer shows that practically all the studied representatives of the class of trinitromethyl-ONN-diazene oxides are significantly superior in terms of ballistic efficiency to the reference plasticizers. [ABSTRACT FROM AUTHOR]

Keeping in view the weightage of electric propulsion over chemical propulsion, materials from metals to polymers and liquid (water) have been tested as propellants in ablative laser propulsion. This emerging propulsion technique can be widely used for aerospace applications like debris removal in the range of cm, pointing micro and nano satellites, laser micro thrusters for spacecraft attitude, and orbit control. Laser propulsion can become a less expensive alternative to chemical propulsion. In this review, we compile the work done in ablative laser propulsion and different modes of propulsion along with the efficiency of different propellants. We summarize the optimized propulsive parameters with solid propellants and liquid propellants along with the efficiencies and theories of laser thrusters with optimized specific impulses. The article provides precise developments done in the field of ablative laser propulsion and deep insights into the analysis done between the different propellants used recently in ablative laser propulsion. [ABSTRACT FROM AUTHOR]

Hybrid rocket motors have great development potential due to their outstanding thrust adjustment flexibility and long-term operation ability. However, nozzle erosion during the motor operation can cause an increase in the throat area of the nozzle, which leads to a decrease in combustion chamber pressure and nozzle efficiency. Therefore, a performance prediction model for hybrid rocket motors considering nozzle erosion has become a key technology that must be addressed when developing hybrid rocket motors. This study uses dynamic grid technology to simulate the regression of the combustion surface and nozzle erosion, which fits well with experimental values. The behavior of high-energy particles in the combustion chamber is simulated through a discrete phase model. Notably, distinctive behavior is observed in Al and Mg droplets, with Al droplets exhibiting incomplete vaporization in the combustion chamber while Mg droplets completely vaporize. A ground firing test using the Φ336 mm hybrid rocket motor lasting 200 s is conducted. The results show that the dynamic numerical simulation, accounting for nozzle erosion, substantially enhances performance prediction accuracy. The average deviation in motor thrust remains below 1.8%, and the combustion chamber pressure deviation stays under 2.6%, confirming the precision of the model. Ultimately, both simulation and experimental outcomes indicate a gradual decrease in specific impulse and characteristic velocity over the long-term operation, attributed to the gradual deviation of the oxygen-fuel ratio. This research provides valuable insights for guiding hybrid rocket motor design and optimizing design parameters to improve overall performance. This model can achieve long-duration and high-precision performance predictions for hybrid rocket motors. [ABSTRACT FROM AUTHOR]

The object of this study is the ST-25 Hall thruster with limited discharge power, no more than 200 W, designed by Flight Control LLC (Ukraine). The problem that was solved in the current paper was to determine the effect of residual gas pressure in vacuum chambers on the operating parameters of the Hall thruster. To solve this task, the operating parameters of the ST-25 thruster were determined, which was tested in three vacuum chambers with different sizes of residual pressure. As a result of the laboratory study into the operating parameters of the ST-25 thruster, the volt-ampere characteristics of the engine discharge at fixed values of the working gas (xenon) flow rate were obtained. The dependences of the engine thrust on the mass flow rate of the working gas at fixed values of the discharge voltage were derived. Based on the experimental data, the dependences of the specific impulse of the engine anode unit on the discharge voltage, as well as the dependence of efficiency of the engine anode unit on the discharge voltage were calculated. The studies showed that when the residual pressure in the vacuum chamber is reduced by 2–3 times, the operating parameters of the engine increase by 15–20 %. Such a reduction in residual pressure increases thrust by 25–40 %. Special feature of the results is the determination of threshold values of residual gas pressure in vacuum chambers during experimental studies, in which the operating parameters of the Hall thruster are similar to its operating parameters under space conditions. This work’s findings could be used in practice when conducting experimental studies of electric rocket engines, when it is necessary to estimate the operating parameters of Hall thrusters that will be obtained under actual space conditions. [ABSTRACT FROM AUTHOR]

