DESING AND ANALYSIS OF SATELLITE SOLAR PANEL GRADUATION PROJECT. Büşra ÇATALBAŞ. Department of Astronautical Engineering

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1 ISTANBUL TECHNICAL UNIVERSITY FACULTY OF AERONAUTICS AND ASTRONAUTICS DESING AND ANALYSIS OF SATELLITE SOLAR PANEL GRADUATION PROJECT Büşra ÇATALBAŞ Department of Astronautical Engineering Thesis Advisor: Prof. Dr. Zahit MECİTOĞLU JANUARY 2019 i

2 ISTANBUL TECHNICAL UNIVERSITY FACULTY OF AERONAUTICS AND ASTRONAUTICS DESING AND ANALYSIS OF SATELLITE SOLAR PANEL GRADUATION PROJECT Büşra ÇATALBAŞ ( ) Department of Astronautical Engineering Thesis Advisor: Prof. Dr. Zahit MECİTOĞLU JANUARY 2019 ii

3 Büşra Çatalbaş, student of ITU Faculty of Aeronautics and Astronauticsstudent , successfully defended the graduation entitled Design and Anaylsis of Satellite Solar Panel which she prepared after fulfilling the requirements specified in the associated legislations, before the jury whose signatures are below. Thesis Advisor: Prof. Dr. Zahit MECİTOĞLU... İstanbul Technical University Jury Members: Prof. Dr. Metin O. KAYA... İstanbul Technical University Prof. Dr. A. Rüstem ASLAN... İstanbul Technical University Date of Submission: 02 January 2019 Date of Defense: 16 January 2019 iii

4 To my family, iv

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6 TABLE OF CONTENTS Page FOREWORD... Error! Bookmark not defined.v TABLE OF CONTENTS... vi ABBREVIATIONS... vierror! Bookmark not defined. LIST OF TABLES... Error! Bookmark not defined. LIST OF FIGURES... Error! Bookmark not defined. SUMMARY... Error! Bookmark not defined. 1. INTRODUCTION Purpose of Thesis Literature Review... Error! Bookmark not defined. 1.3 Design Specification of the Satellite... Error! Bookmark not defined. 2. CONCEPTUAL DESIGN(spacecraft structure)... Error! Bookmark not defined. 2.1 Purpose... Error! Bookmark not defined. 2.2 Mission Requirements... Error! Bookmark not defined. 2.3 Design Requirements... Error! Bookmark not defined. 2.4 Power Requirements... Error! Bookmark not defined. 3. DETAILED DESIGN(solar panel design)... Error! Bookmark not defined. 3.1 Purpose Design and Sizing Trades... Error! Bookmark not defined Autoregressive models... Error! Bookmark not defined Process based model: SWAT... Error! Bookmark not defined Multi variable analysis... Error! Bookmark not defined. 3.2 Deployment Mechanisms and Structures Error! Bookmark not defined. 3.3 Application Data... Error! Bookmark not defined. 4. (IF NECESSARY) CHAPTER 4... Error! Bookmark not defined. 4.1 Practical Application of This Study... Error! Bookmark not defined. 4.2 Second Level Title: First Letters Capital Error! Bookmark not defined Third level title: Only first letter capital... Error! Bookmark not defined Fourth level title: Only first letter capital Error! Bookmark not defined. Fifth level title: No numbering after fourth level titles... Error! Bookmark not defined. 5. (IF NECESSARY) CHAPTER 5... Error! Bookmark not defined. 5.1 Practical Application of This Study... Error! Bookmark not defined. 5.2 Second Level Title: First Letters Capital Error! Bookmark not defined Third level title: Only first letter capital... Error! Bookmark not defined Fourth level title: Only first letter capital Error! Bookmark not defined. Fifth level title: Only first letter capital... Error! Bookmark not defined. 6. CONCLUSIONS AND RECOMMENDATIONS... Error! Bookmark not defined. 6.1 Practical Application of This Study... Error! Bookmark not defined. 6.2 Second Level Title: First Letters Capital Error! Bookmark not defined. vi

