Optimization of micro air launch vehicle for micro and nano satellite /
Air launch vehicles provide a prominent solution for low cost launch of micro and nano satellites with much less launching constraints as compared to ground launch. In this research, evolutionary optimization techniques is used to design a micro air launch vehicle (MALV) which is specifically design...
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| Main Author: | |
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| Format: | Thesis |
| Language: | English |
| Published: |
Kuala Lumpur:
Kulliyyah of Engineering, International Islamic University Malaysia,
2012
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| Subjects: | |
| Online Access: | Click here to view 1st 24 pages of the thesis. Members can view fulltext at the specified PCs in the library. |
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| LEADER | 053680000a22003130004500 | ||
|---|---|---|---|
| 008 | 120919t2012 my ad g m 000 0 eng d | ||
| 040 | |a UIAM |b eng | ||
| 043 | |a a-my--- | ||
| 050 | |a TL785.8.L3 | ||
| 100 | 1 | |a Aldheeb, Mohammed Abdulmalek | |
| 245 | 1 | |a Optimization of micro air launch vehicle for micro and nano satellite / |c by Mohammed Abdulmalek Aldheeb | |
| 260 | |a Kuala Lumpur: |b Kulliyyah of Engineering, International Islamic University Malaysia, |c 2012 | ||
| 300 | |a xx, 187 leaves : |b ill, charts ; |c 30cm. | ||
| 500 | |a Abstract in English and arabic. | ||
| 500 | |a " A thesis submitted in fulfiiment of the requirement for the degree of Master of Mechanical Engineering."--On t.p. | ||
| 502 | |a Thesis (MSc.MEC)--International Islamic University Malaysia, 2012. | ||
| 504 | |a Includes bibliographical references (leaves 163-165). | ||
| 520 | |a Air launch vehicles provide a prominent solution for low cost launch of micro and nano satellites with much less launching constraints as compared to ground launch. In this research, evolutionary optimization techniques is used to design a micro air launch vehicle (MALV) which is specifically designed for very small payloads. To perform the optimization, a two-step design optimization algorithm is developed. In the first step, a preliminary design optimization of the air launch vehicle is performed to determine its design and performance parameters. In the second step, a trajectory optimization is performed to determine optimal trajectory and estimate the velocity losses of the designed vehicle. For the first step, a preliminary design model is developed for air launch vehicles. This model is used to calculate the masses, dimensions and performance parameters of the air launch vehicle. The inputs for this model are the payload mass; launching velocity losses due to drag, steering and gravity; and initial guess values for the vehicle specific impulses and inert mass fractions. Given these values, mission analysis and preliminary design is performed to determine, for each stage of the vehicle, the geometric and performance parameters taking into consideration some important inputs such as thrust to weight ratio, length to diameter ratio, chamber pressure, type of propellant and structure materials. For the second step, a 3 degree-of-freedom trajectory simulation model is developed. The inputs of this model are the masses and performance parameters obtained from vehicle design and a predetermined set of angles of attack. Angles of attack are used to steer the vehicle along its trajectory using aerodynamic forces. The coefficients of lift and drag are obtained using Missile DATCOM, which calculates the aerodynamic coefficients within the flight envelope of the vehicle. Two evolutionary optimization techniques are used for the optimization of the air launch vehicle trajectory and design; namely particle swarm optimization (PSO) and differential evolution (DE). These techniques were selected for their ability to handle large design spaces with many dimensions while making no assumptions about the design problem itself. The optimization problem is formulated with the vehicle initial mass taken as the objective function. A set of design variables are used in vehicle design, which are thrust to weight ratio, length to diameter ration, exit area to rocket diameter ratio, chamber pressure and velocity fraction for each stage. The micro air launch vehicle is designed to launch a 20-kg payload into a 400-km circular polar orbit after release from mother plane at 12-km with initial velocity of 300 m/s. Realizing the important role of the fairing mass on vehicle performance, a geometric method is used to estimate the fairing mass for micro air launch vehicle. The estimation gives fairing mass between 8 kg and 12 kg. Using PSO, the optimal micro air launch vehicle initial mass achieved is 1287 kg and the maximum payload mass is 20.7 kg for 12-kg fairing mass. When the 8-kg fairing mass is assumed, the initial mass achieved is reduced to 1267 kg and the maximum payload mass is reduced to 20.64 kg. DE is used to solve the same optimization problem with 8-kg fairing mass for comparison with PSO. The initial mass obtained from DE using fairing mass of 8 kg is 1267 kg and payload mass is 20.61 kg. Therefore, both evolutionary techniques give very close optimal results for the same objective functions and same design variables. | ||
| 596 | |a 1 | ||
| 650 | 0 | |a Launch vehicles (Astronautics) |x Design and construction | |
| 655 | 7 | |a Theses, IIUM local | |
| 690 | |a Dissertations, Academic |x Department of Mechanical Engineering |z IIUM | ||
| 710 | 2 | |a International Islamic University Malaysia. |b Department of Mechanical Engineering | |
| 856 | 4 | |u http://studentrepo.iium.edu.my/handle/123456789/5052 |z Click here to view 1st 24 pages of the thesis. Members can view fulltext at the specified PCs in the library. | |
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