scholarly journals Theoretical Analysis of Ammonium-perchlorate Based Composite Propellants with RDX Containing Small Size Particles of Beryllium

2020 ◽  
Author(s):  
Paulo Alexandre Rodrigues de Vasconcelos Figueiredo ◽  
Francisco Miguel Ribeiro Proença Brojo
Keyword(s):  
Ammonium Perchlorate ◽  
Rocket Engine ◽  
Specific Impulse ◽  
Combustion Products ◽  
Rocket Engines ◽  
Oxygen Balance ◽  
Metallic Fuels ◽  
Ignition And Combustion ◽  
Operational Envelope ◽  
Propellant Rocket

Rocket engines have been developed for at least the last six decades. There is a need to improve the actual solid propellant grain for rocket engines through the addiction of metallic fuels in the mixture as well as the addiction of energetic binders to stabilize the combustion. The rocket industry expects the launchers to be reliable, to be faster, stable and to have longer times of operation for the most possible payload weight (operational envelope). New propellants should have optimized ignition and combustion time rates reducing the possibility of negative oxygen balance thus reducing detonation process. Deflagration process should be optimized for best performance of the rocket. In this evolution, small quantities of explosives have been used in the propellant in order to increase the operational burning time, hence, the specific impulse. Adding metallic fuels such as aluminum, boron or beryllium on double based composite propellants and ammonium perchlorate are expected to increase the propellant density over chemical stability and aging resistance. The study of heterogeneous propellants containing large amounts of fine beryllium and ammonium perchlorate,   it is necessary to understand the combustion products so to a proper evaluation of specific impulse, Mach number and mass flow of the mixture. In this study a mixture with nitramides (RDX – Cyclotrimethylene trinitramide) and ammonium perchlorate was analyzed with and without the addiction of small size particles of beryllium using a numerical algorithm. Therefore, this study relates the influence of beryllium  in the performance parameters of ammonium perchlorate based composite propellants. Keywords: Propellant, Rocket engine, RDX, Ammonium perclorate

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

2019 ◽  
Author(s):  
V.A. Sorokin ◽  
O.V. Mokretsova ◽  
P.V. Valuy ◽  
D.Yu. Fedorov ◽  
A.N. Molodtsov ◽  
...  
Keyword(s):  
Solid Propellant ◽  
Rocket Engine ◽  
Combustion Products ◽  
Rocket Engines ◽  
Additional Energy ◽  
Ramjet Engine ◽  
Structure Reliability ◽  
Solid Propellant Rocket ◽  
Solid Propellant Rocket Engine ◽  
Propellant Rocket

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.

Large Rocket Engines for Space Vehicles and Missiles

1961 ◽  
Vol 65 (605) ◽  
pp. 321-331
Author(s):  
S. K. Hoffman
Keyword(s):  
Rocket Engine ◽  
Space Vehicles ◽  
Rocket Engines ◽  
Liquid Propellant ◽  
New Developments ◽  
Engine Component ◽  
Space Projects ◽  
History Of ◽  
Liquid Propellant Rocket Engine ◽  
Propellant Rocket

SummaryThe history of Rocketdyne's activity in the field of large liquid-propellant rocket engines is outlined in a chronological review of applicable United States ballistic and space projects. Within security limitations, major rocket engine component improvements and general fabrication techniques are discussed. The trends and new developments in liquid-propellant rocket engine designs are presented and a forecast of future engines is made.

Research of Faulty Sensors Data Reconstructed for Liquid-Propellant Rocket Engines

2012 ◽  
Vol 229-231 ◽  
pp. 1449-1453 ◽  
Author(s):  
Yan Jun Li ◽  
Xiao Hui Peng ◽  
Yu Qiang Cheng ◽  
Jian Jun Wu
Keyword(s):  
Fuzzy Inference System ◽  
Fuzzy Inference ◽  
Rocket Engine ◽  
Fuzzy Model ◽  
Rocket Engines ◽  
Liquid Propellant ◽  
Inference System ◽  
Neuro Fuzzy ◽  
Faulty Sensors ◽  
Propellant Rocket

In this paper, the data of faulty sensors reconstruct algorithm of liquid-propellant rocket engine is developed based on adaptive neuro-fuzzy inference system. First, the input parameters selected for method is according to regularity criterion and the relationships between each parameter; second, adaptive neuro-fuzzy inference system is train by normal test, finally, the fuzzy mode is validated by normal data and the data of faulty sensor is reconstructed. The results indicate that this algorithm can reconstruct the data of faulty sensors accurately and show that the fuzzy model approach has good performance in faulty sensors data reconstruct for LRE.

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

On the feasibility of developing a single-stage acceleration/deceleration unit based on the oxygen-hydrocarbon engine 11D58M for a super-heavy launch vehicle

2021 ◽  
pp. 62-75
Author(s):  
Boris A. SOKOLOV ◽  
Nikolay N. TUPITSYN ◽  
Evgeniy N. TUMANIN ◽  
Igor A. KRYUKOV ◽  
Andrey V. KISELEV ◽  
...  
Keyword(s):  
Rocket Engine ◽  
Launch Vehicle ◽  
Single Stage ◽  
Rocket Engines ◽  
Liquid Propellant ◽  
Gravity Assist ◽  
Upper Stage ◽  
Gravity Assist Maneuver ◽  
Lunar Gravity Assist ◽  
Propellant Rocket

