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As a multiple-start ignition method for liquid-fuel rocket engines, the gas-dynamic igniter has many advantages, such as a simple configuration, low weight and high structural strength. However, because of the complexity of the flow of the working fluid, the details of the thermal processes involved are not clearly understood. In this study, the thermal and flow characteristics of a gas-dynamic igniter are investigated through numerical simulations using the software OpenFOAM. The simulation results show that the pressure within the igniter undergoes oscillations. The axial flow velocity decreases across the pressure wave front so that the kinetic energy of the flow is converted to thermal energy. The temperature increase within the oscillation tube of the igniter is strongly correlated with the entry mass flow. Therefore, the tube inlet area should be designed according to the igniter nozzle flow to achieve the best performance from a gasdynamic igniter. [ABSTRACT FROM AUTHOR]

This paper presents transient simulations and experimental studies of the ignition process of the hybrid rocket motors (HRMs) using 90% hydrogen peroxide (HP) as the oxidizer and polymethyl methacrylate (PMMA) and Polyethylene (PE) as fuels. A fluid-solid coupling numerically method is established based on the conserved form of the three-dimensional unsteady Navier-Stokes (N-S) equations, considering gas fluid with chemical reactions and heat transfer between the fluid and solid region. Experiments are subsequently conducted using high-speed camera to record the ignition process. The flame propagation, chamber pressurizing process and average fuel regression rate of the numerical simulation results show good agreement with the experimental ones, which demonstrates the validity of the simulations in this study. The results also indicate that the flame propagation time is mainly affected by fluid dynamics and it increases with an increasing grain port area. The chamber pressurizing process begins when the flame propagation completes in the grain port. Furthermore, the chamber pressurizing time is about 4 times longer than the time of flame propagation. [ABSTRACT FROM AUTHOR]

Purpose The capability to predict and evaluate the motor pressure during each phase by means of a numerical analysis can significantly increase the efficiency of the preliminary design process with a reduction of both the motor development and operational costs. This paper aims to perform numerical simulation to analyze the ignition transient in solid rocket motor by solving Euler equation coupled with some semi-empirical correlations. These relations take into account the main phenomena affecting the ignition transient. Coupling relationships include the heat transfer of the gas to the propellant and erosive burning rate relationship.Design/methodology/approach The current research effort divides motor into series of control volumes along the port axis, and the variation of port area, burning surface and burning rate along the port are taken into account. A set of governing equations are then solved using explicit, time-dependent, predictor-corrector finite difference method. The numerical model helps to capture and embed shock wave associated with igniter flow within the solution. Second-order artificial viscosity dampens out the numerical oscillations due to sharp gradient within the flow field. The developed computer code predicts the start-up characteristics of motor. The study also provides comparison of simulation results with in-house experimental motor.Findings Simulations are performed with and without erosive burning to demonstrate that the flow model is a good physical approximation of motor. Numerical results calculated by this model without erosive burning are not in good agreement with experimental results. This minor discrepancy has motivated the inclusion of erosive burning in numerical model. The simulated results are then compared with the experimental data for head-end and rear-end pressure. The agreement between simulation and experiment is remarkable. In summary, major finding of this study is that unsteady quasi-one-dimensional gas dynamic model can capture the flow field in the motor during ignition transient effectively.Research limitations/implications Unsteady quasi-one-dimensional gas dynamic model can capture the flow field in the motor during ignition transient effectively. However, in systems where two- and three-dimensional effects are pre-dominant, one would require to develop a more elaborate, multi-dimensional model which will allow for further understanding of the flow behavior and eventually lead to modeling of rocket motors with more complex geometries.Practical implications The close agreement between experimental and simulation results can be considered as forced to some degree, because the general mathematical model of erosive burning contains a free variable erosive burning exponent. However, in future, this variable can be established a priori by erosive burning tests.Originality/value The solid propellant ignition process consists of series of rapid events and must be completed in a fraction of a second. An understanding of the dynamics of ignition has become increasingly vital with the development of larger and more sophisticated solid propellant rocket motors. This research effort provides the simulation framework to predict and evaluate the motor pressure during each phase by means of a numerical analysis, thus significantly increasing the efficiency of the preliminary design process with a reduction of both the motor development and operational costs. [ABSTRACT FROM AUTHOR]

