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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 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 (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.

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.

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.

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.

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.

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 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.

Objectives. The addition of high-density carbon materials to jet fuels can lead to a significant increase in the volumetric energy of the fuel combustion. The purpose of the current study was to thermodynamically analyze the possibility of obtaining model hydrocarbon fuels from toluene and T-1 using stacked-cup multiwall carbon nanotubes (MWCNTs). Methods. Bomb combustion calorimetry was used to define the combustion energy of the MWCNTs in the crystalline state. The temperature dependence of the MWCNTs’ heat capacity in the range 5–370 K and the fusion parameters were estimated using low-temperature adiabatic calorimetry. The physical density of MWCNTs was measured using the pycnometric method. The sedimentation stability of the mixtures of MWCNTs with liquids was determined using centrifugation at 7000 g. The calculations were carried out in MS Excel. Results. The energy and enthalpy of combustion of a technical sample of MWCNTs in the crystalline state were determined. Based on the smoothed heat capacity values, the standard thermodynamic functions (enthalpy, entropy, and Gibbs reduced energy) of MWCNTs in the crystalline state were obtained in a temperature range of 0–2000 K. The extrapolation of the MWCNTs’ heat capacity was carried out at a temperature of up to 2000 K using the heat capacity of crystalline graphite. It has been established that mixtures of MWCNTs with liquids containing more than 33 mass % of MWCNTs are stable during centrifugal sedimentation at 7000 g. For the toluene–MWCNTs and fuel T-1–MWCNTs model systems, the specific and volumetric combustion energies, the adiabatic combustion temperatures, and the conditional final maximum speed of the model rockets with fuel of various compositions were also calculated. Conclusions. The thermodynamic analysis showed that the addition of MWCNTs can significantly increase the volumetric energy intensity of traditional jet fuels, which can in turn improve the operational characteristics of drones and rockets.

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.

Pulse Detonation Engine (PDE), is an emerging and promising propulsive technology all over the world in the past few decades. A pulse detonation engine (PDE) is a type of propulsion system that uses detonation waves to combust the fuel and oxidizer mixture. Theoretically, a PDE can be operate from subsonic to hypersonic flight speeds. Pulsed detonation engines offer many advantages over conventional air-breathing engines and are regarded as potential replacements for air-breathing and rocket propulsion systems, for platforms ranging from subsonic unmanned vehicles, long-range transportation, high-speed vehicles, space launchers to space vehicles. This article highlights the operating cycle of PDE, starting with the fuel-oxidizer mixture, combustion and Deflagration to detonation transition (DDT) followed by purging. PDE combustion process, a unique process, leads to consistent and repeatable detonation waves. This pulsed detonation combustion process causes rapid burning of the fuel-oxidizer mixture, which cannot be seen in any other combustion process as it is a thousand times faster than any other mode of combustion. PDE not only holds the capability of running effectively up to Mach 5 but it also changes the technicalities in space propulsion. The present paper is the extension of the previous study which is also a well characterized status report of PDE in different areas. The present study deals with the categorization of the design approach, computations & simulations, flow visualization, DDT & Thrust enhancement, PDRE’s, experimental detonation engines with some of the experience and research undertaken in Punjab Engineering College under the complete supervision and guidance of Prof. Tejinder Kumar Jindal followed by applications of PDE technology.

The performances of a dual-combustion ramjet (DCR) and a scramjet were compared via computational fluid dynamics numerical simulation to provide theoretical guidance for engine selection for a hypersonic vehicle. Kerosene, C12H23, with an equivalence ratio of 0.8, was employed as the fuel, and the reactive flow was modeled using six-species and four-step chemistry. The results show that the DCR has a central combustion mode, which has a smaller temperature gradient and more uniform heat release, resulting in higher combustion efficiency, compared to the near-wall combustion mode of the scramjet. The total pressure recovery coefficient of scramjet is 0.9% lower than that of DCR under the Ma6 condition, but 5.6% higher than that of DCR under the Ma7 condition. The combustion efficiency of DCR is 35.6% and 25.4% higher than that of the scramjet under Ma6 and Ma7 conditions, respectively. The decrease in the combustion efficiency of the DCR is caused by the increase in the dissociation rate of CO2 into CO with the increase in temperature. The performance of DCR is better than that of scramjet under both conditions. However, the performance advantage of DCR decreases as the Mach number increases. Specifically, under the conditions of Ma6 and Ma7, the specific impulse or specific thrust of DCR was 2.67 times and 1.51 times that of scramjet, respectively.

