Design of a New Multi-Variant Solar Panel Deployment System (MVSPDS)

Design of a New Multi-Variant Solar Panel Deployment System (MVSPDS)

Souradeep Hazra

Final Year of Aerospace Engineering

Department of Aerospace Engineering, Chandigarh University Punjab, India

Abstract Multi-Variant Solar Panel Deployment System (MVSPDS) is defined as the satellite deployment System which can change its orientation according to the power supply required for the satellite or power supply source available to the satellite. To deploy the solar panels completely, it is necessary to design the deployment mechanism which has high precision and reliability. So this Method of deployment will not only provide the said benefit but also it will allow a wide range of application. Consequently, the analysis on the dynamic characteristic of the deployment mechanism must be done at an initial design stage. The design effectiveness and structural safety of the proposed solar panel module were validated by launch vibration and in- orbit environment tests at the qualification level.

In this paper, the complete design of a new Multi-Variant Solar Panel Deployment System in a Satellite is proposed, where I have inculcated various deployment methods and proposed a new method of satellite deployment. The complete design is done in the Autodesk 360 software (Educational license) where the design and animation method are extensively used to make this possible. It will be clearly depicted from the design that the structure is very compact during the process of orbit insertion. Yet when it come into working it definitely works according to the power need of the satellite. For an example when the satellite needs less amount of power it will transform itself in such a way that the amount of power generation will be less on the other hand when the satellite needs more power than the Solar Panel will change itself in such a way that it will have the maximum surface area as a result of which the power generation will be maximum. Moreover, the variety which provides is really unique as it also has the power to change its solar panel into different structure for an example it can change itself into a circle like structure it can change itself in the canister to form another structure and moreover it will work completely according to the need if programmed properly for space flight mission. And hence can give rise to a wide variety of different deployment methods.

KeywordsSatellite, solar panel orientation, Solar panel deployment, multi-Variant Solar Panel

1. INTRODUCTION

   Multi-Variant Solar Panel Deployment System (MVSPDS) is defined as the satellite deployment System which can change its orientation according to the power supply required for the satellite or power supply source available to the satellite. The sole purpose of this project is based on designing a of a new Multi-Variant Solar Panel Deployment System. Solar technology can be a political football on the ground tossed around and tackled oftenin space, it encounters little opposition. For space limited operations many components were successfully miniaturized.

   The stabilization types changed from simple spin stabilization and passive stabilization using permanent magnets to fully three-axis stabilized spacecraft. Some recent missions also validated propulsion systems for satellites. Despite this significant progress, two problems remain unsolved due to the small surface area of satellite. The commonly used body fixed panels produce insufficient energy and similarly provide too little surface area for adequate energy dissipation. This project will address the basic arrangement along with various variation in the arrangement of solar array. One of the possible solutions is the use of deployable structures to increase the available surface area in many possible ways and with different transformation mode. This paper deals exclusively with a new multi-Variant structural design of a single satellite although dual system is possible. There are many aspects of the design, construction, and operation of such a space solar power system that are outside the scope of this project (e.g., the design of lightweight photovoltaic cells, large-scale phased arrays across independent spacecraft, integrated circuits for microwave signal synthesis, and formation flying).

   Image of Basic Deployable Structure

   However, the concern herein is purely structural design and its deployment, and, as such, only relevant metrics will be

   considered which include the operation of the solar array. This paper will present a preliminary structural design of a spacecraft that has less area with small packaged volume, and is sufficiently stiff.

   The developed design is based on a generic satellite mission and an orbit with an altitude of 400 kilometers. The satellite is assumed to be three-axis stabilized. The panels are usually in stowed position to avoid damage during the launch phase and are deployed later on when the satellite reaches its desired orbit.

   In general, mechanisms on board of satellites are limited to small dimensions and are required to be extremely light weight by the Satellite launch system. Working in an environment with no possibility of later corrections or modifications requires a high functional reliability and repeatability. The two major impacts on the satellite are the compatibility of the thermal heat

   expansion ratios and the large occurring temperature changes. These result in thermally induced stresses which might cause cracks in components or delamination of glued components like solar cells.

