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During the spacecraft (SC) mission, some of orbital elements change during the flight, the need to study the spacecraft maneuvers become paramount importance need. A full mathematical model for the Spacecraft lifetime maneuvers is deduced, implemented in a simulation scenario with a GUI, and then desired simulation is verified during the different operations' phases such as orbit transfer, propagation, maintenance and deorbiting. In the current article, an interpretation of the carried-out model will be introduced to match a real case of mission analysis as possible for newly designed satellite. Two types of orbits (inclined and Sun synchronous Orbit SSO) are investigated for the SC mission analysis over the whole lifetime phases. The investigation is carried out using two types of power systems (electrical and chemical) using the same propellant tank model with capacity 60 kg. A numerical comparison was concluded for both studied power systems.

Hybrid rocket motor propellant offers a number of advantages over the traditional liquid and solid propellants, particularly in the areas of safety, cost and environmental effects. The combustion mechanism and internal ballistics of hybrid rocket motor (HRM) differs markedly from those in liquid engine and solid motor. Prediction of internal ballistic parameters of HRM is greatly affected by the instantaneous port geometry of the solid fuel grain during firing and types of oxidizer and fuel.This paper discusses the design and selection criteria of HRM propellant (solid fuel, gas or liquid oxidizer and energetic additives), fuel grain geometry and operating parameters. The three basic elementary geometries namely, tube, pie-shaped, and triangular ports are used to describe more complicated hybrid grain configurations. Consequently burning perimeter, effective port area, fuel grain cross section area, filling coefficient (greater than 50%) and sliver ratio (about 6%) can be calculated.Comprehensive static tests for small scale hybrid rocket motor have been carried out for various solid fuel materials (PMMA, PE, PE+AL, paraffin, bees wax) with gas oxygen as oxidizer. Finally, the designer selects a propellant type that is likely to give the required performance, considering cost, mechanical properties as well as the necessary storage stability and the best safety properties (non-toxic and smokeless).

Many phenomena occur during Solid Propellant (SP) Rocket Motor (RM) firing, one of the most important is known as SP erosive burning, which occurs at the end of grain, during the earliest period of firing after ignition. Calculating and analyzing erosive burning rate are an essential requirement for evaluating SPRM performance parameters. Intensive theoretical investigations are performed for the prediction of erosive burning and RM internal ballistic parameters. Investigations help in minimizing experimental work, which is expensive, risky and time consuming.This paper presents a SPRM erosive burning computer code with its mathematical simulation, based on heat transfer theory of Lenoir and Robillard. Results of modeling are compared with the results of experimental study and a satisfactory agreement is indicated by an error less than 5%. The experimental work is based on full-scale firing of artillery RM that has higher propellant filling coefficient (volumetric loading).A comprehensive analysis of the erosive burning phenomena and simplified formula are presented. The presented simulation code can be considered as a powerful tool for design parameters and analysis of the erosive burning rate for SPRM.

The main vital purpose of gas generator (GG) solid propellant (SP) is, the burning of SP to produce the required gas amount with low temperature and without solid precipitate. Numerous variety of propellants and configurations have been used to create hot gas. A comprehensive study of several important aspects of SPGG design has been investigated in the present work. End burning mechanism, GG solid propellant types, various burning surface configurations, simplified design procedure, insulation evaluation, analysis of experimental works, numerical models have been developed to describe pressure variation with time and performance parameters. Effect of design parameters were also discussed. Good agreement is obtained between experimental firing data and the established models for SPGG parameters performance. Finally; the several design parameters of the SPGG are analyzed and investigated, each design parameter is treated separately while the other parameters are assumed constant.

Hybrid rocket solid fuel grain regression rate is one of the significant values to accurately determine in the hybrid rocket motor (HRM) design parameters and performance prediction. The comprehensive method to measure and evaluate HRM solid fuel regression rate was investigated theoretically and experimentally .The mathematical model has been introduced and a computer program was built to describe HRM regression rate. A series of experimental tests using a small scale HRM employs Polyethylene (PE) and aluminized PE as solid fuel (up to 15% by mass metal powder additive) with different initial port diameter of fuel grain (5, 20 and 28mm) and gaseous oxygen (GO2) as oxidizer to evaluate the HRM regression rate performance. The measured experimental data are compared with results obtained from theoretical computation. The comparison shows good agreement, which proves the validity of the developed program. Regression rate of hybrid fuel grain was enhanced by addition of Al powder compared to basic configuration (0% Al) and it was found that it is more significant to use Al powder additive increased up to 7.5% by mass

