Gas dynamics of combustion-chamber processes in solid-fuel rocket motors

1992 ◽  
Vol 35 (8) ◽  
pp. 769-776
Author(s):  
E. A. Kozlov ◽  
A. B. Vorozhtsov ◽  
S. S. Bondarchuk
Keyword(s):  
Combustion Chamber ◽  
Solid Fuel ◽  
Gas Dynamics ◽  
Rocket Motors

Variability analysis of ABS solid fuel manufactured by fused deposition modeling for hybrid rocket motors

2021 ◽  
Vol 15 (2) ◽  
pp. 8029-8041
Author(s):  
Jose Alejandro Urrego ◽  
Fabio Arturo Rojas ◽  
Jaime Roberto Muñoz
Keyword(s):  
Combustion Chamber ◽  
Solid Fuel ◽  
Fused Deposition Modeling ◽  
Rocket Engine ◽  
Acrylonitrile Butadiene Styrene ◽  
Hybrid Rocket ◽  
Rocket Motors ◽  
Fused Deposition ◽  
Manufacture Process ◽  
Direct Digital Manufacturing

The process of fused deposition material (FDM) was used to manufacture propellant grains of Acrylonitrile Butadiene Styrene (ABS) as novel rocket fuel grain, with three types of geometry in the burning port. These solid fuel grains were used to measure the typical characteristics of combustion in rocket motors such as thrust and pressure inside the combustion chamber, seeking to obtain preliminary characteristics of operation and analyze the effect of combustion port geometry on pressure and thrust, using Multivariate Analysis of Variance (MANOVA) as statistical method. Two of the three geometries were manufactured with a helical-finocyl configuration, specially designed to be fabricated by Direct Digital Manufacturing (DDM), the other one was a straight-bore geometry also by DDM. This characterization experiment was performed on a static hybrid rocket engine, designed to inject 99.98% pure nitrous oxide into a combustion chamber with capacity to withstand 6.9 MPa of pressure, with an easy-to-exchange nozzle, avoiding erosive behavior in the throat. Statistical analyses made with the ABS fuel grains, suggest a significant effect on rocket motor pressure and thrust, due to helical geometric changes made to the combustion port of solid fuel grains made by FDM manufacture process.

scholarly journals Challenges of Ablatively Cooled Hybrid Rockets for Satellites or Upper Stages

Aerospace ◽  
2021 ◽  
Vol 8 (7) ◽  
pp. 190
Author(s):  
Francesco Barato
Keyword(s):  
Combustion Chamber ◽  
Solid Fuel ◽  
Diameter Ratio ◽  
Propulsion System ◽  
Burning Time ◽  
Regression Rate ◽  
Hybrid Propulsion ◽  
Deep Impact ◽  
Time Ratio ◽  
Hybrid Rockets

Ablative-cooled hybrid rockets could potentially combine a similar versatility of a liquid propulsion system with a much simplified architecture. These characteristics make this kind of propulsion attractive, among others, for applications such as satellites and upper stages. In this paper, the use of hybrid rockets for those situations is reviewed. It is shown that, for a competitive implementation, several challenges need to be addressed, which are not the general ones often discussed in the hybrid literature. In particular, the optimal thrust to burning time ratio, which is often relatively low in liquid engines, has a deep impact on the grain geometry, that, in turn, must comply some constrains. The regression rate sometime needs to be tailored in order to avoid unreasonable grain shapes, with the consequence that the dimensional trends start to follow some sort of counter-intuitive behavior. The length to diameter ratio of the hybrid combustion chamber imposes some packaging issues in order to compact the whole propulsion system. Finally, the heat soak-back during long off phases between multiple burns could compromise the integrity of the case and of the solid fuel. Therefore, if the advantages of hybrid propulsion are to be exploited, the aspects mentioned in this paper shall be carefully considered and properly faced.

An Accurate Performance Prediction of Solid Fuel Ramjets Using Coupled Intake-Combustor-Nozzle Simulation

2020 ◽  
Vol 36 (6) ◽  
pp. 933-941
Author(s):  
A. M. Tahsini
Keyword(s):  
Combustion Chamber ◽  
Solid Fuel ◽  
Solid Phase ◽  
Three Dimensional ◽  
Flow Fields ◽  
Reacting Flow ◽  
Computational Domain ◽  
Regression Rate ◽  
Fuel Flow ◽  
Accurate Performance

ABSTRACTThe performance of the solid fuel ramjet is accurately predicted using full part simulation of this propulsion system, where the flow fields of the intake, combustion chamber, and the nozzle are numerically studied all together. The conjugate heat transfer is considered between the solid phase and the gas phase to directly compute the regression rate of the fuel. The finite volume solver of the compressible turbulent reacting flow is utilized to study the axisymmetric three dimensional flow fields, and two blocks are used to discretize the computational domain. It is shown that the combustion chamber's pressure is changed due to the fuel flow rate's increment which must be taken into account in predictions. The results demonstrate that omitting the pressure dependence of the regression rate and also the effect of the combustor's inlet profile on the regression rate, which specially exists when simulating the combustion chamber individually, under-predicts the solid fuel burning rate when the regression rate augmentation technique is applied to improve the performance of the solid fuel ramjets. It is also illustrated that using the inlet swirl to increase the regression rate of the solid fuel augments considerably the thrust level of the considered SFRJ, while the predictions without considering all parts of the ramjet is not accurate.