2. LITERATURE REVIEW

   This section covers the literature review of solar arrays used and the method of deployment for space applications. The first section gives a brief overview of the principal solution options. Afterward currently used techniques are evaluated to determine the most valuable direction of development.

   For a systematic review of existing solar array systems, the complete system of a deployable solar array was split into several subsystems:

   • Actuation mechanism for deployment

   • Guiding mechanism for deployment

   • Damping mechanism for deployment

   • Initial release mechanism

   • Actuation mechanism for articulating

   • Control mechanism for articulating

   • Solar cell technologies

   • Solar panel materials

   Solution options were generated by analysis of current solar panel systems and research in scientific journals, patent databases and books. The options originate from satellites ranging from small satellites to very large satellites, like the ISS or the Hubble Space Telescope. But is concept of design is completely new.

3. SOLAR CELL

   Although many types of solar cell are to be used but it is recommended to use a triple junction solar cells because it has high efficiency and good radiation resistance. While the earlier discussed solar cells use just one n-p junction, triple junction solar cells combine three material pairs to cover a larger part of the spectrum of the electromagnetic radiation. A possible combination is the three junctions are Gallium-Indium- Phosphorus, Gallium-Indium- Arsenide and Germanium. Current solar cells provide a conversation efficiency of 28.3 percent and experience a degradation of 15 percent by 1 MeV in 33 years.

   Triple-Junction solar cells offer the highest conversation rate and a good radiation resistance. This enables this project of reaching the optimum power output from the solar cells. Therefore triple- Junction solar cells will be used within this paper. Each panel contains 6 solar cell of 12V. Hexagonal shape is given to fit the six panel all together apt for design.

4. DESIGN OF SINGLE PANEL

   Initial side of the panel is kept constant so as to keep the solar cell in a less complicated manner. Although the hexagonal panel is done but different shape is also possible.

   Detail Design Front View

5. DESIGN OF DIFFERENT VARIANTS Different variant accounts for different range of power supply

   to the satellite. Possible variant of the MVSPDS is given below.

6. STAGES OF DEPLOYMENT

Different stages of deployment are given below in a tabular form.

ACKNOWLEDGMENT

I take upon this opportunity to acknowledge such people whose participation and help has played an integral part in completion

this research project with flying colors. I am deeply indebted to my Autodesk 360 team for helping me with this project with proper guidance. I would like to thank our honorable Head of the Department (Aerospace) for furnishing us with proper reasoning and knowledge, helping and supporting me at every moment. I owe my sincere gratitude towards other faculty member who cordially gave ideas regarding the paper. My heartfelt thanks to the lab assistant who has provided us the required information in crucial time. I also express my deepest gratitude to the institution for providing us the opportunities for such interesting paper and file a patent on it. Finally, I would like to wind up by paying my heartfelt thanks to all faculty member who has helped us in doing this paper and to the institution for providing us the opportunity for such enthusiastic paper. Last but not least, I want to express my special regards and thankfulness to my family for their great and continuous support and motivation without my research paper without whom it wouldn't have been possible.

CONCLUSION

From the above design it can be concluded that the design is a unique design which is not only having multi variant structure but also having wide range and dynamic performances.

It is clearly depicted from the design that the structure is very compact it during the process of orbit insertion. Yet when it come into working it definitely works according to the need of the satellite. For an example when the satellite needs less amount of light it will transform itself in such a way that the amount of power generation will be less on the other hand when the satellite needs more power than the Solar Panel will change itself in such a way that it will have the maximum surface area as a result of which the power generation will be maximum. Moreover, the variety which provides is really e unique as it also has the power to change its solar panel into different structure for an example it can change itself into a circular structure it can change itself into a canister form and moreover it will work completely according to the need if programmed properly. And hence can give rise to a wide variety of different deployment methods for solar panels in space.

REFERENCES

1. S. Lee, A. Hutputanasin, A. Toorian, W. Lan, and R. Munakata, CubeSat Design Specification, 01-Aug-2009.

2. P. Höhn, Entwicklung und Bau eines ausklappbaren Sonnensegels für den Cube- Satelliten Uwe 2, Diplomarbeit, University of Applied Sciences Coburg, 2008.