The NATO agreement STANAG 4569 defines the protection levels for the occupants of logistic and light armored vehicle. The Allied Engineering Publication, AEP-55, Volume 2 document outlines the test conditions for underbelly improvised explosive device (IEDs), which must be buried in water-saturated sandy gravel. The use of sandy gravel has some drawbacks, for instance reproducibility, time consumption, and cost.This paper focuses on the investigation of four alternatives to sandy gravel, which could produce similar specific and cumulative impulses: a concrete pot filled with water, a concrete pot filled with quartz sand, a steel pot without filling and a concrete pot filled with glass spheres (diameter 200 μm–300 μm) and different water contents. The impulses are measured with a ring technology developed at the Fraunhofer EMI.A numerical soil model based on the work of Marrs, 2014 and Fišerová, 2006 and considering the soil moisture was used to simulate the experiments with glass spheres at different water contents, showing much better agreement with the experiments than the classical Laine & Sandvik model, even for high saturation levels. These results can be used to create new test conditions at original scale that are more cost-effective, more reproducible and simpler to manage in comparison to the current tests carried out with STANAG sandy gravel.

Background This study compares propellant fuels’ specific thrust and impulse parameters using nanocarbon variant fuels from coconut shells and coal. Specific impulses and impulses are essential parameters that determine rocket performance. The specific thrust and impulses are influenced by fuel type, material composition, heat flow, and burning time parameters. The characteristics of nanocarbon as a fuel are proven to have high-value features and a long combustion time. This parameter is essential to add value to specific thrusts and impulses. Methods The method to prove the quality of solid fuel material was done experimentally. The composition of the propellant fuel variant mixed with Ammonium Perchlorate (NH 4ClO 4) as an oxidizer through the printing process and sample testing was carried out using scanning electron microscope (SEM), X-ray diffraction analysis (XRD), combustion time, and specific impulse tests. Results Test results with the composition of CBK2 (Coconut Shell + NH 4ClO 4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 267 seconds or an increase of 37 seconds without carbon. CBB2 (Coal + NH 4ClO 4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 261 seconds or an increase of 31 seconds. Conclusions The composition of nanocarbon as a solid fuel propellant has been proven to increase a rocket’s thrust and specific impulses. After all the material variants have been tested for thrust and specific impulse, the best rocket propellant is CBK2.

Copyright of Cyber Security & Data Governance is the property of Editorial Office of Information Technology & Network Security and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

The NATO agreement STANAG 4569 defines the protection levels for the occupants of logistic and light armored vehicle. The Allied Engineering Publication, AEP-55, Volume 2 document outlines the test conditions for underbelly improvised explosive device (IEDs), which must be buried in water-saturated sandy gravel. The use of sandy gravel has some drawbacks, for instance reproducibility, time consumption, and cost. This paper focuses on the investigation of four alternatives to sandy gravel, which could produce similar specific and cumulative impulses: a concrete pot filled with water, a concrete pot filled with quartz sand, a steel pot without filling and a concrete pot filled with glass spheres (diameter 200 mm e300 mm) and different water contents. The impulses are measured with a ring technology developed at the Fraunhofer EMI. A numerical soil model based on the work of Marrs, 2014 and Fiserova, 2006 and considering the soil moisture was used to simulate the experiments with glass spheres at different water contents, showing much better agreement with the experiments than the classical Laine & Sandvik model, even for high saturation levels. These results can be used to create new test conditions at original scale that are more cost-effective, more reproducible and simpler to manage in comparison to the current tests carried out with STANAG sandy gravel. [ABSTRACT FROM AUTHOR]

This paper outlines a methodology for assessing the integral characteristics of a generic scramjet with an integrated propulsion system (PS). The specific impulse and heat flows in the PS for X-43 and X-51 scramjets are calculated. The results obtained are in good agreement with the estimates of other authors (in the speed range 5 < M < 10), as well as with the known graphs of the specific impulse dependence for various types of engines depending on the flight Mach number and the type of fuel used. [ABSTRACT FROM AUTHOR]

Background This study compares propellant fuels’ specific thrust and impulse parameters using nanocarbon variant fuels from coconut shells and coal. Specific impulses and impulses are essential parameters that determine rocket performance. The specific thrust and impulses are influenced by fuel type, material composition, heat flow, and burning time parameters. The characteristics of nanocarbon as a fuel are proven to have high-value features and a long combustion time. This parameter is essential to add value to specific thrusts and impulses. Methods The method to prove the quality of solid fuel material was done experimentally. The composition of the propellant fuel variant mixed with Ammonium Perchlorate (NH4ClO4) as an oxidizer through the printing process and sample testing was carried out using scanning electron microscope (SEM), X-ray diffraction analysis (XRD), combustion time, and specific impulse tests. Results Test results with the composition of CBK2 (Coconut Shell + NH4ClO4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 267 seconds or an increase of 37 seconds without carbon. CBB2 (Coal + NH4ClO4 + Binder + Al = 15: 70: 10: 5) produced a specific impulse of 261 seconds or an increase of 31 seconds. Conclusions The composition of nanocarbon as a solid fuel propellant has been proven to increase a rocket’s thrust and specific impulses. After all the material variants have been tested for thrust and specific impulse, the best rocket propellant is CBK2.