In recent years, mono-propellant propulsion has attracted widespread attention in space due to excessive benefits of safety, low cost, simplicity, clean exhaust and high reliability. A 5N monopropellant Hydrogen Peroxide (H2O2) thruster is selected to be utilized with remote sensing (RS) satellites for attitude control and orbit correction. In the present paper, comprehensive theoretical analysis and design parameters are investigated. The effects of change in various design parameters are studied, such as operating pressures (7-16 bar), H2O2 concentration (70-95%) and initial temperature (25-70℃). The crucial design performance of thruster has been carried out by using thermo-chemical software package. A conceptual design was conducted to determine the dimensions of the thruster prototype including the injector plate, catalytic bed, retainer and nozzle part. 70℃ H2O2 preheating temperature might be needed for catalytic bed fast response. The results are described in the form of graphs and tables for the quick and easy analysis.

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%

For many reasons, electrical propulsion systems are the most expected in space; they are suitable for researchers because they are effective, easy to design and applicable in space. One of the most important parameters that could assess the performance of the propulsion system is the problem of measuring the very low thrust of pulsed plasma thrusters (PPTs). Various test beds, including indirect designs of measuring systems such as pendulums and optical measuring systems, can measure the thrust with very high precision, but it is very complicated. This paper presents the firing test of a new designed coaxial pulsed plasma thruster for proofing the concept and then the thrust is measured by a new simple test bed using a rod in accordance with a loading cell (0-100 gf.) which can measure thrust accurately after good calibration up to μgf. In a medium-sized vacuum chamber (VC) (10−5 bar, D= 1 m and L= 1.4 m), a coaxial PPT is designed and tested under environmental conditions in space. Four proofs of concept took place until the experiment was adjusted and six experiments were carried out to measure the thrust using the new test bed under various operating conditions. Evaluation of the system is presented too and the maximum thrust obtained is 0.68mN at the discharge voltage equals 7kV and the 12kV voltage of the spark plug initiator and the efficiency equals 10%. The steps of fabrication of the coaxial PPT and steps of measuring the thrust is presented in the paper beside the result charts of the thrust under different operating conditions.

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 the present paper is to prove the validity design and fully automatically operation of HPS, seeking to achieve its safety, low cost, and flexibility. In the stage of preliminary design, a coupling between experimental and theoretical work has been done, making use of available laboratory facilities and software packages (thermo-chemical calculation), to fully demonstrate all operating conditions. The HPS design includes the solid fuel grain configuration, essential operating parameters, oxidizer feed through control devices, measuring tools with calibration system, data recording and operating system. Commercial fuel materials as Polymethyl methacrylate (PMMA) and Polyethylene (PE) are used during experimental work to represent conventional generation fuel. Furthermore, new generation fuels as paraffin, beeswax, and paraffin with energetic additive metal (AL powder) are used. Gaseous oxygen is used as oxidizer. That may guide to design and fabricate high performance real rocket weapon or space propulsion system, targeting longer term storage and a propulsion system with available technology. The objective is extended to select appropriate ignition system (NiCr coil igniter) for laboratory firing of small-scale engine. One of the main problems (price, safety, available technology) of propulsion system today has been tackled. To that end, analysis has been established, to investigate the regression rate, solid fuel mass flow rate and combustion efficiency for each fuel grain group types.

