Numerical Computation of Specific Impulse and Internal Flow Parameters in Solid Fuel Rocket Motors with Two-Phase Сombustion Products

Eulerian --- Lagrangian method was used in the Fluent computational fluid dynamics system to calculate motion of the two-phase combustion products in the solid fuel rocket motor combustion chamber and nozzle. Condensed phase is assumed to consist of spherical particles with the same diameter, which dimensions are not changing along the motion trajectory. Flows with particle diameters of 3, 5, 7, 9, and 11 μm were investigated. Four versions of the engine combustion chamber configuration were examined: with slotted and smooth cylindrical charge channels, each with external and submerged nozzles. Gas flow and particle trajectories were calculated starting from the solid fuel surface and to the nozzle exit. Volumetric fields of particle concentrations, condensed phase velocities and temperatures, as well as turbulence degree in the solid propellant rocket engine flow duct were obtained. Values of particles velocity and temperature lag from the gas phase along the nozzle length were received. Influence of the charge channel shape, degree of the nozzle submersion and of the condensate particles size on the solid propellant rocket engine specific impulse were determined, and losses were estimated in comparison with the case of ideal flow

Mathematical Model of Criticality of Control System for Rotary Nozzle of Solid-Propellant Rocket Engine

К системам управления соплами ракетных двигателей предъявляются специальные требования, жесткое соблюдение которых обеспечивает расчетную работоспособность данных устройств в условиях действия повышенных нагрузок при старте ракеты. В условиях такого старта в системе управления соплом неизменно возникает и развивается критичность. Появление такой критичности присуще системам автоматического управления. Развитие критичности может привести к изменению параметров системы управления за весьма малый отрезок времени. Например, критичность возникает при отработке угла поворотным соплом крупногабаритного твердотопливного ракетного двигателя. Изменения параметров устройств при старте могут находиться как в пределах допустимых значений, так и выходить за них, в таких случаях и развивается критичность.В статье рассматривается возможность оценки критичности переходного процесса системы управления поворотного сопла ракетного двигателя в виде математической модели.Оценку критичности предполагается проводить по значениям характеристик переходного процесса. Данные характеристики входят в 12 формул математической модели. Вычисленный по формуле результат – показатель (индикатор), характеризующий процесс развития критичности. Показатели нормированы так, чтобы определить ситуацию, при которой происходит нерасчетное развитие критичности. Значение показателя, близкое и превышающее 1, указывает на критичность.На числовом примере показана методика использования математической модели для определения расчетного (с наибольшей вероятностью) развития критичности при работе поворотного сопла ракетного двигателя.

ROCKET-PROPELLED ANTI-PERSONNEL GRENADE LAUNCHERS EQUIPPED WITH DIFFERENT WARHEADS

At present, the issue of creating means of destruction of domestic production in Ukraine, including rocket- propelled anti-personnel grenade launchers equipped with different warheads (thermobaric, assault, fragmentation and incendiary) is acute. With the advent of new incendiary, thermobaric agents and explosives, the development of multi-purpose rocket-propelled grenade launchers has become one of the promising trends. All of the above is proven by the experience of the Anti-Terrorist Operation/Joint Forces Operation in eastern Ukraine. Analysis of flame troops combat operations during the anti-terrorist operation showed the need to intensify efforts to develop new rocket-propelled anti-personnel grenade launchers with thermobaric, assault, fragmentation and incendiary warheads. Leading specialists from the Russian Federation and Western countries also came to this conclusion. Thus, in recent years in Russia there have been developed, passed state tests and adopted RGSH-1 with a modular warhead in the thermobaric version, RGSH-2 rocket projectile with thermobaric warhead and RPG-32 "Barkas" equipped with thermobaric or tandem cumulative grenade. Nowadays, several enterprises of the defense complex, one of which is the ТОV "Vognyana Varta", are dealing with the creation of a modern domestic rocket-propelled anti-personnel grenade launcher in Ukraine. ТОV "Vognyana Varta" on its own initiative has been developing a rocket-propelled anti-personnel grenade launcher with different warheads using new incendiary, thermobaric mixtures and explosives. The classification of rocket-propelled grenade launchers (flamethrowers) is carried out. The design of modern ammunition to rocket-propelled grenade launchers (flamethrowers) and types of warheads for them are considered. Calculations of a solid-propellant rocket engine for modern rocket-propelled grenade launchers (flamethrowers) are given.

NUMERICAL INVESTIGATION OF UNSTEADY GASDYNAMIC PROCESSES AT THE LAUNCH OF SOLID-PROPELLANT ROCKETS

The paper presents the results of the methodology developed for calculating unsteady gasdynamic processes occurring at the launch of missiles, in the gas-dynamic paths of rocket engines, and in the external regions. The method accounts for the variation in the geometry of the solidpropellant charge in the course of solid-propellant rocket engine operation and in the geometry of the computational domain at the rocket launch. The analysis of the unsteady force impact of the supersonic jet on the launch surface is carried out. It is shown that the maximum force action is located in the vicinity of the Mach disks of the unperturbed jet. Numerical studies of gasdynamic processes at the launch of a model solid-propellant booster rocket are implemented including the case when the nozzle plug opening is taken into account. The contribution of the thrust force components at the stage of bootstrap operation is assessed. The presence of the plug at the initial stage of the engine start leads to an abrupt change in the thrust and minor fluctuations, which are damped as the pressure in the combustion chamber rises.

The Influence of Configuration on the Range of a Hypersonic Flight Vehicle under Space Limitations

The article examines the influence of the configuration of a hypersonic flight vehicle with a hori-zontal cruise flight section on the maximum flight range under the conditions of limiting dimen-sions. The ram jet and the solid propellant rocket engine were chosen as the cruise engine, and re-spective calculations were performed. The hypersonic flight vehicle was configured based on the design patented in the Russian Federation, under the condition of launch using the 3S14 universal launcher. Dependencies of the maximum range on the ratio between the launch mass to the payload were analysed. The mass efficiency indicator for a two-stage ballistic missile was chosen as a crite-rion for a comparison with other available alternatives.

Dynamics of a cylindrical shell with a collapsing elastic base under the action of a pressure wave

In the dynamic and quasi-static statements, the issue of non-stationary deformation and stability of the solid propellant rocket engine (SPRE) was approximately solved. It is modeled by a thin, smooth cylindrical shell, inside of which, on a part of its length, there is an elastic base corresponding to a gradually burning powder charge. A pressure wave is moving along the outer surface of the body, simulated by the running load. The deformed state of the shell is considered axisymmetric and is determined on the basis of the moment theory of the shells. For diverse variants of mounting the ends of the shell in a closed form, expressions were obtained for the critical velocity of the load. Examples were considered.

Analysis of the existing igniter devices and the development of a promising igniter design for the launching booster of modern solid-propellant ramjets

The article briefly reviews the existing designs and technical solutions for solid propellant igniters in rocket engines. The technical and design solutions in the development of solid-propellant rocket engine igniters are analyzed. The results of the development of a promising igniter design with gunpowder filler axially located in the internal channel of a propellant grain of a launching free-flowing booster of a rocket-ramjet engine are presented. Ensuring the required level of structure reliability and durability in the launch mode of the solid propellant rocket-ramjet engine, the igniter will improve the engine traction characteristics due to the combustion of the igniter aluminum shell as an additional energy source and using its combustion products in the stream of solid propellant combustion products.