A new method for determining the burning rate of a solid propellant, called the Impulse Method, is proposed in this paper. It is based on the proportional relationship between the impulse generated and the mass of the burned propellant. The pressure–time and thrust–time curves are obtained from a tubular propellant grain burning in the chamber, whose inner surface serves as the initial burning surface. Consequently, the mass of the propellant that was burned off at different pressures can be determined, and the burning rates at different pressures are derived according to the geometric parameters of the propellant grain. The Impulse Method was applied to test the burning rate of two types of propellants twice. The results show that the burning rates were consistent for the same propellant at corresponding pressures, demonstrating the feasibility and reliability of the Impulse Method. The burning rate of a GAP-based composite propellant at 20 MPa measured using the Standard Motor Method was 22.6 mm/s, and that measured using the Impulse Method was 22.2 mm/s and 22.7 mm/s, respectively. These findings indicate that the two methods have comparable accuracy. However, the Impulse Method has the advantage of obtaining the burning rate of the solid propellant at any pressure through a single test. In addition, the nozzle erosion only affected the pressure and not the burning rate. Finally, the rationality of the approach for determining the actual specific impulse was proven by comparing the results with those from another testing method. [ABSTRACT FROM AUTHOR]

Hybrid rocket motors have great development potential due to their outstanding thrust adjustment flexibility and long-term operation ability. However, nozzle erosion during the motor operation can cause an increase in the throat area of the nozzle, which leads to a decrease in combustion chamber pressure and nozzle efficiency. Therefore, a performance prediction model for hybrid rocket motors considering nozzle erosion has become a key technology that must be addressed when developing hybrid rocket motors. This study uses dynamic grid technology to simulate the regression of the combustion surface and nozzle erosion, which fits well with experimental values. The behavior of high-energy particles in the combustion chamber is simulated through a discrete phase model. Notably, distinctive behavior is observed in Al and Mg droplets, with Al droplets exhibiting incomplete vaporization in the combustion chamber while Mg droplets completely vaporize. A ground firing test using the Φ336 mm hybrid rocket motor lasting 200 s is conducted. The results show that the dynamic numerical simulation, accounting for nozzle erosion, substantially enhances performance prediction accuracy. The average deviation in motor thrust remains below 1.8%, and the combustion chamber pressure deviation stays under 2.6%, confirming the precision of the model. Ultimately, both simulation and experimental outcomes indicate a gradual decrease in specific impulse and characteristic velocity over the long-term operation, attributed to the gradual deviation of the oxygen-fuel ratio. This research provides valuable insights for guiding hybrid rocket motor design and optimizing design parameters to improve overall performance. This model can achieve long-duration and high-precision performance predictions for hybrid rocket motors.

A laser can be used to propel distant objects. The ability of silicon to produce thrust and its propulsive parameters are needed to be studied in ablative laser propulsion (ALP). In this work, pure silicon and silicon doped with indium were subjected to ALP to achieve the momentum coupling coefficient and specific impulse to investigate its worth as a propellant. The experiment was conducted using the Nd:YAG laser (Quantel Brilliant) operating at fundamental harmonic (λ = 1064 nm and 5 ns pulse duration). In the given range of fluence, 1 × 105–5 × 105 J/m2, no significant difference among both samples for values of momentum coupling coefficient (Cm) and specific impulse (Isp) is observed; for instance, both propellants follow a decreasing trend for Cm. However, maximum enhancement for Cm and Isp is observed with a cavity aspect ratio of one. Cm and Isp are enhanced about 1.75 and 4.5 times, respectively, for pure silicon. The external cavity does not have any impact on Cm values for indium-doped silicon while values of Isp showed considerable enhancement. [ABSTRACT FROM AUTHOR]