In the present study, the ignition effect on detonation initiation was investigated in the air-breathing pulse detonation engine. Two kinds of fuel injection and ignition methods were applied. For one method, fuel and air was pre-mixed outside the PDE and then injected into the detonation tube. The droplet sizes of mixtures were measured. An annular cavity was used as the ignition section. For the other method, fuel-air mixtures were mixed inside the PDE, and a pre-combustor was utilized as the ignition source. At firing frequency of 20 Hz, transition to detonation was obtained. Experimental results indicated that the ignition position and initial flame acceleration had important effects on the deflagration-to-detonation transition. [ABSTRACT FROM AUTHOR]

On the basis of the international processing approach was adopted to solve the problem of engine hydrogen emissions, by direct combustion method to handle the cryogenic hydrogen emitted from the engine. Using high temperature two-phase jet flow field from the ignition motor, which contains a large number of metal particles as the ignition source of hydrogen emitted by Rocket Engine. The numerical simulation of ignition motor's high speed multiphase exhaust plume flow field was performed by using FLUENT software, temperature, velocity field and particle distribution with high temperature was obtained in the exhaust plume. Meanwhile, the installation form of ignition motor was designed and optimized, including its location, number, combination structure of ignition motor, in order to obtain the best plan to ensure the reliability of successful ignition. [ABSTRACT FROM AUTHOR]

The transient characteristics of kerosene/oxygen spray were investigated using a windowed rocket combustor. In order to investigate the effects of ignition conditions, viz. aerodynamic Weber number and the oxygen fraction of pre-filled gas, on GOx/kerosene ignition, the ignition process was visualized directly using the shadowgraph technique. In addition, the dynamic pressure and global OH * chemiluminescence were measured simultaneously. The transient ignition process was significantly changed by the type of pre-filled gas (nitrogen or oxygen). The nitrogen-filled ignition condition resulted in smooth ignition, while the oxygen-filled ignition condition induced two-step ignition following the sequence of ignition with high-pressure peak, temporary extinction, and re-ignition. The ignition process was less affected by the spraying condition ( We a ), except that the change of the spraying condition influenced ignition to steady state transition, such as a combustion chamber pressure overshoot. It was revealed that the oxygen-filled ignition condition caused drastic energy generation at the moment of ignition, producing a high-pressure peak and causing reverse flow of combustion gas through the gaseous oxidizer injector. This resulted in temporary extinction of the initial flame followed by re-ignition. Once the two-step ignition phenomenon occurred, the combustion chamber pressure did not increase to its steady state value during the entire run time (1.5 s). As the oxygen fraction in the pre-filled gas increased because of the increase in oxygen pre-injection time, the ignition delay time decreased and the peak value of global OH * intensity increased exponentially. When the oxygen pre-injection time became more than 14 ms, the ignition process transformed from smooth ignition to two-step ignition, and the peak value of global OH * intensity began increasing, the ignition process was converted from the smooth ignition to the two step ignition. [ABSTRACT FROM AUTHOR]

This paper is to provide up-to-date information on the state of acoustic ignition system development, with such systems seen as one of the most promising alternative to the traditional electrical systems. The article summarizes the principles of the gas-dynamic (acoustic) ignition systems, describes the most commonly used igniter designs, and also describes some unusual and potentially promising igniter designs. It also contains a survey of contemporary domestic and foreign theoretical and experimental research in the field of gas-dynamic ignition. Descriptions of promising igniter designs are provided. The paper analyses identified strengths and weaknesses of the construction designs described. On the basis of accumulated and presented materials, recommendations are drawn for the future course of development, which comprise the need to address the thermal issues of such igniters, their lifetime and the need for more comprehensive studies on atypical designs, which can be more efficient than their currently used counterparts. [ABSTRACT FROM AUTHOR]