The current transmission and reflection laser ablation micropropulsion modes have the problem of a complex working medium supply system in engineering. Therefore, we propose large-spot laser ablation with a one-dimensional supply mode. In order to verify this ablation mode, a multipulse ablation experiment of submillimeter-scale light spots was carried out on the surface of pretreated copper and nickel under the atmosphere using an ultrafast laser with a pulse width of 290 fs and 10 ps. The results show that femtosecond laser multipulse ablation (FLMA) leads to the grain refinement of copper, the crater quality of the two metals under FLMA is better, and picosecond laser multipulse ablation (PLMA) causes the crater of nickel to form a dense remelting bulge that affects laser absorption; both metals have obvious heat-affected zones after FLMA and PLMA, the heat-affected zones of nickel are 5–10% larger than those of copper, and the ablation depth of copper is deeper. Under the same conditions, the ablation mass of copper is smaller than that of nickel, and the specific impulse performance of laser ablation micropropulsion is better.

The paper presents a fast mathematical model that can be used to quickly asses the propulsive characteristics of liquid propelled rocket engines. The main propulsive parameters are computed using combustion surfaces obtained after a nonlinear data fitting analysis. This approach is much more time efficient than using standard codes which rely on frequent calls of the Fuel Combustion Charts and interpolating their data. The tool developed based on the proposed mathematical model can be used separately or it can be integrated in a multidisciplinary optimisation algorithm for a preliminary microlauncher design.

Pulse Detonation Engine (PDE), is an exciting propulsion technology for the future and has been able to seek considerable attention over the last era. It has the potential to work efficiently in the modern cosmos. It works on a Humphrey cycle offering a great opportunity, which outweighs the conventional Brayton cycle. The operating cycle of PDE starts with the fuel-oxidizer mixture, combustion and DDT followed by purging. The PDE combustion process, which is a unique process, leads to consistent and repeatable detonation waves. This pulsed detonation combustion process causes rapid burning of the fuel-oxidizer mixture, which cannot be seen in any other combustion process as it is a thousand times faster than any other mode of combustion. PDE not only holds the capability of running effectively up to Mach 5 but it also changes the technicalities in space propulsion. The present study deals with the categorization of design approach, thermal analysis, performance analysis, fuel-based analysis, PDE combustors, detonation propagation, experimentation analysis, valving techniques, ignition studies, parameters & specification optimization, instrumentation, thrusters, parameters affecting specific impulse & thrust, hybrid PDE’s & turbine integration.

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.

The energy characteristics and theoretical performance of the hybrid rocket fuels are discussed in this paper. Aluminum (Al) and boron (B) metal additives were used to increase the energy density of the paraffin-based solid fuels. To predict the energy characteristics, the heat of combustion was evaluated by adiabatic bomb calorimetry. Theoretical performance parameters such as specific impulse (Isp), flame temperature, and characteristic velocity were obtained with NASA Chemical Equilibrium with Applications (CEA) code. Calorimetric test results revealed that paraffin/polyethylene/boron (P/PE/B)-based fuel formulations exhibited the highest heat of combustion among all the tested fuels. The heat of combustion value of the P/PE/B sample at 25 wt% B loading was found to be 9612 ±16 cal/g and 9293±17 cal/g for the P/PE/Al fuel formulation. The CEA results showed that the addition of Al to paraffin is noneffective in improving specific impulse performance. When B loading increased from 5 to 25 wt% in the P/PE/B, the Isp increased by 47 s compared to pure paraffin. A specific impulse increase implies the possible propellant mass saving. The reduction of the oxidizer and fuel masses may yield increased payload performance for given boundary conditions. The P/PE/B25 formulation has reported the highest value of characteristics velocity (C*) compared to other paraffin-based formulation.