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 binary mixture of dimethyl aminoethyl azide (DMAZ) and tetramethyl ethylene diamine (TMEDA) has been introduced to replace the hydrazine group in space industries. Both fuels are noncarcinogens. The former has a slightly longer ignition delay (ID) time, whereas the latter has low density. Adding DMAZ to TMEDA may have a synergetic effect on the ID time of the binary mixtures TMEDA-DMAZ and better performance effects than each fuel. In this research, the ID time and specific impulse (Isp) of the binary mixtures were assessed with nitric acid-based oxidizers such as pure HNO3, AK14, AK27, and AK44. Using these oxidizers, the binary mixtures containing 50, 40, 50, and 65 wt% of TMEDA in DMAZ were hypergolic with ID times of 22, 18, 14, and 45 ms, respectively. The density-specific impulse of the binary fuel mixtures with the oxidizers was higher than that of unsymmetrical dimethylhydrazine (UDMH). Therefore, the binary fuel mixtures of TMEDA-DMAZ with optimum compositions can be proposed as a good substitution for UDMH in space programs. [ABSTRACT FROM AUTHOR]

Copyright of Industrial Minerals & Processing / Huagong Kuangwu yu Jiagong is the property of Industrial Minerals & Processing Editorial Office and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

What distinguishes the rocket system from airbreathing engines is that the rocket carries its own propellants, fuel and oxidants. The modern rocket was originally developed for ballistic missiles. The operation of the German V2 ballistic missile was not satisfactory because of excessive nose-cone heating, The Soviet Union launched Sputnik 1 to orbit the earth for 21 days. The solid and liquid propellants are widely used in rocket propulsion. Thrust production is necessary in the propulsion of the rocket. Therefore, the purpose of this study is the development of expression through thrust for the study of the performance characteristics of the rocket. Mathematical model has been developed and illustrative solutions obtained, plotted and analysed for the relationship between parameters that affect rocket performance.

A mathematical model of thermodynamic processes occurring in the arcjet thruster device with using argon for arcjet working process carring out. A comparative analysis of the results of experimental data with theoretical data has been carried out. The purpose of the study is to develop a mathematical model of thermodynamic processes in the arcjet thruster device and simulate the interaction of argon with arc discharge plasma. The comparative analysis of the calculated parameters, thanks to the proximity of electron temperature values with the previously experimentally obtained under the static state in the arcjet thruster device, confirms the reliability of the presented mathematical model of plasma processes and the need for its improvement.

Micro-propulsion is considered to be the emerging technology for the propulsion of micro and micro aerospace vehicles as it is preferred over mesoscale thruster due to lower overall life-cycle cost and launching costs. Hence this paper investigates the influence of critical parameters like the Nozzle Pressure Ratio (NPR) and Reynolds number (Re) on the operational characteristics of the micronozzle. A conical nozzle with throat diameter 710 µm and exit/throat area ratio ∼2.14 has been designed and is analyzed numerically by using a model based on pressure-based coupled implicit for various NPR, the backpressure with three Res namely, 1000, 1500, and 2000. The performance of this micronozzle has been characterized in terms of thrust, thrust coefficient, and specific impulse for all three Re cases. A subsequent analysis of the subsonic layer reveals that the nozzle is subjected to high viscous losses at low NPRs, which are independent of Re. [ABSTRACT FROM AUTHOR]

In this paper, an 80-ton thrust liquid rocket engine (hereinafter referred to as an LRE) with a gas generator cycle, a 5:1 thrust throttling ratio, and an integrated flow regulator/gas generator (hereinafter referred to as an IFRGG) is analyzed. This LRE can be used during the first stage of launching, second-stage and upper-stage space missions, and moon/mars low-orbit hovering and soft landing, and it can also be used with various near-space multipurpose flight vehicles. The thermal performance model of the variable-thrust LRE is established, the influence of the main LRE design parameters on the performance optimization is analyzed, and an optimal selection of the design parameters under certain constraints is completed. A performance evaluation was successfully conducted, and we determined the main structural parameters at the sea-level design point; additionally, we evaluated the LRE performance, matched the system parameters under a 100–20% variable-thrust operation, and conducted an analysis on the LRE operating characteristics under a wide range of variable-thrust operations. The LOX/kerosene propellants were selected, and the vacuum-specific impulse of the LRE decreased from 303.2 s to 289.2 s; this followed an approximate linear law, with a decrease of about 4.62% when the thrust was varied in the wide range of 100–20%. The variable-thrust LRE still had a better vacuum performance under a very deep throttling condition. The reason why the specific impulse was low under the deep throttling condition is that it was greatly affected by the different atmospheric pressure that was caused by the varying flight height and the insufficient atomization and combustion of the propellant; however, because of its wide range of variable-thrust working abilities, it is suitable for various special flight missions. [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]