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9 1. INTRODUCTION Satellites are used for several services and applications. They also need energy to function like other machines. On the ground, there are many ways to produce electrical power from several energy supplies. However, in the space, the problem is where to get that power from. Hence, the power generation source is one of the most sensitive issue in the satellite science. Photovoltaic solar arrays have been commonly used to electrical energy for satellites since Vanguard 1 was launched in It was the first satellite to use solar energy, used to obtain geodetic measurements through orbit analysis. Vanguard 1 is also the oldest satellite still orbiting the Earth. Single crystal silicon solar cells mounted on the body of the satellite produced a total of approximately 1 Watt at 28 C with an efficiency of 10%. [1] Power needs of spacecraft increased with the development of vehicles and space missions. Therefore, design and efficiency of solar panels have been significantly improved. Today, the largest space solar power system with 120 kw is used by The International Space Station. [7] Despite increasing efficiency of solar cells, higher power demands may lead to larger solar panel requirements. Nevertheless, there are also various design constraints with low mass and low volume to match the dimensions of launch vehicle. Additionally, reducing the size of the solar panel increases the spacecraft's maneuverability. Consequently, the high power requirement of aerospace applications, has led to innovative, high-performance materials and structures as well as the rising efficiency of solar cells. The sandwich structures have good stiffness properties and energy absorbing capacity without a significant increase in weight. Because of these crucial advantages and simplicity of production, sandwich constructions have been widely used in aerospace structures. There are three main layers of sandwich structures: core, facing sheet and adhesive bond. Each layer carries different loads and moments. Therefore, the use of this robust sandwich structure provide high power-toweight ratios for solar panels. During launch of satellite, combination of accelerations of launch vehicle and random vibrations cause axial and lateral loads on the satellite components. 9

10 1.1 Purpose of Thesis 10

11 1.2 Design Specification of Satellite Water is the key to sustain organic life on the Earth, Moon and everywhere in the known universe. The poles of the moon are noteworthy places to find water based on previous research. Objective of this mission is to determine the location and quantities of water deposits. Mission Concept Remote sensing from the orbit may be used to determine the ratio of water to regolith with the position data. However, in order to satisfy accessibility of the water inside the craters, an additional surface mission is necessary that remote sensing cannot provide a total evidence for accessibility. Consequently, remote sensing is not enough in order to provide heritage for accessibility to the water. Hence, there are a lander, a rover and an orbiter to determine the water quantity analysis on this mission. Lander and rover are designed to determine ratio of water to regolith with surface operations. Therefore, this mission uses a main lander spacecraft, a rover, which is a payload to lander and an orbiter, which is a support for the lander. Scheme of the whole operations can be seen on Figure 1.1. Figure 1.1: Concept of Operation [2] Date of launch from Earth is chosen as May A two-stage launch vehicle, Atlas V, will be used to place Lunar Aqua onto direct lunar transfer orbit. Time of flight will be 5 days from 11

Centaur s re-ignition, second burn, to lunar transfer orbit insertion. Orbiter and lander will travel in space in integrated to each other as shown in the Figure 2. Figure 1.

12 Centaur s re-ignition, second burn, to lunar transfer orbit insertion. Orbiter and lander will travel in space in integrated to each other as shown in the Figure 2. Figure 1.2: Isometric View of the Spacecraft [2] Orbiter with the lander on top will take the spacecraft to desired Moon s orbit, which is 50 km altitude by use of two different maneuvers, which will performed by orbiter s propulsion system. Orbiter and lander after decreasing to 50 km altitude from lunar surface will separate. Orbiter will stay in circular orbit for two years. Every month it will use station-keeping maneuver with value of 11 m/s and this maneuver will consume 33.9 kg propellant totally. [2] Spacecraft Concept Orbiter s dimension is 1250X800X800 mm and structure includes solar panel, thruster, and electronic devices. Its total dry is kg and total wet is 706 kg. Orbiter will stay in circular orbit for two years. Every month it will use station-keeping maneuver with value of 11 m/s and this maneuver will consume 33.9 kg propellant totally. [2] Projects main goal is to determine lunar ice ratio to the regolith. This will be achieved by surface experiments with additional remote sensing. This mission aims to determine water mass ratio to the regolith with two ways: Remote sensing from the orbit and surface experiment with laser induced breakdown spectroscopy. Therefore, there are two scientific payloads in orbite: LEND and mini-sar. 12

