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Energy Engineer - Renewable Energy, 180 credits
Dark Ages Interferometer (DALI)
Deployment Rover: Energy System
Gustav Andersson, Emil Ericsson
Energy Technology
Energy Technology
Dark Ages Lunar Interferometer
(DALI) Deployment Rover:
Energy System
Emil Ericsson
Gustav Andersson
2014-06-02
Supervisors:
Professor Jonny Hylander
Professor Lars Bååth
Doctor Ashitey Trebi-Ollennu
Examiner :
Abstract
The cosmic “Dark Ages” is the cosmic era between the epochs of recom-bination of cosmic microwave background and the formation of the first stars. The only signal from this epoch is from neutral hydrogen, which could represent one of the richest data sets in cosmology. In order to extract this data, NASA/JPL has proposed a rover mission to the far side of the moon to deploy several radio arrays. Here the arrays would gather data undisturbed by human interference. This thesis examines the possibility of using photovoltaic and electric batteries as an energy solution for a rover on the moon. The requirement for such a system to survive on the moon is discussed in a literature study. A proof of con-cept simulation using a Simulink model has also been done. The thesis concludes that a rover can deploy the radio array using solar energy. It would be able to hibernate through the night using radioisotope heating. It would need to wait for its batteries to charge before each night.
Sammanfattning
I kosmologi kallas epoken mellan “rekombinationen till väte” och bildandet av de första stjärnorna för “den mörka tidsåldern”. Från denna tid finns endast spår i form av strålning från neutralt väte. Denna strålning kan enligt astronomer vara en viktigare källa till data om universums uppkomst än den kosmiska mikrovågsstrålningen. Därför arbetar NASA/JPL med att hitta metoder att observera denna rika källa till data. Den mest använda metoden är att använda lågfrekventa radioteleskop för att observera strålning med våglängder mellan 3-30 m och frekvenser mellan 10-100 Mhz. Ett stort problem med så kort strålning är den lätt störs ut av mänsklig påverkan och andra radiokällor, till exempel solen. Ett sätt att undvika antropogen störning är att bygga ett radioteleskop på månens baksida. Eftersom månen är i en låst bana runt jorden vänder den alltid samma sida bort från planeten. Därför är platsen alltid i radioskugga från jorden och störs inte av mänsklighetens radiotrafik. JPL har lång erfarenhet av användandet av robotfarkoster för att utforska himlakroppar. År 2030 vill de sända en så kallad rover för att upprätta en grupp radioteleskop på månen med syftet att införskaffa data om “den mörka tidsåldern.” Högskolan i Halmstad erbjuder sedan 2013 studenter möjligheten att skriva sitt examensarbete i samarbete med NASA/JPL om konstruktionen av denna rover. Detta arbete har ämnat finna en lösning på rovens energiförsörjning genom att använda solceller och batterier. Slutsatsen har varit att det är möjligt att driva en rover på månen med solceller samt batterier. Ett krav är att rovern värms med radioisotoper under natten för att minska energianvändningen.
Preface
This work is the result of a five month project preformed in collaboration with NASA/JPL and Halmstad University. The work is the bachelor degree thesis of the renewable energy engineering program at Halmstad University and represents 15 hp or 15 ECTS. As stu-dents this project has been an enormous opportunity for us to test our knowledge and our skills at solving problems. It has been an interesting journey, but not one we have travelled alone. We would therefore like to thank our supervisors, Professors Jonny Hy-lander and Lars Bååth for their valuable help and support. We would also like to thank Dr. Ashitey Trebi-Ollennu from JPL for his feedback. This project would not have been completed without their help.
