Performance Analysis of Axial-Injection, End-Burning Hybrid Rocket Propulsion System

2020 ◽  
Vol 57 (6) ◽  
pp. 1408-1413
Author(s):  
Matthew A. Hitt
Keyword(s):  
Performance Analysis ◽  
Propulsion System ◽  
Hybrid Rocket ◽  
Axial Injection ◽  
Rocket Propulsion

Design and Characterization of Optically Clear PMMA/ GOx Hybrid Rocket Propulsion System.

2021 ◽  
Author(s):  
Jordi Estevadeordal ◽  
Holton Miller ◽  
Joshua Yurek ◽  
Elliot Omerza ◽  
Al Habib Ullah ◽  
...  
Keyword(s):  
Propulsion System ◽  
Hybrid Rocket ◽  
Rocket Propulsion

scholarly journals Trade-off Evaluation due to Application of Mixing Chamber for Hybrid Rocket-Propulsion System

2016 ◽  
Vol 10 (3) ◽  
pp. 23-31
Author(s):  
Hakchul Kim ◽  
Keunhwan Moon ◽  
Heejang Moon ◽  
Jinkon Kim
Keyword(s):  
Propulsion System ◽  
Hybrid Rocket ◽  
Trade Off ◽  
Rocket Propulsion ◽  
Mixing Chamber

A Numerical Study of Coupling of Thermal and Hydrodynamic Oscillations in a Hybrid Rocket Motor

2011 ◽  
Author(s):  
K. M. Akyuzlu ◽  
K. Albayrak
Keyword(s):  
Mathematical Model ◽  
Combustion Chamber ◽  
Numerical Study ◽  
Rocket Motor ◽  
Propulsion System ◽  
Working Fluid ◽  
Hybrid Rocket ◽  
One Dimensional ◽  
Rocket Propulsion ◽  
Hybrid Rocket Motor

A one-dimensional, mathematical model is adopted to investigate, numerically, the instabilities experienced inside a hybrid rocket propulsion system. The presumption is that such oscillations feed into combustion instabilities and result in poor performance of the propulsion system and/or result in mechanical vibrations that lead to failure of the rocket motor. The model adopted for the numerical study is a one-dimensional, multi-node representation of a subscale hybrid rocket propulsion system. A one dimensional channel with circular cross-section is configured to simulate a combustion chamber of a rocket hybrid rocket motor and is connected to a converging–diverging nozzle in the downstream and to a plenum with a flow straightener in the upstream side. The working fluid is supplied from a pressurized storage tank to the upstream plenum through a throttle valve. A multi-component approach is used to model, mathematically, the propulsion system. In this integrated-component model, the unsteady flow through the throttle valve and the nozzle is assumed to be one-dimensional and isentropic whereas the flow in the forward plenum and in the combustion chamber is assumed to be a one-dimensional, unsteady, compressible, turbulent, and subsonic. The physics based mathematical model of the flow in the channel consists of conservation of mass, momentum and energy equations subject to appropriate boundary conditions as defined by the physical problem stated above. The working fluid is assumed to be compressible through a simple ideal gas relation. The governing equations of the compressible flow in the combustion chamber are discretized using the second order accurate MacCormack finite difference scheme. Convergence and grid independence studies were done to determine the optimum mesh size and computational time increment needed for the present simulations. Furthermore, steady state results of the proposed model are compared to the results of the isentropic, Fanno (viscous 1-D flow), and Rayleigh (1-D flow with heat input) case studies to verify the accuracy of the numerical predictions. Numerical experiments were then carried out to simulate the flow oscillations in the combustion chamber of a sample subscale hybrid rocket motor. Experiments were repeated for various operating conditions (Re numbers between 104 and 106) to determine the flow regions where these oscillations are sustained. The numerical simulation results indicate that the proposed mathematical model predicts the expected unsteady axial distributions of temperature, velocity, and pressure in the combustion chamber and the general characteristics of the experimentally observed instabilities associated with hybrid rocket propulsion systems.

scholarly journals Hybrid Propulsion System: Novel Propellant Design for Mars Ascent Vehicles

2021 ◽  
Author(s):  
Ozan Kara ◽  
Arif Karabeyoglu
Keyword(s):  
Low Cost ◽  
Flame Temperature ◽  
Propulsion System ◽  
Hybrid Rocket ◽  
General Audience ◽  
Hybrid Propulsion ◽  
Test Setup ◽  
Rocket Propulsion ◽  
Hybrid Motor

This chapter briefly introduces hybrid rocket propulsion for general audience. Advantageous of hybrid rockets over solids and liquids are presented. This chapter also explains how to design a test setup for hybrid motor firings. Hybrid propulsion provides sustainable, safe and low cost systems for space missions. Therefore, this chapter proposes hybrid propulsion system for Mars Ascent Vehicles. Paraffin wax is the fuel of the rocket. Propulsion system uses CO2/N2O mixture as the oxidizer. The goal is to understand the ignition capability of the CO2 as an in-situ oxidizer on Mars. CO2 is known as major combustion product in the nature. However, it can only burn with metallic powders. Thus, metallic additives are added in the fuel grain. Results show that CO2 increase slows down the chemical kinetics thus reduces the adiabatic flame temperature. Maximum flammability limit is achieved at 75% CO2 by mass in the oxidizer mixture. Flame temperature is 1700 K at 75% CO2. Ignition quenches below the 1700 K.

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