13 LEND (Lunar Exploration Neutron Detector) is a neutron spectrometer that is integrated on the orbiter. It uses the neutron albedo of moon to map the hydrogen rich regions on moon. If the hydrogen is associated with water, a detection limit of 100-ppm hydrogen corresponds to ~ 0.1% weight water ice in the regolith. From a 50km orbit, the instrument has a 1-2m spatial resolution inside the surface and has a spatial resolution of 5 km on the lunar surface. [3] In addition, Mini-SAR (Miniature Synthetic Aperture Radar) is the second instrument on the orbiter. The radar uses the polarization properties of reflected radio waves to characterize surface properties. [4] Orbiter is the main tool for communicating with the Earth. It will not only communicate with DSN, but also communicate with lander. As it was stated in the lander section, communication data between orbiter and lander will be carried in X-band. High gain Helical X-band Antenna will also be used in orbiter to communicate to lander with the supporting of XTx-400 and Mitsubishi X-band receiver. DSN will be used as main communication architecture. Image data rover gathered will be received from lander by cabled communication. Ka-band will be used for sending image and communication data to DSN. In all other operations, X-band will be used. [2] Orbiter need precise attitude control and determination systems due to detailed mission requirements. ADS system of the orbiter includes two star trackers and one inertial measurement unit. Four reaction wheel and four thruster combination is used for the ADC system. ADC system has attitude sensing requirements and should counterbalance the disturbance torques for the 50 km Moon orbit. Power Requirement of Satellite Solar panels are the primary source of power of our orbiter. Orbiter will be in daylight region for the most of time. In daylight operations mode, power is distributed to all systems from the solar panels. In addition, charging of the batteries takes place in this mode. In eclipse operations mode, power to subsystems is supplied from the batteries. The batteries are capable of supplying the maximum power draw of the subsystems for the whole eclipse duration. 13

Table 1.1: Subsystem Power Allocation for the Orbiter Spacecraft [2] 1.2.3.1 Maximum Eclipse Time As shown in Table 1.1, requirement power of subsystems is 386 W per one hour.

14 Table 1.1: Subsystem Power Allocation for the Orbiter Spacecraft [2] Maximum Eclipse Time As shown in Table 1.1, requirement power of subsystems is 386 W per one hour. For design calculation of solar array size and battery systems, the maximum eclipse time must be calculated from below formulas; Figure 1.3: Maximum Eclipse Period, Circular orbit [5] Figure 1.3 shows the maximum eclipse period condition, which occurs when the sun is in the orbit plane. As shown in Figure 1.3, is the shadow region half-angle, in terms of degree. R is the radius of the central body, in other words it is Moon radius. h is the orbit altitude. From Figure 1.3: 14

15 = sin 1 ( R R + h ) (1.1) h = 50 km R = 1737 km With calculation of Eq 1.1, the shadow region was found as: = deg For circular orbit, Τ n is the maximum eclipse perio and Ρ is the orbit period. Τ n = 2 Ρ 360 (1.2) Ρ = 113 min With calculation of Eq 1.2, the maximum eclipse perios was found as: Τ n = min In 48-minute-long maximum eclipse region, solar panels should both generate enough power to charge the batteries for the next eclipse region and to meet the power requirement of the subsystems Bateries The orbiter will experience eclipse periods as much as 48 minutes per orbit. Therefore, the battery system on the orbiter should be able to supply the total power requirement for 48 minutes. Eagle-Picher Technologies 43Ah Space grade battery cells will be used on the orbiter to provide the necessary amount of energy in the eclipse regions. To have a low DOD value, eight of these batteries will be used in parallel. They have a long life over 40,000, LEO DOD over 10 years of operation. [6] Taking into account filling of the batteries, 600 W energy must be obtained from the solar arrays during daylight mode. On eclipse mode, power management unit minimizes the use of energy, because the batteries must meet the energy needs. In the Figure 1.4 can be seen power management system. 15