This paper presents a study on the trajectory of a two stage launch vehicle. The study includes dynamic responses of motion parameters as well as the variation of angles affecting the orientation of the launch vehicle (LV). LV dynamic characteristics including state vector variation with corresponding altitude and velocity for the different LV stages separation, as well as the angle of attack and flight path angles are also discussed. A flight trajectory study for the drop zone of first stage and the jettisoning of fairing are introduced in the mathematical modeling to study their effect. To increase the accuracy of the LV model, atmospheric model is used taking into consideration geographical location and the values of solar flux related to the date and time of launch, accurate atmospheric model leads to enhancement of the calculation of Mach number, which affects the drag force over the LV. The mathematical model is implemented on MATLAB based software (Simulink). The real available experimental data are compared with results obtained from the theoretical computation model. The comparison shows good agreement, which proves the validity of the developed simulation model; the maximum error noticed was generally less than 10%, which is a result that can lead to future works and enhancement to decrease this level of error., {"references":["Roger R., Donald D., Jerry E., \"Fundamentals of Astrodynamics\", Dover publications, New York, 1971.","Kirsanov V. G., \"Tutorial on Ballistics to do the academic year work for post graduate students \", Moscow aviation institute, 1981.","K. H. WELL, \"Neighboring Vehicle Design for a Two-Stage Launch Vehicle\" University of Stuttgart, 70550 Stuttgart, Germany,2003.","Michael D. Griffin, James R. French, \"Space Vehicle Design\", second edition, AIAA Education Series, Reston, Virginia, 2004.","J. M. Picone, A. E. Hedin 1, and D. P. Drob, \"NRLMSISE-00 Empirical Model of the Atmosphere: Statistical Comparisons and Scientific Issues\", E. O. Hulburt Center for Space Research Naval Research Laboratory Washington, DC 20375, 2001.","Greg A. Dukeman and Ashley D. Hilly, \"Rapid Trajectory Optimization for the ARES I Launch Vehicle\", AIAA 2008-6288, AIAA Guidance, Navigation and Control Conference and Exhibit 18 - 21 August 2008, Honolulu, Hawaii, 2008.","XUAN Ying†, ZHANG WeiHua and ZHANG YuLin, \"Pseudospectral method based trajectory optimization and fairing rejection time analysis of solid launch vehicle\", Sci China Ser E-Tech Sci, vol. 52 | no. 11 | 3198-3206, 2009.","Saqlain Akhtar and He Linshu, \"Support Vector Machine Based Trajectory Metamodel for Conceptual Design of Multi-Stage Space Launch Vehicle\", School of Astronautics Beihang University, 37-XueYuan Lu 100083, Beijing, China, 2005."]}

In this paper, there is an implementation, verification, and graphical demonstration of a software application, which can be used swiftly over different preliminary orbit determination methods. A passive orbit determination method is used in this study to determine the location of a satellite or a flying body. It is named a passive orbit determination because it depends on observation without the use of any aids (radio and laser) installed on satellite. In order to understand how these methods work and how their output is accurate when compared with available verification data, the built models help in knowing the different inputs used with each method. Output from the different orbit determination methods (Gibbs, Lambert, and Gauss) will be compared with each other and verified by the data obtained from Satellite Tool Kit (STK) application. A modified model including all of the orbit determination methods using the same input will be introduced to investigate different models output (orbital parameters) for the same input (azimuth, elevation, and time). Simulation software is implemented using MATLAB. A Graphical User Interface (GUI) application named OrDet is produced using the GUI of MATLAB. It includes all the available used inputs and it outputs the current Classical Orbital Elements (COE) of satellite under observation. Produced COE are then used to propagate for a complete revolution and plotted on a 3-D view. Modified model which uses an adapter to allow same input parameters, passes these parameters to the preliminary orbit determination methods under study. Result from all orbit determination methods yield exactly the same COE output, which shows the equality of concept in determination of satellite’s location, but with different numerical methods., {"references":["Administrator, NASA. Sputnik 1|NASA. https://www.nasa.gov/multimedia/imagegallery/image_feature_924.html October 4, 2011.","Howard D. Curtis, \"Orbital Mechanics for Engineering Students\", ISBN 0-7506-61690, ELSEVIER, Embry Riddle Aeronautical University Daytona Beach, Florida, 2005.","Bate R., Mueller D. and White J., \"Fundamentals of Astrodynamics\", ISBN 0-486-60061-0, Dover Publications, Inc., New York, 1971.","Vladimir A. Chobotov, \"Orbital Mechanics Third Edition\", ISBN 1-56347-537-5, AIAA, Reston, Virginia, 2002.","The Math Works, Inc. \"MATLAB R2012b (8.0.0.783)\", copyright 1984-2012.","Nancy Teresa Cabrera. \"MATLAB® GUI Visualization of Classical Orbital Elements\", San Luis Obispo, CA, 93407, 2010.","Analytical graphics, Inc. \"Satellite Tool kit 9.1.0\", copyright 1989-2009.","Willium F. Huseonicu, \"A Numerical Evaluation of Preliminary Orbit Determination Methods\", NASA TN D-5649, WASHINGTON, D. C.,1970.","Montenbruck, Gill., \"Satellite Orbits\", ISBN 3-540-67280-X, Springer, Wessling, Germany 2000."]}