Copyright of Journal of National University of Defense Technology / Guofang Keji Daxue Xuebao is the property of NUDT Press and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

A pre-cooled combined engine based on a closed helium (He) cycle offers high specific thrust and high specific impulse. Therefore, evaluation of the performance of such an engine is crucial for engine applications and key technology research. This study employs an analytical approach to investigate the effects of key parameters on the performance of a pre-cooled core engine, assuming a perfect gas model. The findings revealed that the specific thrust and specific impulse of the pre-cooled core engine are related to the pressurized coefficient of the airflow passage and equivalence ratio (ER). An increase in the pressurized coefficient leads to an increase in both specific thrust and specific impulse. However, within a certain range, although the specific thrust is positively correlated with the ER, the specific impulse is greatly reduced. With specific component parameters and a fixed thermodynamic cycle, a minimum ER exists, which satisfies the cycle-matching requirement. Moreover, the value of the minimum ER is related to the closed-cycle efficiency. For a pre-cooled core engine with a simple closed He cycle, the minimum ER is approximately 2.5–3.5.

A wide range of propellants, and propulsion systems in space exploration by aircrafts or space vehicles was studied, developed, investigated, and commercialized. Liquid, solid, or hybrid propellants have been used for rocket’s launches. In this review, a consistent definition of space propulsion systems, including solid, liquid and hybrid has been given with up-to-date state of developments. A comparison of their performances was made by theoretical and experimental specific impulses. On the other hand, ammonium perchlorate and hydrazine were used as propellants for rocket’s launches and for satellite’s maneuverings; respectively. However, their high toxicity and their storage problem pushed researchers and scientists to investigate and develop other eco-friendly, propellant systems, so called “green propellants”, for launch or reaction control systems of satellites.

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]

As a key component of the solid rocket motor, nozzle ablation is very important to the performance of the motor. In order to study the influence of the law of nozzle ablation on motor performance, a two-dimensional axisymmetric model of a ground state nozzle and vacuum state nozzle before and after ablation was constructed on the basis of a reference profile, and the change law of pressure, temperature and velocity with ablation were studied. The comprehensive performance of the nozzle before and after ablation was calculated to evaluate the influence of ablation on motor performance. It was found that the profile ablation changed the shock wave distribution in the flow field of the expansion section, and the appearance of the ablation step also led to a new internal shock wave. The gas was compressed, and the flow direction, velocity, pressure, temperature, etc., were changed. After the ablation, the temperature and pressure of the nozzle outlet location increased, the velocity decreased, and the total thrust and specific impulse decreased. [ABSTRACT FROM AUTHOR]

A wide range of propellants, and propulsion systems in space exploration by aircrafts or space vehicles was studied, developed, investigated, and commercialized. Liquid, solid, or hybrid propellants have been used for rocket's launches. In this review, a consistent definition of space propulsion systems, including solid, liquid and hybrid has been given with up-to-date state of developments. A comparison of their performances was made by theoretical and experimental specific impulses. On the other hand, ammonium perchlorate and hydrazine were used as propellants for rocket's launches and for satellite's maneuverings; respectively. However, their high toxicity and their storage problem pushed researchers and scientists to investigate and develop other eco-friendly, propellant systems, so called "green propellants", for launch or reaction control systems of satellites. [ABSTRACT FROM AUTHOR]

A numerical method for thermogasdynamic calculation of low-thrust liquid rocket engines is presented. These engines are used as end organs of the system of space attitude control for nanosatellites and small spacecraft. The method is based on the use of the TERRA software package for thermodynamic calculation and the Ansys CFX software package for gas-dynamic calculation using the Navier-Stokes equations. The results of the thermogasdynamic calculation, as well as the flow pattern of the working fluid in the chamber, are presented. The results of validating the described method are also presented. Its capabilities and limitations are analyzed. The validation procedure is based on the comparison with the results of experimental data on the Reynolds number and the momentum thickness.

