Numerička analiza putanje aktivno-reaktivnog projektila

A computational analysis was carried out in order to study the effect of solid rocket motor operating parameters on the trajectory elements of 155mm rocket-assisted projectile. Two numerical solutions have been programmed for this purpose. The first for the rocket motor internal ballistic calculation. The second is intended for the six degree-of-freedom (6-DOF) trajectory determination for both unpowered and powered flight. The effect of three operating parameters of the solid rocket motor on the trajectory elements was taken into account, namely: (1) thrust-time profile (neutral, regressive or progressive), (2) working time and (3) ignition delay. The results obtained following this analysis are helpful as a preliminary stage in the design of extended range projectiles in order to optimize flight performance.

Study on cook-off behavior of HTPE propellant in solid rocket motor

3-D grain burnback analysis of solid propellant rocket motors: Part 1 – ballistic motor tests

Accurate prediction of ignition delay and flame spread rate in solid propellant rocket motors is of great topical interest. In this paper using a standard k-ω turbulence model numerical studies have been carried out to examine the influence of solid rockets port geometry on ignition delay and the flame spread pattern. We observed that with the same inflow conditions and propellant properties heat flux histories and ignition time sequence are different for different port geometries. We conjectured from the numerical results that in solid rocket motors with highly loaded propellants, mass flux of the hot gases moving past the burning surface is large. Under these conditions, the convective flux to the surface of the propellant will be enhanced, which in turn enhance the local Reynolds number. This amounts a reduction in heat transfer film thickness and enhanced heat transfer to the propellant with consequent enhancement in the dynamic burn rate resulting the undesirable starting pressure transient. We concluded that, the more accurate description of gas phase to surface heat transfer process will give a better prediction and control of ignition delay and flame spread rate in solid propellant rockets.

The efficient development of hypergolic fuels requires an interdisciplinary approach involving ab initio modeling, synthesis, and experimental physical chemistry. Candidate molecules must exhibit hypergolic ignition delay times that are fast enough to warrant further testing for safety and performance criteria. Hypergolic ignition delay apparatus has been mentioned in the open literature for six decades, but accurate, detailed, modern ignition delay hardware that uses inexpensive laboratory building blocks and a minimum of custom circuitry is still needed. This article details line-of-sight electro-optical circuitry with direct digital readout and additional oscilloscope recording that can be used to measure total ignition and chemical delay times for screening candidate fuels. We also illustrate the value of high speed video and quantum chemical calculations to supplement the ignition delay measurements for a comprehensive approach to hypergolic fuel research.

The combustion instability remained as a problem during solid rocket motors progress. In this study, combustion instability of a double base solid rocket motor was estimated wherein acoustical and erosive burning was considered. The ballistic parameters of the solid rocket motor were presented and the results were compared with simulations, which show a good agreement. The result shows that specific impulse of the solid propellant rocket motor was obtained 211 s and the maximum pressure of the combustion chamber reaches 13.8 MPa. Combustion instability analysis indicates that the motor became unstable after 0.5 s.

This chapter describes miniature stress sensor technology for monitoring the thermal stresses and ignition pressurization loads in solid rocket motors. The study was part of a larger international (TTCP) collaborative effort carried out from 1988 to 2002 to validate the instrumentation and analytical stress analysis and service life prediction methodologies for solid composite rocket motors, and thus establish improved, more reliable, cheaper and non-destructive capabilities for service life prediction and extension. Different motor configurations were used by the different countries in this collaborative program. The Australian effort, described in this chapter, used an end burning generic research motor (Pictor). The embedded transducers, in this end burning motor and in the different motor designs used by the other TTCP countries, were found to be stable in the temperature range used in the environmental testing program and gave consistent data during propellant cure, environmental testing and static firing of the motors. The rocket motor instrumentation and data reduction techniques were described. The data from the instrumented motors under various thermal storage loading conditions (multiple thermal cycling, shocking, accelerated ageing at elevated temperature) were used to validate the stresses and critical failure modes predicted by structural finite element modelling and a modified fracture mechanics approach for nonlinear viscoelastic materials. These studies verified the ability of the miniature bond stress sensors to detect cracking / damage in the propellant charge. The advancement in bond stress sensor technology was further used to investigate failure analysis of rocket motors under ignition pressurization conditions. Results from these studies demonstrated that the sensors are safe for static firing and could accurately measure pressure in different regions of the burning motor. The stress sensor data from this international collaborative program showed that the stress sensor technology could be used for real-time structural health monitoring of solid rocket motors to detect cracks and debonds in the propellant and to continuously monitor the extent of damage. The results from the instrumented Pictor motor verified and validated the thermal distributions, stress / strain states and regions of high propensity for crack propagation predicted by finite element modelling and fracture mechanics. The use of the instrumented motor data in probabilistic service life prediction methodologies and other NDE methods are also discussed.