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

2021 ◽  
pp. 98-107
Author(s):  
T.S. Sultanov ◽  
G.A. Glebov
Keyword(s):  
Combustion Chamber ◽  
Condensed Phase ◽  
Solid Fuel ◽  
Solid Propellant ◽  
Rocket Engine ◽  
Specific Impulse ◽  
Two Phase ◽  
Solid Propellant Rocket ◽  
Solid Propellant Rocket Engine ◽  
Propellant Rocket

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

Modelling of a rotating detonation engine combustion chamber

2021 ◽  
pp. 33-49
Author(s):  
Елена Викторовна Михальченко ◽  
Валерий Федорович Никитин ◽  
Любен Иванович Стамов ◽  
Юрий Григорьевич Филиппов
Keyword(s):  
Boundary Conditions ◽  
Combustion Chamber ◽  
Chemical Reactions ◽  
Gas Dynamics ◽  
Turbulent Transport ◽  
Kinetic Equations ◽  
Problem Solution ◽  
Engine Combustion ◽  
Detonation Engine ◽  
Rotating Detonation

Рассмотрено трехмерное численное моделирование камеры сгорания двигателя с непрерывной детонационной волной с помощью авторского программного пакета. Программное обеспечение использует для многокомпонентного химически реактивного газа математическую модель с опциональным подключением модели турбулентности. В основе модели химической кинетики лежит механизм элементарных реакций, в зависимости от механизма меняется число реакций. В программе, в том числе, реализован авторский кинетический механизм. Рассмотрены шесть кинетических механизмов: Мааса-Варнаца-Поупа, Хонга, Вильямса, Gri-Mech 3.0, Ли-Джоу-Казакова-Драера и авторский, проведено их сравнение. Код распараллелен с помощью технологий OpenMP и MPI. В результате работы программы получена оптимальная форма камеры сгорания с самоподдерживающейся детонационной волной на смеси водорода с кислородом. Purpose. To create software for studying the features of the transition from ignition and deflagration to a detonation mode in a three-dimensional configuration, including the formation and propagation of a rotating detonation complex, which takes transient processes into account. Methodology. The software is based on a mathematical model for multi-component gas dynamics with chemical reactions and turbulent transport for diffusion, viscosity, and thermal conductivity. High-order calculation schemes are used. To solve a stiff subsystem of kinetic equations, a hybrid implicit-explicit Novikov method is used (a specific variant of a Rosenbrock method). Findings. Authors created a code which calculates physical processes within a multi-component gas dynamics paradigm. The code accounts for chemical processes and turbulence modelling. The shape of computation domain and the type of boundary conditions is user defined. These include boundary conditions at the wall, as well as inflow and outflow conditions for both subsonic, and supersonic modes. Initial conditions can be set up differently in different regions of the domain. The software consists of several modules: a mesh-building module, initial state creation, calculation of new time layers saving the intermediate and final results at control points with a possibility to resume interrupted calculations, and post-processing modules. Authors developed blocks of solutions for various elementary chemical kinetic mechanisms, one of considered mechanisms is build up by themselves, others are published previously. It was obtained that the details of the 3D transient problem solution significantly depend on the chosen mechanism. Оriginality/value. The software complex makes it possible to process numerical modelling of a detonation engine combustion chamber in a 3D configuration considering chemical reactions and turbulent transport. Different chemical kinetics mechanisms are utilizable, and thrust characteristics could be obtained.

The response of the solid fuel ramjet combustor to the inflow excitations

2021 ◽  
pp. 095441002110534
Author(s):  
AmirMahdi Tahsini ◽  
Seyed Saeid Nabavi
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Solid Fuel ◽  
Ambient Pressure ◽  
Excitation Frequency ◽  
Engine Performance ◽  
Computational Domain ◽  
Rocket Motors ◽  
Fuel Mass

The response of the solid fuel ramjet to the imposed excitations of the ambient pressure is investigated using full part computation of the system including the intake, combustion chamber, and exhaust nozzle. The finite volume solver of the turbulent reacting compressible flow is used to simulate the flow field, where two grid blocks are considered for discretizing the computational domain. Both impulsive and oscillatory excitations are imposed to predict the response of the solid fuel mass flow rate. The results demonstrate that strong fuel flow overshoot occurs in the case of sudden impulsive excitation which is omitted for gradual impulsive excitations. In addition, the oscillatory excitations eventually lead to regular oscillatory response with frequencies similar to the imposed excitations and decrease the average fuel mass flow rate independent of the excitation frequency. But the amplitude of the response depends on the excitation frequency and amplification occurs in some frequencies. This behavior is not related to the combustion instabilities and is similar to the L-star instability in the solid rocket motors. In the design and analysis of the solid fuel ramjets, the coupling of the flight dynamics and the engine performance must be considered, and this study is the first step of such complete methodology to have more accurate predictions.

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