3. M. Kashiwa u. a., Tokyo Tech CubeSat: CUTE-I, presented at the AIAA 21st International Communications Satellite Systems Conference and Exhibit, 2003.

4. V. G. Baghdasarian, Hybrid solar panel array, U.S. Patent EP0754625A1.

5. B. Duperray, A. Donzier, and J. Sicre, Automotive, self-locking and damping articulated joint and articulation equipped with same, U.S. Patent US 2001/0037538 A108-Nov-2001.

6. W. W. Vyvyan, Self-Actuating, Self-Locking Hinge, U.S. Patent US3,386,12804-Juni- 1968.

7. A. Carpine, C. Martin, C. Rinn, and C. Verne, Architecture and Deployment of a high power Solar Array, presented at the European Space Mechanism and Tribology Symposium, 2009.

8. S. C. Meyers and M. D. Greenbelt, AIAA/USU Conference on Small Satellites, 10 th, Utah State Univ, Logan, UT, 1996.

9. D. C. Lagoudas u. a., Introduction to Shape Memory Alloys, in Shape Memory Alloys – Modeling and Engineering Applications, Springer Science+Business Media, 2008, S. 1-51.

10. T. Bieler, in Handbuch der Raumfahrttechnik, 3. Aufl., Munich: Hanser, 2008, S. 238.

11. H. Heidt, J. Puig-Suari, A. S. Moore, S. Nakasuka, and R. J. Twiggs, CubeSat: A new generation of picosatellite for education and industry low-cost space experimentation, S. 12, 2001.

12. L. Herbeck, C. Sickinger, and A. Herrmann, Ultraleichte entfaltbare Maststrukturen aus CFK, DGLR Jahrbuch, S. 739748, 1999.

13. Y. Miyazaki, H. Isobe, T. Kodama, M. Uchiki, and S. Hinuma, AIAA/USU Annual Conference on Small Satellites, 15 th, Utah State University, Logan, 2001.

14. G. Tibert, Deployable tensegrity structures for space applications, Royal Instituteof Technology, 2002.

15. A. Binetti, ABLE Engineering Company, Inc, ABLE Engineering Company, Inc, 16- Feb-2005.

16. B. Schoneberger, Coilable PDS, 2003.

17. STS117 Array Deployment. 2008.

18. J. R. Wertz and W. J. Larson, in Space Mission Analysis and Design, Third Edition, Ninth Printing., Hawthorne: Microcosm Press, 2007, S. 417.

19. N. J. Leon, ST8 Technologies, ST8 Technologies, 23-Apr-2010. [Online].Available: http://nmp.nasa.gov/st8/tech/technologies.html. [Accessed: 23-Apr-2010].

20. A. R. Lavoie, Post-Flight Engineering Report Performance TSS-1R (STS-75). NASA, 1996.

21. ACE Stoßdämpfer GmbH, Rotationsbremsen, 2007.

22. C. V. Enriquez u. a., A preliminary design for a satellite power system. NASA,1991, S. 97.

23. S. H. Smith, D. Dowen, E. Fossness, and A. Peffer, AIAA/USU Annual Conference on Small Satellites, 13 th, Utah State University, Logan, 1999.

24. B. Huettl and C. E. Willey, AIAA/USU Annual Conference on Small Satellites, 14th, Utah State University, Logan, 2000.

25. on Small Satellites, 14 th, Utah State University, Logan,2000.

26. SQL-RV-1.8 Reduced Voltage SQUIGGLE RV linear drive system. [Online]. Available:

    http://www.newscaletech.com/doc_downloads/SQL-RV-1- 8_datasheet.pdf.

27. P. Fortescue, J. Stark, and G. Swinerd, in Spacecraft Systems Engineering, 3. Aufl., J Wiley & Sons, 2009, S. 328-336.

28. W. Rubel, Satellite Design – Power System, Dez-2008.

29. J. R. Wertz and W. J. Larson, in Space Mission Analysis and Design, ThirdEdition, Ninth Printing., Hawthorne: Microcosm Press, 2007, S. 414.

30. J. Brown, Test Pod Users Guide, 11-Juni-2006.

31. P. Forescue, J. Stark, and G. Swinerd, in Spacecraft Systems Engineering,Third Edition., J Wiley & Sons, 2009, S. 363.