The energy abilities of solid composite propellants based on some nitrosubstituted tetrazole derivatives, which are still hypothetical but they are promising for create on their basis rather powerful propellants, have been considered. It is confirmed by the thermodynamic calculations

The kinetic theory of gases methods has been used to evaluate plasma arc parameters in the arcjet thruster device. It greatly means the compression process in adiabatic character and gas temperature increasing for excited atoms Boltzmann distribution over the energy levels. Using argon gas for arcjet working process carring out, the way of gas temperature eveluation is offered. The results of the study can be use in the spacecraft arcjet development

The energy abilities of solid composite propellants based on a couple of new high-enthalpy N-oxides (furazano-terazine-dioxide and tetrazino-tetrazine-1,3,6,8-tetraoxide) have been considered, it is found that these two compounds show record energy characteristics — specific impulse up to 273 s (at pressure in the combustion chamber and at the nozzle section 4,0 and 0,1 MPa respectively) at the combustion temperature not higher than 3700 K. Formulations based on these two compounds are very powerful if any binder (an active binder or a hydrocarbon one) is used

Various cyclic N8 isomers are considered at the levels of B3LYP/6–311++G(d,p) and B3LYP/cPVTZ. Some energies and molecular orbital properties are obtained. The structures are found to be stable in the singlet state but mostly unstable in the triplet. The heats of formation values calculated by means of T1 recipe reveals that they are highly endothermic. NICS(0) values have been calculated and aromaticity/antiaromaticity of the rings are discussed. Additionally, theoretical IR,UV-VIS spectra and the calculated impulse values have been obtained.

To develop elements of a system for contact-free transportation of objects in space has now become an urgent task for the contemporary space-related activities. The purpose of work that is presented hereinafter was to conduct ground tests of the ion source, which is a key element of the above-mentioned system, and to compare the obtained experimental data with the mathematical simulation results in order to build a refined physical and mathematical model of the ion source. Such model was built on the basis of the classical problem regarding the motion of charged particles in an electrostatic field. Parameters of the ion source have been determined experimentally for several operating modes using various structural designs of the ion source electrodes. Two types of ion optics were tested—with slit and round apertures. Good correlation between simulation results and experimental data has been demonstrated. The optimum ion source operation modes have been identified to ensure minimum divergence angles for the plasma beam exiting from the ion source, which in its turn maximizes the pulse transmitted to the transported object.

Buried high explosive (HE) charges represent a high threat to military vehicles. The detonation of these charges can lead to significant momentum transfer onto vehicles and their occupants. A detailed understanding of the physical processes involved in the loading of vehicle structures is necessary for an optimization of effective countermeasures and protection systems. A quantitative description of the local momentum distribution on the vehicle underbody due to the detonation process is of special importance. In the following, a new test setup is presented that allows the experimental determination of the specific impulse distribution. It is based on a ring arrangement where the elements are nested into each other and the velocity of each ring is correlated with the local specific impulse at its position. The momentum transfer to a vehicle depends on a number of influencing factors such as: charge mass, embedding material (e.g. sand, gravel, clay), density, water content, saturation, depth of burial, ground clearance and vehicle shape. The presented technology is applied to quantify the influence of the embedding material (alluvial sand, quartz sand), the burial depth and the water content on the local specific impulse distribution. The obtained data can be used as initial condition for the numerical simulation of occupant safety assessment and as input for empirical modeling of momentum transfer on structures.

Due to the continuous run for a green environment the current article proposes a new type of solid propellant based on the fairly new synthesized oxidizer, ammonium dinitramide (ADN). Apart of having a higher specific impulse than the worldwide renowned oxidizer, ammonium perchlorate, ADN has the advantage, of leaving behind only nitrogen, oxygen and water after decomposing at high temperatures and therefore totally avoiding the formation of hydrogen chloride fumes. Based on the oxidizer to fuel ratios of the current formulations of the major rocket solid booster (e.g. Space Shuttle’s SRB, Ariane 5’s SRB) which comprises mass variations of ammonium perchlorate oxidizer (70-75%), atomized aluminum powder (10-18%) and polybutadiene binder (12-20%) a new solid propellant was formulated. As previously stated, the new propellant formula and its variations use ADN as oxidizer and erythritol tetranitrate as fuel, keeping the same polybutadiene as binder.