Die Saat-Emulsionspolymerisation ist eine effiziente Herstellungsmethode für nanoskalige multifunktionale Hybridmaterialien. In einem zweistufigen Prozess wurden zunächst Eisenoxidnanopartikel mit einer Hülle aus Polystyrol verkapselt. Diese wurden anschließend mit Quantenpunkten in Quantenstäben (QDQRs) belegt, sodass Nanohybride mit Eisenoxidkern und QDQR-Satelliten resultierten. Die Synthese einer nahtlosen Polystyrolhülle ermöglicht sowohl die Fixierung des Konstruktes als auch die Steuerung seiner Zusammensetzung. Dadurch bleiben die Fluoreszenzeigenschaften der QDQRs und die magnetischen Eigenschaften des Eisenoxids unverändert, was mittels Einzelpartikeluntersuchung, Messung der Fluoreszenzabklingzeit und Relaxometrie nachgewiesen werden kann. Die Biofunktionalisierung erfolgte direkt durch Copolymerisation von entsprechenden Affinitätsmolekülen. Die Anwendbarkeit der Methode auf verschiedenartige Nanopartikel wurde demonstriert. [ABSTRACT FROM AUTHOR]

A laser-ignition micro solid rocket for application as propulsion on a 1-10 kg-class microspacecraft was proposed and designed. The laser-ignition micro solid rocket uses boron/potassium nitrate as propellant, and a number of solid propellant blocks are ignited by irradiating their surface with a 1 W diode laser through a laser-incident window. According to previous experiments, contamination of the laser-incident window can cause ignition failure to as high as 39% (7 thrusters out of 18) as a result of the gas generated at the initial phase of the laser irradiation. To reduce this contamination, the distance between propellant and laser-incident window is the parameter of interest in this study. The ignition probability was measured using a newly designed propellant stand with variable distance between propellant and laser-incident window, and an optimized window-to-propellant distance was applied to the thrusters. As a result, the ignition probability reached 100% (26 thrusters out of 26) while maintaining the specific impulse. [ABSTRACT FROM AUTHOR]

Highlights: [•] Long-duration experiments were conducted to study effects of wall-temperature. [•] It was observed that wall-temperature had a positive effect on the DDT process. [•] Wall-temperature was also found to have a relationship with detonation pressure. [•] Extraordinary high wall-temperature would cause pre-ignition of fresh reactants. [Copyright &y& Elsevier]

In order to investigate the ignition mechanism of composite solid propellant under waterjet high pressure from the dust ignition mode perspective, the modified Hartman bomb device was used to study experimentally the minimum explosive mass concentration of AP/A1/ ferrocene mixed system. Processes of dust explosion of the typical composite propellant were discussed. Effects of components, temperature and humidity on the minimum explosive mass concentration were investigated. Results show that the minimum explosive mass concentration increases with the increase of humidity, while it decreases with the increase of temperature. Also, the minimum explosive mass concentration of the mixed system is sensitive to the content of AP/ferrocene. [ABSTRACT FROM AUTHOR]

We have developed a sensitivity analysis permitting effective estimation of the change in the impulse responses of a microthrusters and in the ignition characteristics of the solid-fuel charge caused by the variation of the basic macrokinetic parameters of the mixed fuel and the design parameters of the microthruster's combustion chamber. On the basis of the proposed sensitivity analysis, we have estimated the spread of both the propulsive force and impulse and the induction period and self-ignition temperature depending on the macrokinetic parameters of combustion (pre-exponential factor, activation energy, density, and heat content) of the solid-fuel charge of the microthruster. The obtained results can be used for rapid and effective estimation of the spread of goal functions to provide stable physicochemical characteristics and impulse responses of solid-fuel mixtures in making and using microthrusters. [ABSTRACT FROM AUTHOR]

In the secondary chamber of ducted rocket, there exists a relative speed between boron particles and air stream. Hence, the ignition laws under static conditions cannot be simply applied to represent the actual ignition process of boron particles, and it is required to study the effect of forced convective on the ignition of boron particles. Preheating of boron particles in gas generator makes it possible to utilize the velocity difference between gas and particles in secondary chamber for removal of the liquid oxide layer with the aid of Stoke's forces. An ignition model of boron particles is formulated for the oxide layer removal by considering that it results from a boundary layer stripping mechanism. The shearing action exerted by the high-speed flow causes a boundary layer to be formed in the surface of the liquid oxide layer, and the stripping away of this layer accounts for the accelerated ignition of boron particles. Compared with the King model, as the ignition model of boron particles is formulated for the oxide layer removal by considering that it results from a boundary layer stripping mechanism, the oxide layer thickness thins at all times during the particle ignition and lower the ignition time. [ABSTRACT FROM AUTHOR]