Contents
Contents iv
List of Figures vi
List of Tables vii
Nomenclature viii
1 Introduction 1
1.1 Background . . . 1
1.2 Company Presentation . . . 2
1.3 Aims and Objectives . . . 3
1.4 Problem Statement . . . 3
1.5 Assumptions and Delimitation . . . 3
2 Methods and Materials 4 2.1 Work Schedule . . . 4
2.2 Sources and Materials . . . 4
3 Literature Study 5 3.1 Information Regarding the Moon . . . 5
3.1.1 Cosmic Radiation and Solar Irradiance . . . 5
3.1.2 Lunar Topography and Environment . . . 8
3.2 Previous Rovers . . . 11
3.2.1 Lunokhod 1 . . . 11
3.2.2 Lunokhod 2 . . . 12
3.2.3 Lunar Roving Vehicle . . . 13
3.2.4 Sojourner . . . 14
3.2.5 Spirit and Opportunity . . . 15
3.2.6 Curiosity . . . 16
3.3 Halmstad University Rover Concept . . . 18
3.3.1 Warm Electronics Box . . . 20
3.3.2 Photovoltaic Power Generation . . . 21
3.3.3 Energy Storage . . . 24
4 Results 28 4.1 Rover Power Usage . . . 28
4.2 Power Generation . . . 29 4.3 Energy Storage . . . 31 5 Discussion 35 5.1 Power Usage . . . 35 5.2 Power Generation . . . 35 5.3 Energy Storage . . . 36 6 Conclusions 39 Bibliography 40 Appendices 45 A Rover Power Intensive Components . . . 46
B Simulation Model . . . 47
C Simulation Model: Subsystems . . . 48
D Solar Power Simulations . . . 49
List of Figures
1.1 Concept picture of deployed array. . . 2
3.1 Amount of solar power available at the moon’s surface per square meter. 8 3.2 Location of lunar marinas. . . 9
3.3 Sunlight hitting the center of the crater. . . 10
3.4 Lunokhod 1 . . . 12 3.5 Lunokhod 2 . . . 13 3.6 Apollo 15. . . 14 3.7 Sojourner on Mars. . . 15 3.8 Opportunity . . . 16 3.9 Curiosity . . . 17 3.10 Yutu . . . 18
3.11 Initial process schematic. . . 19
3.12 Concept picture of the Halmstad University Rover. . . 20
3.13 Solar cells laboratory efficiency. . . 21
3.14 Gallium Arsenide band gap dependency on temperature. . . 23
3.15 Image of a simple electrochemical cell. . . 25
4.1 Rover simulation at varying loads. . . 30
4.2 Solar power simulations. The x-axises show the time in seconds and the y-axises show the SOC percentage, voltage and irradiance. Higher resolutions are available in appendix D. . . 31
4.3 Li-ion battery performance at different loads. . . 32
4.4 Hibernation simulations. The x-axises show the time in seconds, and the y-axises show SOC percentage, voltage and current. Higher resolutions are available in appendix E. . . 33
4.5 Battery charging with a load of 5 W. . . 34
List of Tables
1.1 Student groups and components area. . . 2
3.1 Penetration and energy levels of particles. . . 6
3.2 Measured temperatures and thermal conductivity of lunar surface and sub-surface. . . 8
3.3 Sunlight percentage. . . 11
3.4 Comparison of NiCd, NiH2, Silver-zinc, and Li-ion performance . . . 27
4.1 Insulation depending on thickness. . . 28
4.2 Time estimation. . . 29
4.3 Simulation results. . . 29
4.4 Materials for solar cells. . . 30
4.5 Volumetric and mass properties of batteries. . . 31
Nomenclature
Symbolsα Energy per Kelvin [eV/K]
β Material dependent constant
ω Angular frequency [rad/s]
Φ Thermal flux [W]
Θ Total amount of unshaded sky [◦]
θ Solar height above mountainous horizon [◦]
4T Temperature dfference [K]
4x Insulation width [mm]
ϕ Latitude [◦]
E0 Bandgap at absolute zero [eV]
EG Bandgap dependent on temperature [eV]
I Intensity of the sun (Irradiance) [W/m2]
k Thermal conductivity [W/m·K]
S Surface area [m2]
T Temperature [K]
t Time [h, min, s]
tsun Time without shading each day [h, min, s]
VOC Open-circuit voltage [V]
Abbreviations
Caltech California Institute of Technology
CdS Cadmium Sulphide
CIGS Copper Indium/Gallium (di)Selenide
CMB Cosmic Microwave Background CNSA Chinese National Space Agency DALI Dark Ages Lunar Interferometer
DC Direct Current
DFMR Design for Minimum Risk
DOD Depth of Discharge [%]
GaAs Galium Arsenide Solar Cell
GCR Galactic Cosmic Rays
HGR Heat Generating Resistor
HI Neutral Hydrogen
ISS International Space Station
JPL Jet Propulsion Laboratory
LOFAR Low Frequency Array
LRV Lunar Roving Vehicle
LWA Long Wavelength Array
MER Mars Exploration Rovers
MPF Mars Pathfinder
MSL Mars Science Laboratory
MWA Murchison Widefield Array
NASA National Aeronautics and Space Administration
NiCd Nickel-cadmium
NiH2 Nickel-hydrogen
NREL National Renewable Energy Laboratory
PV Photovoltaics
RHU Radioisotope Heater Unit
RTG Radioisotope Thermoelectric Generator
SOC State of Charge [%]
WEB Warm Electronic Box
Glossary
Band gap The energy level required to excite an atom’s electrons. Rover A robotic land vehicle used for exploration of celestial bodies.