The paper presents results of unsolicited exploratory design studies done by the authors into the feasibility of developing for a super-heavy launch vehicle a single-stage oxygen-hydrocarbon acceleration/deceleration unit (ADU) with two liquid-propellant rocket engines 11D58M developed by RSC Energia, intended for insertion of manned spacecraft into lunar orbit, as well as for insertion of super-heavy spacecraft into geostationary orbit (including the orbital module high-apogee transfer profile using lunar gravity assist maneuver). It demonstrates that the single-stage ADU will have a number of important advantages over both a single-stage oxygen-hydrogen ADU and a functionally similar two-stage acceleration/deceleration system of an orbital module in the form of a tandem stack of an oxygen-hydrogen acceleration stage and correction and braking stage. To assure the start-ups of the main liquid propulsion system of the ADU, it proposes a new method for inertial propellant component phase separation in the tanks in zero-gravity environment using a pre-startup pre-programmed ullage separation turn maneuver of the orbital unit about its transverse axis of inertia. Key words: Integrated launch vehicle, launch vehicle, orbital module, upper stage, orbital transfer vehicle, acceleration/deceleration unit, ullage maneuver, liquid-propellant rocket engine.

scholarly journals Comparison of cylindrical and non-cylindrical grain internal ballistic behavior of hybrid rocket engines and solid rocket motors

2021 ◽  
Author(s):  
Zhiliu Lu
Keyword(s):  
Specific Impulse ◽  
Propulsion System ◽  
Simulation Program ◽  
Rocket Engines ◽  
Hybrid Rocket ◽  
Ballistic Performance ◽  
Flow Process ◽  
Rocket Motors ◽  
Internal Ballistic ◽  
Propellant Rocket

Hybrid rocket engines (HREs) are a chemical propulsion system that nominally combine the advantages of liquid-propellant rocket engines (LREs) and solid-propellant rocket motors (SRMs). HREs in some cases can have a higher specific impulse and better controllability than SRMs, and lower cost and engineering complexity than LREs. For HREs and SRMs, both kinds of rocket engine employ a solid fuel grain, and the chosen grain configuration is a crucial point of their design. Different grain configurations have different internal ballistic behavior, which in turn can deliver different engine performance. A cylindrical grain design is a very common design for SRMs and HREs. A non-cylindrical-grain is a more complex grain configuration (than cylindrical) that has been used in many SRMs, and is also a choice for some HREs. However, while an HRE and an SRM can employ the same fuel grain configuration, the resulting internal ballistic behavior would not be expected to be the same. Pressure-dependent burning tends to dominate in SRMs, while axial flow-dependent burning tends to dominate in HREs. To help demonstrate in a more direct manner the influence of the differing combustion processes on the same fuel grain configuration used by an HRE and SRM, a number of internal ballistic simulations are undertaken for the present study. For the reference SRM cases looked at, an internal ballistic simulation program that has the capability of predicting head-end pressure and thrust as a function of time into a simulated firing is utilized for the present investigation; for the corresponding HRE cases, a simulation program is used to simulate the burning and flow process of these engines. For the present investigation, the two simulation programs are used to simulate the internal ballistic performance of various HREs and SRMs employing comparable cylindrical and non-cylindrical fuel grain configurations. The predicted performance results, in terms of pressure and thrust, are consistent with expectations that one would have for both propulsion system types.

scholarly journals On the question of solid-propellant rocket engine design preventing unstable operation in the combustion chamber

2017 ◽  
pp. 63-72
Author(s):  
G. A. Glebov ◽  
S. A. Vysotskaya
Keyword(s):  
Combustion Chamber ◽  
Solid Propellant ◽  
Rocket Engine ◽  
Rocket Engines ◽  
Pressure Oscillations ◽  
Engine Design ◽  
Solid Propellant Rocket ◽  
Solid Propellant Rocket Engine ◽  
Flow Duct ◽  
Propellant Rocket

The paper presents results of a numerical investigation concerning the effect that the flow duct shape and combustion rate equation have on the gas dynamic vortex flow pattern and self-excited pressure oscillations in the combustion chamber of a solid-propellant rocket engine. We provide guidelines on upgrading solid-propellant rocket engines in order to decrease the magnitude of pressure pulses in the case of pulsating combustion.

scholarly journals Factors Affecting Characteristic Length of the Combustion Chamber of Liquid Propellant Rocket Engines

2019 ◽  
Vol 38 (3) ◽  
pp. 729-744
Author(s):  
Tajwali Khan ◽  
Ihtzaz Qamar
Keyword(s):  
Combustion Chamber ◽  
Characteristic Length ◽  
Rocket Engine ◽  
Droplet Diameter ◽  
Computer Code ◽  
Rocket Engines ◽  
Fuel Droplet ◽  
Liquid Propellant ◽  
Proposed Model ◽  
Propellant Rocket

Optimum characteristic length of the combustion chamber of liquid rocket engine is very important to get higher energy from the liquid propellants. Characteristic length is defined by the time required for complete burning of fuel. Combustion reactions are very fast and combustion is evaporation dependent. This paper proposes fuel droplet evaporation model for liquid propellant rocket engine and discusses the factors which can affect the required size of characteristic length of the combustion chamber based on proposed model. The analysis is performed for low temperature combustion chamber. A computer code based on proposed model is generated, which solve analytical equations to calculate combustion chamber characteristic length under various input conditions. The analysis shows that characteristic length is affected by combustion chamber temperature, pressure, fuel droplet diameter, chamber diameter, mass flow rate of propellants and relative velocity of the droplet in the combustion chamber.

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