The design of liquid propellant rocket engine (LPRE) is a very complicated process; this is due to two main concerns: First, the requirements to satisfy the issues of performance, stability and compatibility. Second, the complicated, interacting processes inside thrust chamber. In this paper, an attempt to illustrate the importance of different parameters affecting performance, stability and compatibility is performed, followed by extensive study of processes inside thrust chamber. The result of processes study is developing the concept of “rate limiting process” which means that the process that can be considered the most important hence the design can be done mainly by considering it alone. This is done by developing a 1D vaporization-controlled model with its application to two case studies to illustrate model validation and application. It was found that the 1D model is valid as long as the vaporization process is the slowest process in this case the error in computing chamber cylindrical length is ∼15%. However, if the mixing process is slow, or the reaction process in gas phase is slow as in the second case study of RFNA/Tonka250 case, the error grow and may reaches 50%

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The hybrid combustion phenomenon is more complex than the case of solid or liquid combustion. An understanding of hybrid combustion is dependent upon understanding of the interrelationships between the boundary layer zones characteristics and fuel grain geometry during burning time and along fuel grain. The main objectives of the article are to predict hybrid combustion boundary layer geometry and characteristic parameters variation with fuel grain length during combustion. Specially, the relation between flame zone (propagation) position and regression rate change values during combustion. Based on this mathematical model, a computer code was implemented to identify the hybrid combustion boundary layer (BL) parameters. The comparison between Schlieren photograph of Plexiglas-oxygen flat plate burner and computational results show good agreement. The presented code can be considered as a powerful tool for the design, analysis and can be used to evaluate quantitatively the effect of changes in various design parameters of hybrid rocket motor.

Low Earth Orbit satellite(LEO) missions have potentially attractive features such as low launch cost, and wide remote sensing applications. There are three vital missions during the satellites life time including orbit transfer from the parking to the operating orbit, maintaining the operating orbit and de-orbiting mission to a disposal orbit at the end of the satellite lifetime according to the standard disposal orbit law. The mathematical model has been implemented to study various LEO mission orbits including the three vital missions. The mathematical model evaluates the required performance parameters, velocity budget, number of manoeuvers, duration and consumed propellant mass for each mission. The mathematical algorithm is carried out using Matlab/Simulink, the complete mathematical flight dynamic model has been verified through comparisons with the generated mission scenarios from Satellite Tool kit (STK), the graphical user interface is designed to display the model. The carried out algorithm can be applied in satellite mission design to help predicting the parameters for each mission to minimize the risks and intensive tests. it can be considered as powerful tool to design and analysis LEO satellite missions.

The objective of this research is to design, build, and test about a 5N Hydrogen Peroxide(H2O2) monopropellant thruster (MPT). It is utilized in remote sensing satellites for attitude control and orbit manoeuvres. The MPT uses high test peroxide (HTP) of 85 % concentration. Firstly, H2O2 ≈85% concertation by weight is prepared in the laboratory. A distillation and filtration units are built. The distillation and filtration processes are performed. Next, the design of the monopropellant thruster is done based on the developed mathematical model using NASA CEA rocket performance code. The test facility is developed which consists of the thruster, the feeding system, static test stand and data acquisition system with measuring sensors. An experimental test stand is designed and fabricated with Pendulum thrust mechanism for measurements of thrust. Finally, the silver catalyst is prepared and packed inside the MPT chamber where silver screens of high purity 99.96 % are used. The 10-firing tests are conducted under atmospheric conditions. The firings performed without heating are not completely successful. The analysis of the results shows that the thruster has a thrust range from 3.8-4.2 N. The performance of thruster starts to decay after consuming 6 kg of stabilized H2O2. The specific impulse (Is) is evaluated to be ≈93-97s at decomposition pressure of ≈10 bars and mass flow rate conf≈4.18 g/s. The performance evaluation is judged to be successful. However, using the whole potential of the 85% concentrated H2O2, is expected to increase (Is) up to ≈111.5s.