The solid rocket motor upper stage for a space launch vehicle is a more efficient propulsion technology than the liquid rocket motor upper stage. Its grain design has the potential to be crucial in terms of lowering inert mass by adopting improved volume efficiency with the lowest practicable sliver size while keeping maximum strength. Specifically, the strategy for (3D) grain arrangement of the slot for the upper stage solid rocket engine has been described in this paper. The complex configuration is established by the design process, which takes place under a parametric model of geometry in (CAD) software and is typified by varied dynamics. When constructing solid propellant rocket motors, grain arrangement is a vital and critical step. Accurate estimates of grain geometric properties play a key role in performance prediction and can be a vital and critical stage in the design of solid propellant rocket motors. This research study proposes an effective performance-matching design framework for solid rocket motors that are tuned to suit a range of thrust performance criteria. The framework is constructed utilising an innovative and specialised general design technique that was designed to evaluate the general design parameters, which is given in this study. Because of the findings obtained, it can be stated that the recommended framework is a practical and efficient approach for solid rocket engine design and development.

: The ignition delay and combustion of amorphous and crystalline boron particles is investigated at elevated temperatures and pressures for wet, dry, and fluorine-containing atmospheres. Particles ranging from submicron to 32 microns in diameter are ignited in the ambient conditions produced by a reflected shock wave in a shock tube. The ignition delay and combustion times are examined as a function of temperature for pressures of 8.5, 17, and 34 atm and for oxidizer mixtures of 100% oxygen, 30% water vapor, 1-3% sulfur hexafluoride, and 6-12% hydrogen fluoride. Results indicate that pressure in the range studied does not affect the ignition delay or burn time. The additives, water vapor and sulfur hexafluoride, reduce the ignition delay time for amorphous and sub-micron crystalline boron when compared to oxygen. For 20 microns particles, H2O and SF6 reduce the ignition temperature limit from 2500 deg K in pure oxygen to 2200 deg K and 1900 deg K, respectively. Burn time is unaffected by the additives. Hydrogen fluoride did not show any change in ignition delay or burn time compared to pure oxygen. At the range of temperatures tested, very little (less than 2%) of HF is dissociated into H and F atoms. The report also presents reviews of previous chemical and physical models that have attempted to explain why boron powder is relatively difficult to ignite. Ignition of metal powder, Shock tube initiation.

: The ignition delay and combustion of amorphous and crystalline boron particles is investigated at elevated temperatures and pressures for wet, dry, and fluorine-containing atmospheres. Particles ranging from submicron to 32 microns in diameter are ignited in the ambient conditions produced by a reflected shock wave in a shock tube. The ignition delay and combustion times are examined as a function of temperature for pressures of 8.5, 17, and 34 atm and for oxidizer mixtures of 100% oxygen, 30% water vapor, 1-3% sulfur hexafluoride, and 6-12% hydrogen fluoride. Results indicate that pressure in the range studied does not affect the ignition delay or burn time. The additives, water vapor and sulfur hexafluoride, reduce the ignition delay time for amorphous and sub-micron crystalline boron when compared to oxygen. For 20 microns particles, H2O and SF6 reduce the ignition temperature limit from 2500 deg K in pure oxygen to 2200 deg K and 1900 deg K, respectively. Burn time is unaffected by the additives. Hydrogen fluoride did not show any change in ignition delay or burn time compared to pure oxygen. At the range of temperatures tested, very little (less than 2%) of HF is dissociated into H and F atoms. The report also presents reviews of previous chemical and physical models that have attempted to explain why boron powder is relatively difficult to ignite. Ignition of metal powder, Shock tube initiation.

In the present research, an experimental and numerical investigation has been performed to study initial pressure rise in multi grain solid rocket motor. A cluster of seven uninhibited tubular grains of double base propellant is analysed to get near neutral pressure in rocket motor. It is experimentally observed that in some static firings initial pressure rise in first 100 ms is higher than the predicted value of ∼7.5 MPa. The average pressure in the rocket motor is ∼7 MPa and total burning time of these grains is ∼2 s. ANSYS software is used for numerical analysis to investigate the effects of various % openings of nozzle end (NE) metal grid on pressure rise in the rocket motor with and without peripheral opening. The analysed results are compared with experiments. It is found that, as the %opening of NE metal grid increases, the initial pressure in the rocket motor decreases. The peripheral opening is also important to reduce pressure rise in rocket motor. This study helps in reducing the initial pressure rise in the combustion chamber which is essential for safe working of the rocket motor.