16 Figure 1.4: Sapcecraft Power Diagram [2] 16

17 2. DESIGN 2.1 Conceptual Design The solar array technology started to improve with Vanguard's 10% efficient single-crystalline silicon solar cells. This oldest solar panels had limited space because they were made up of solar cells mounted on the body of a spacecraft. Thereafter, both solar cell technology and sequence structure have made significant development in power capacity, mass and cost efficiency. Today there are various solar cell and array types with different materials and sizes with 30% efficiency. Therefore, trade-off study is required for the selection of the optimum solar array and cell technology for the design of spacecraft solar panel. Selection of Solar Cells There are many criteria to choose the most suitable solar cell such as power per unit mass ( W kg ), power per unit mass ( W m 2 ) (it is also known as efficiency (percentage)), cost per unit power ( $ W ), radiation and atomic oxygen resistance, and reliability. Cell efficiencies can be reported at different time interval. For example beginning of life, (BOL) refers to the efficiency of the cell at the time it was first used, and End of life (EOL) shows efficiency at end of the mission. The efficiency of solar cells is based on standard assumptions like air mass zero (AMO) environment that is outside of the Earth s atmosphere at 1 AU from the sun where the temperature is 28 C and the solar intensity is 1367 W m 2. [8] A smaller solar panel area is easier to integrate into the spacecraft, easier to place and steer in space, and with less mass. Therefore, the highest efficiency value must be priority preference. Primarily, silicon solar cells are used since the beginning of solar cells technology; however, silicon solar cells are very sensitive to space radiation and experience significant loss of 17

18 productivity over time. By contrast, multi-junction solar cells advantage is the multiple layers of light-absorbing materials, GaInP2 and GaInAs, and a germanium substrate with an active 18

1, the widest bandgap material GalnP2 with wavelength region from 300 nm to 1800 nm, is the top junction and each another

19 junction, each of which efficiently convert specific wavelength regions of the solar spectrum into energy. Figure 2.1: Illustration of a Triple-Junction Solar Cell and Solar Irradiance Spectrum [7] As shown in the Figure 2.1, the widest bandgap material GalnP2 with wavelength region from 300 nm to 1800 nm, is the top junction and each another subcell has a lower bandgap material. Thus, photons of many different wavelengths are transformed into energy. [7] Figure 2.2: Efficiency of Solar Cells[9] 19

20 In Figure 2.2, there are 26 different categories with distinctive color. Cell efficiencies are shown with different semiconductor families such as multijunction, single-junction gallium arsenide and crystalline silicon cells. As shown in Figure 2.2, multi-junction GaAs-based cell technology are more efficient than silicon solar cells. Solar cell efficiency is the most significant parameter to optimize in order to achieve minimum mass and volume of the solar cell power system. Therefore, triple junction GaAs based solar cells was preferred for using on this mission because of high efficiency gains and high environment tolerance. As a consequence, selected multijunction cells are the XTJ-Prime Solar Cells are produced by Spectrolab, its efficiency is 30.7% beginning of life in AM0 conditions. [10] Selection of Solar Arrays There are some criteria to make a comparison when choosing solar panels: one of the most important is the Technology Readiness Levels (TRL). It is a type of measurement system used to assess the matureness of a particular technology. There are nine technology readiness levels. TRL 1 is the lowest value; TRL 9 is the highest value. If a technology has been "flight proven" during a successful mission, it can be called TRL 9. [11] As shown in Table 2.1, the reliability of the arrays is quite high with TRL 9. Table 2.1: Efficiency of Solar Arrays [7] Flexible arrays are not considered because they are mainly used to larger power needs. As shown Table2.2, deployable rigid solar array is the best option for orbiter power supply. Main reason is the high efficiency with 330 W m 2 and deployable arrays are more suitable for the spacecraft. Selection of Deployment Mechanisms 20

21 Configuration of Solar Array 2.2 Detailed Design Selection of Core Material Selection of Face Material Selection of Adhesieve and Coating Material Material Orientations 3. ANALYSIS 3.1 Purpose 21

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Lunar Flashlight Project

Lunar Flashlight Project ABSTRACT Recent observations of the Moon with the Moon Mineralogy Mapper (M3), Lunar Crater Observation and Sensing Satellite (LCROSS), the Lunar Reconnaissance Orbiter (LRO) and other evidence suggest