1 Introduction
1.1 Background
In astronomy, the “Dark Ages” is the cosmic era between the epochs of recombination of cosmic microwave background (CMB) and the formation of the first stars. During this darkened time the only detectable signal is likely to be from neutral hydrogen (HI), that appears in the absorption against the CMB. The HI absorption occurs in areas with large concentration of dark matter. These places were the cradles from which the first stars eventually formed. NASA/JPL considers the HI absorption to potentially represent:
“The richest of all data sets in cosmology, not only is the underlying physics relatively simple so that the HI absorption can be used to constrain fundamen-tal cosmological parameters in a manner similar to that of CMB observations, but the spectral nature of the signal allows the evolution of the Universe as a function of redshift (z) to be followed” [1]
NASA/JPL hopes that further studies into the field will increase understanding of the early Universe, and that any deviation from the expected data will be a “Clean signature of fundamentally new physics”.
A suggested site for radio telescope arrays to observe the HI absorption is on the far side of the moon, since it is shielded from human generated interference during the lunar day and from solar emissions during the night. The array will be made of current ground-based telescopes such as Long Wavelength Array (LWA), Murchison Widefield Array (MWA) and Low Frequency Array (LOFAR). As such, it will observe wavelengths between 3-30 meters, and frequencies between 10-100 MHz. Each array will consist of a base unit with six dipole mats acting as antennas connected to it. These arrays will be deployed by rovers over an area of approximately 50 kilometers in diameter to obtain the required angular resolution.
The Dark Ages Lunar Interferometer (DALI) project is still in its infancy. NASA/JPL estimates the mission will commence in 2030. Since the development time of anything space related is considerable, it has been deemed important to start research into key components now. One such component is the rover which will be used for transporting and deploying the arrays on the moon. Halmstad University is involved in an annual project aimed at creating ideas for the rover design. The project is carried out by students from different courses designing different parts of the rover. The first project was completed in 2013 by three groups of mechanical engineering students and created solutions for the rover’s chassis, deployment system and mobility system. This year the project is conducted by two groups of students, one in mechanical engineering and one in energy engineering. The energy engineering students will in this report study the rover energy system, how it can generate energy and what it can be used for.
Figure 1.1 – Concept picture of deployed array.
Table 1.1 – Student groups and components area.
Student group Year Components
Johan Winberg & Tomislav Stanimirovic 2013 Chassis
Tobias Johannesson & Karl Hansson 2013 Deployment mechanism
Erik Andersson & Per-Johan Bengtsson 2013 Mobility system
Björn Bernfort & Haris Pasalic 2014 Suspension
Emil Ericsson & Gustav Andersson 2014 Energy system
1.2 Company Presentation
Jet Propulsion Laboratory (JPL) primarily designs and constructs robotic planetary spacecrafts. JPL is a part of California Institute of Technology (Caltech) and is funded through NASA.
The origin of JPL can be traced back to 1930 and Caltech Professor Theodore Von Kár-mán and a group of his graduate students. The first rocket launch took place in October 1936 in the deserts of California. The group took up work making rocket designs for the army until they were transferred in late 1958 to the newly formed National Aeronautics and Space Administration (NASA). Earlier that year USA had sent the satellite Explorer 1 into space. Explorer was built by JPL and marked a shift from rockets to payload. Since then, JPL has built many important space explorers, including the Mariner 2, both Voyagers and the Viking probes. In more recent years, JPL has constructed NASA’s rovers for Mars, including Sojourner, Spirit, Opportunity and Curiosity.
1.3 Aims and Objectives
The long term objective of the project is to create several concepts for small rover designs. The rovers are to be used in deploying a number of low frequency radio antennas on the far side of the moon. The goal of this deployment is to create a large lunar array capable of detecting radiation from the Cosmic Dark Ages without the disturbance of human interference.
The aim for the 2014 energy engineering students is to develop a concept energy system for the rover. The students aims to answer questions such as how the rover can be supplied with energy and what storage of this energy is possible.
1.4 Problem Statement
• How much energy can be extracted from solar irradiance on the moon?
• What are the power requirements of a rover? Can enough energy be supplied for continuous operation?
• How much energy can be stored on board the rover in batteries onboard the rover? • What environmental conditions needs to be addressed for the rover to be able to
work continuously?
1.5 Assumptions and Delimitation
Since this project takes place early in the expected mission timeline, there is a signif-icant lack of knowledge concerning several key factors of the mission and rover. Some assumptions and deliminations are therefore necessary.
• Previous student groups have assumed delivery will be performed by an Atlas V 401 4S. We will assume usage of this system as well.