On September 26, 2016 Algerian Space Agency launched AlSat-1B, the second in a series of medium resolution earth observation satellites mission. AlSat-1B electro-thermal propulsion system follows the classic design principle that has been used on previous SSTL-100 missions. With a total dry weight of 5.17 kg, the system is filled with 2.35 L self-pressurized butane. The thruster assembly is designed to use 15 W (at 28 V) redundant heater elements, it is designed to have an average mission thrust of 50 mN. With a warm up time of 10 min, the temperature of the butane will be increased by over 250 °C. This is the equivalent of about 35% gain in specific impulse over ambient temperature butane. The spacecraft dry weigh at launch site was 107 kg, the propulsion system is designed to provide at least 19 m/s Delta-v to meet the spacecraft's mission requirement. To investigate the system readiness for flight, a series of performance tests have been completed at the Surrey Satellite Technology Ltd laboratories (United Kingdom), Satellite Development Centre (Algeria) and Satish Dhawan Space Centre rocket launch site (India), the results of these tests as well as initial flight test results are presented in this paper.

A new method for determining the burning rate of a solid propellant, called the Impulse Method, is proposed in this paper. It is based on the proportional relationship between the impulse generated and the mass of the burned propellant. The pressure–time and thrust–time curves are obtained from a tubular propellant grain burning in the chamber, whose inner surface serves as the initial burning surface. Consequently, the mass of the propellant that was burned off at different pressures can be determined, and the burning rates at different pressures are derived according to the geometric parameters of the propellant grain. The Impulse Method was applied to test the burning rate of two types of propellants twice. The results show that the burning rates were consistent for the same propellant at corresponding pressures, demonstrating the feasibility and reliability of the Impulse Method. The burning rate of a GAP-based composite propellant at 20 MPa measured using the Standard Motor Method was 22.6 mm/s, and that measured using the Impulse Method was 22.2 mm/s and 22.7 mm/s, respectively. These findings indicate that the two methods have comparable accuracy. However, the Impulse Method has the advantage of obtaining the burning rate of the solid propellant at any pressure through a single test. In addition, the nozzle erosion only affected the pressure and not the burning rate. Finally, the rationality of the approach for determining the actual specific impulse was proven by comparing the results with those from another testing method.

Liquid propellant missiles are commonly applied to launch satellites that must be located in upper orbits. These systems normally use hydrogen and oxygen propellants that are non-hypergolic mixtures. In this study, the test results of a designed hydrogen/oxygen engine were evaluated, and a designed spark igniter was successfully used to start the engine. Seven hot tests were carried out to determine the performance of the engine. The effect of oxygen to fuel ratio (O/F ratio) at a constant combustion chamber pressure (800 kPa) showed that the maximum value of the specific impulse (Isp) and characteristic velocity (C*) occurs at an O/F ratio of about 2.8. Experimental tests at the constant O/F ratio of 2.8 showed that performance parameters, such as Isp and C*, were enhanced when the chamber pressure was increased. However, the trend was sluggish at pressures higher than 800 kPa.

Hydrazine and its derivatives have been used as standard propellants for spacecraft propulsion systems since the 1960s, despite being highly toxic and carcinogenic. The propellant synthesis community has constantly been looking for green alternatives for the same. Hypergolic ionic liquids (HILs) with several attractive properties, such as high energy content, high bulk density, low vapor pressure, and low toxicity, have been proposed as an alternative to hydrazine and its derivatives. In the present study, the theoretical performance of sixty-eight HILs was studied at a combustion chamber pressure of 3 MPa and a nozzle expansion ratio of 40. The specific impulse and density specific impulse of the HILs were calculated with white fuming nitric acid (WFNA), inhibited red fuming nitric acid (IRFNA), and nitrogen tetroxide (NTO) as oxidizers. The specific impulse of 2,2-dimethyltriazanium nitrate (HIL-1) was found to be 23 s higher than monomethylhydrazine (MMH), whereas its density-specific impulse was found to be 123 g-s/cm3 higher than MMH. The gains in the specific impulse and density specific impulse coupled with other desirable “green” properties for several HILs are expected to establish them as potential replacements for hydrazine and its derivatives.