<div class="htmlview paragraph">Cycle-to-cycle cylinder pressure variation has been observed in a CFR prechamber diesel engine when low ignition quality (low cetane number) fuels are burned. A statistical analysis of this phenomenon for various fuels and blends with cetane numbers as low as zero has been made. Operating conditions used were those specified by the ASTM Cetane Method for rating diesel fuels, in which the inlet air temperature is 150°F. Additional analysis was made at increased inlet air temperatures of 250°F and 350°F.</div> <div class="htmlview paragraph">The cycle-to-cycle variation has been characterized by the variation in the ignition (or pressure rise) delay time. It has been found to increase sharply as fuel cetane number is decreased below 20. The variation in dynamic injection timing was also measured and correlated with that for ignition delay. It has been concluded that the observed injection timing irregularities have no significant effect on ignition delay time, and that the autoignition process becomes unstable due to other factors when fuels of low ignition quality are burned.</div>

Piezoelectric initiators are a unique form of ignition for energetic material because the current and voltage are tied together by impact loading on the crystal. This study examines the ignition response of an energetic composite composed of aluminum and molybdenum trioxide nanopowders to the arc generated from a lead zirconate and lead titanate piezocrystal. The mechanical stimuli used to activate the piezocrystal varied to assess ignition voltage, power, and delay time of aluminum–molybdenum trioxide for a range of bulk powder densities. Results show a high dielectric strength leads to faster ignition times because of the higher voltage delivered to the energetic. Ignition delay is under 0.4 ms, which is faster than observed with thermal or shock ignition. Electric ignition of composite energetic materials is a strong function of interparticle connectivity, and thus the role of bulk density on electrostatic discharge ignition sensitivity is a focus of this study. Results show that the ignition delay times are dependent on the powder bulk density with an optimum bulk density of 50%. Packing fractions and electrical conductivity were analyzed and aid in explaining the resulting ignition behavior as a function of bulk density.

In order to better understand how future candidate diesel fuels may affect combustion characteristics in diesel engines, 21 pure component hydrocarbon fuels were tested in a single-cylinder diesel engine. These pure component fuels included normal alkanes (C6–C16), normal primary alkenes (C6–C18), isoalkanes, cycloalkanes/-enes, and aromatic species. In addition, seven fuel blends were tested, including commercial diesel fuel, U.S. Navy JP-5 aviation fuel, and five Fischer–Tropsch synthetic fuels. Ignition delay was used as a primary combustion metric for each fuel, and the ignition delay period was analyzed from the perspective of the physical delay period followed by the chemical delay period. While fuel properties could not strictly be varied independently of each other, several ignition delay correlations with respect to physical properties were suggested. In general, longer ignition delays were observed for component fuels with lower liquid fuel density, kinematic viscosity, and liquid-air surface tension. Longer ignition delay was also observed for component fuels with higher fuel volatility, as measured by boiling point and vapor pressure. Experimental data show two regimes of operation: For a carbon chain length of 12 or greater, there is little variation in ignition delay for the tested fuels. For shorter chain lengths, a fuel molecular structure is very important. Carbon chain length was used as a scaling variable with an empirical factor to collapse the ignition delay onto a single trend line. Companion detailed kinetic modeling was pursued on the lightest fuel species set (C6) since this fuel set possessed the greatest ignition delay differences. The kinetic model gives a chemical ignition delay time, which, together with the measured experimental ignition delay, suggests that the physical and chemical delay period have comparable importance. However, the calculated chemical delay periods capture the general variation in the overall ignition delay and could be used to predict the ignition delay of possible future synthetic diesel fuels.

A new method for pulse rocket motor designing is proposed in this paper, on basis of analyzing flow, heat transfer, ignition, combustion and extinction of a kind of nonNewtonian pasty propellant, in which oxidizer and fuel are mixed homogeneously. The new motor is of the advantage of solid propellant rocket motor and liquid propellant rocket motor both. This kind of pulse rocket motor is also thrust adjustable. Moreover, because propellant is ignited by combustion chamber surplus heat, no additional ignition energy is needed. The key technology of this new method is the designing of the re-ignitable igniter. A 2D flow and heat transfer model of propellant inside the reignitable igniter is established, control equations are solved numerically. Temperature distribution of propellant inside re-ignitable igniter during its flow to combustion chamber is obtained, in conjunction with a simple ignition criterion. Relationship among igniter temperature, igniter length, igniter diameter and drive pressure are also obtained. A set of prototype test apparatus is also built for multi-pulse firing. Some seven operation pulses are obtained, each lasts about five seconds, time interval among each pulse is about two second. * Associate Professor, The Fourth Academy of China Aerospace Corporation. ALAA member. + Professor and President, The Fourth Academy of China Aerospace Corporation. Copyright © 1998 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Nomenclature A section area of the tube k consistency coefficient L length of igniter tube P drive pressure m mass n propellant flow exponent Q propellant flow mass rate R radius of igniter tube r propellant burning rate T temperature Tig propellant ignition temperature t time tid ignition delay time tf propellant flow time u velocity x, y Catherine coordinate P propellant density /I heat conduct coefficient T shear stress T] propellant viscous coefficient u dynamic viscous coefficient