[Objectives]The influence of the explosion venting plate thickness on the effect of the venting in the new type broadside structure adopted by the DDG 1000 destroyer is studied in this paper.[Methods] First,the reliability of the simulation method was verified by experimental data. Then the simulation model of the explosion venting cabin was established using the finite element analysis software. And the principle of thin-plate explosion venting structure was analyzed;the damage failure and load variation under different thin plate thicknesses were studied. Finally,the quadratic function model of specific impulse and deflection changing with the thickness is obtained by function fitting.[Results]The results show that the failure of the cabin first occurred at the connection between the thin plate and bulkhead,and gradually extended to the corners and edges of bulkhead;the thinner the plate,the likely an explosion venting opening is to form. When an explosion venting opening forms,the whole thin plate flies out of the cabin;the existence of the explosion venting structure has little effect on the initial shock wave overpressure,but it can effectively reduce the quasi-static pressure and specific impulse in the cabin. The main factors causing the deformation of the protective bulkhead are the initial shock wave and the reflected shock wave,and the main factor causing the ultimate damage of the protective bulkhead is the quasi-static pressure acting for a long time.[Conclusions]The research results can provide some reference for the explosion venting design of the ship side cabin structure.

Four types of nitrogen-rich heterocycles substituted with -NO2, -NHNO2 and -C(NO2)3 explosophoric groups were explored as potential explosives and propellants materials. The calculated crystal density (ρ0)and the condensed phase heat of formation (ΔH·0f)for each of the twelve structures investigated shows that all these derivatives possess high (1.834-1.980 g cm-3)(ΔH·0f) and (605-2130 kJ kg-1) values. Interesting properties such as detonation velocity (D), pressure (P) and specific impulse (Isp) were calculated using the Kamlet-Jacobs method and ISPBKW thermochemical code. Detonation velocity and pressure in excess of 8.44 km s-1 and 32.87 GPa was obtained in all cases. Furthermore, trinitromethyl substituted derivatives shows performance exceeding that of HMX with an estimated D = 9.32-9.72 km s-1 and P = 40.61-43.82 GPa. Some -NO2 and -NHNO2 substituted derivatives were shown to be impact insensitive while retaining good detonation performance and thus are regarded as potential replacement for current RDX -based explosives. Finally, the calculated specific impulse (Isp between 248 and 270 s) of all investigated derivatives indicate that these energetic materials can be considered as possible ingredient in future rocket propellant compositions.

The energy characteristics and theoretical performance of the hybrid rocket fuels are discussed in this paper. Aluminum (Al) and boron (B) metal additives were used to increase the energy density of the paraffin-based solid fuels. To predict the energy characteristics, the heat of combustion was evaluated by adiabatic bomb calorimetry. Theoretical performance parameters such as specific impulse (Isp), flame temperature, and characteristic velocity were obtained with NASA Chemical Equilibrium with Applications (CEA) code. Calorimetric test results revealed that paraffin/polyethylene/boron (P/PE/B)-based fuel formulations exhibited the highest heat of combustion among all the tested fuels. The heat of combustion value of the P/PE/B sample at 25 wt% B loading was found to be 9612 ±16 cal/g and 9293±17 cal/g for the P/PE/Al fuel formulation. The CEA results showed that the addition of Al to paraffin is noneffective in improving specific impulse performance. When B loading increased from 5 to 25 wt% in the P/PE/B, the Isp increased by 47 s compared to pure paraffin. A specific impulse increase implies the possible propellant mass saving. The reduction of the oxidizer and fuel masses may yield increased payload performance for given boundary conditions. The P/PE/B25 formulation has reported the highest value of characteristics velocity (C*) compared to other paraffin-based formulation.

In work explored effect of mixed metallic fuel on the thermodynamic and ballistic characteristics of High-Energy Materials (HEMs). Efficiency of mixed metallic fuel estimated in relation to the characteristics of HEMs, containing conventional metallic fuel - aluminum. The main criteria of comparison are the specific impulse and the composition of the combustion products HEMs.