The energy and ballistic characteristics of solid composite propellants (SCPs) based on 3,6-dinitro-1,4-bis(trinitromethyl)-1,4-dihydropyrazolo[4,3-c]pyrazole (II) are compared with three N-trinitromethyl biazole derivatives. Of the four oxidizing agents considered, compound II turned out to be the most effective compound in a simpled binary formulation with a hydrocarbon binder. The quantitative dependences of the energy parameters of the propellant on the properties of the compound under study (the main filler), the proportion of aluminum and additional oxidizing agents in the formulation, and the type of binder are established. [ABSTRACT FROM AUTHOR]

Novel 1,4‐disubstituted cubane derivatives have been designed and selected ones have been successfully synthesized and characterized by various analytical and spectroscopic techniques, including single‐crystal X‐ray analysis. A detailed computational study at B3LYP/6‐311++G(d,p) level of theory revealed that all newly designed 1,4‐disubstituted cubane derivatives possess higher densities, higher density‐specific impulse and superior ballistic properties when compared to conventional fuels, for example, RP‐1. These compounds also exhibit acceptable kinetic and thermodynamic stabilities which were evaluated in terms of their HOMO‐LUMO energy gap and bond dissociation energies, respectively, and are superior to TEX and many other compounds containing explosophoric groups. These results provide novel insights into the possible application of cubane‐based energetic materials. [ABSTRACT FROM AUTHOR]

A new type of marine transportation engine, the pulsed detonation hydroramjet (PDH), which was first designed, manufactured, and tested by the present authors, has been further investigated in terms of the potential improvement of its propulsive performance. PDH is composed of a pulsed detonation tube (DT) inserted in the flow-through water guide. Thrust is developed by shock-induced pulsed water jets which are periodically emitted from the water guide nozzle. The measured values of the time-averaged thrust and specific impulse in the first operation cycle were shown to always be considerably higher than those in subsequent cycles, indicating the possibility of improving the overall thrust performance. The present manuscript is aimed at clarifying the reasons for, and eliminating, cycle-to-cycle variability during PDH operation, as well as optimization of the PDH design. An experimental model of the PDH with an optically transparent water guide was designed and manufactured. The cycle-to-cycle variability was found to be caused by the overexpansion of gaseous detonation products in the DT due to the inertia of water column in the water guide. Gas overexpansion caused the reverse flow of the gas–water mixture which filled the water guide and penetrated the DT, thus exerting a strong effect on PDH operation. To eliminate the cycle-to-cycle variability, a new PDH model was developed, manufactured, and tested. The model was equipped with a passive flap valve and active rotary valve and operated on the stochiometric propane–oxygen mixture. Its test firing showed that use of the valves made it possible to eliminate the cycle-to-cycle variability and nearly double the time-averaged thrust and specific impulse reaching 40 N and 550 s, respectively. [ABSTRACT FROM AUTHOR]

Methods for the synthesis of 1,2-bis(3,4-dinitropyrazol-1-yl)diazene (1), 1,2-bis(3,5-dinitropyrazol-1-yl)diazene (2), and 1,2-bis(3-nitro-1,2,4-triazol-1-yl)diazene (3) were improved. The standard enthalpies of formation of these compounds were determined experimentally using combustion calorimetry. The density of compound 1 was refined. The possibilities of optimal use of the synthesized diazenes as energetic fillers for solid composite propellants (SCPs) were investigated. The main energy characteristics of SCPs containing diazenes were compared with those of the corresponding propellants based on hexanitrohexaazaisowurzitane, octogen, ammonium dinitramide, and ammonium perchlorate using thermodynamic analysis. [ABSTRACT FROM AUTHOR]

In order to screen novel high energetic density compounds with high power and specific impulse, eight compounds containing an aza-cage or fused ring structure were designed. The B3PW91/6–31G(d,p) theoretical method was used to optimize their structures. Their power, specific impulse, impact sensitivity and the other detonation performance were predicted, and compared with that of HMX and CL-20. The results showed that the explosive powers and the specific impulses of all eight compounds were over 150 and 2.60 N s g−1, respectively, and higher than that of HMX and CL-20. Especially, the compd-1 and compd-4 compounds have excellent comprehensive performance.

• 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 purpose of this paper is to present method of the numerical modeling of the hybrid rocket engine's work as a tool for designing engines of this type. The model is intended as an accurate and simple to use development tool for use in preliminary design stages of hybrid rocket engines, improving the effectiveness and quality of this process. General assumptions underlying the use of a model are presented, together with an analysis of past work in this field. Results of an extensive experimental campaign are presented and compared with the results of numerical modelling in order to calibrate the proposed model and evaluate its accuracy. Parameter variation and optimization were conducted, proving functionality of the methodology. Presented numerical calculations show that the adopted approach to reduce the analysis time and complexity was correct. This method of numerical calculation of hybrid engine working parameters combines aspects of accuracy and simplicity at the early design stage to avoid time-consuming and costly changes in subsequent detailed stages of design.

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]

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