• The deployment site has been prepared prior to arrival.
• This project does not concern electronics and autonomy. However, survivability of electronics will be discussed.
Furthermore, this project will not attempt to create a detailed construction plan. Due to time constraints and the ever speeding development of technology, this report limits itself to dimension the physical properties that the energy system could utilize.
2 Methods and Materials
2.1 Work Schedule
In order to ensure that there was a sufficient understanding of the environment of the moon and what a rover is, this report contains a short literature study of the moon and rovers used during the previous exploration of planetary bodies. This study has been conducted in literature and databases and was completed in April 2014. Since the project takes place early in the design stage of the rover, there were no components available for study on power requirements. In order to calculate how much power the rover needs, a mock up list of assumed components has been created using specifications from other recent rovers, such as the Mars Exploration Rovers and the Mars Science Laboratory. The power input from solar panels varies during the day. It is therefore important to examine how long the photovoltaics cells can provide energy to power the rover. This also helped show how much batteries were needed to store the energy. A simulation has been used for this purpose.
During the project period, there were three planned seminars. During these seminars, the project group gave a presentation of the state of the project and discussed ideas with other students. For the whole project period weekly meetings were held with the supervisors Professor Lars Bååth and Professor Jonny Hylander. During the meetings problems that had appeared were discussed. Discussions were also held via telephone and mail with Dr. Ashitey Trebi-Ollennu of JPL during the project.
2.2 Sources and Materials
For the literature study scientific articles, theses, web sites and relevant literature have been used. The main method of finding relevant sources was from database searches. Databases used were, among others NASA’s databases, Halmstad University library database. For simulation purposes, the program Simulink has been used. Simulink is part of Matlab and provides an toolbox for solving physics and engineering problems.
3 Literature Study
To confidently perform the task at hand, the following three aspects were researched: the radiation environment around the moon, the lunar topography with consideration to mountain ranges and shadows, and a detailed overview of previous rover missions.
3.1 Information Regarding the Moon
In this section, the enviromental conditions of the moon are addressed.
3.1.1 Cosmic Radiation and Solar Irradiance
There are three major types of particle radiation on the moon: solar wind, galactic cosmic rays and solar cosmic rays. Other types of radiation exists, but are too insignificant to be accounted for. The moon is very exposed to all types since it does not have a strong magnetic field or a thick atmosphere to absorb any part of the radiation. The largest decider of which type of particles that will be present at any given time is the sun, partly because it emits radiation itself, partly because it also creates a magnetic field protecting our solar system from other forms of radiation. The strength of the magnetic field is dependent on the activity of the sun, which varies in an 11-year cycle.The common particles in the radiation are atomic nuclei and electrons with energy levels between 0.3 eV and 10 GeV. The penetration and effect of these particles varies [2].
Table 3.1 – Penetration and energy levels of particles. Source: [2], page 48.
Type Solar Wind Solar Cosmic Rays Galactic Cosmic Rays Nuclei energies ∼0.3-3 keV/u ∼1 to >100 MeV/u ∼0.1 to >10 GeV/u Electron energies ∼1-100 eV <0.1 to 1 MeV ∼0.1 to >10 GeV/u Fluxes (protons/cm2sec) ∼3 × 108 ∼0-106 2-4
Particle ratios
electron/proton ∼1 ∼1 ∼0.02 proton/alpha ∼22 ∼60 ∼7 Light (3≤Z≤ 5)/alpha n.d. ∼0.0001 ∼0.015 Medium (6≤Z≤9)/alpha ∼0.03 ∼0.03 ∼0.06 Light Heavy (10≤Z≤14)/alpha ∼0.005 ∼0.009 ∼0.014 Medium Heavy (15≤Z≤19)/alpha ∼0.0005 ∼0.0006 ∼0.002 Very Heavy (20≤Z≤29)/alpha ∼0.0012 ∼0.0014 ∼0.004 Very Very Heavy (30≤Z)/alpha n.d. n.d. ∼3×10−6 Lunar Penetration Depths
protons and alphas <micrometers centimeters meter heavier nuclei <micrometers millimeters centimeters
Solar Wind
Solar wind is a flux of charged particles continuously emitted by the sun. It is emitted in all directions and is present everywhere in the solar system. It also creates the magnetic field that surrounds the solar system, called the heliosphere. Mostly consisting of electrons and protons, the wind is passing the Earth and the moon at speeds around 500 km/s. The solar wind is the main source of several elusive substances on the moon, including hydrogen, helium, carbon and nitrogen [2].
Solar Cosmic Rays
Solar Cosmic Rays (SCR), also called solar energetic particles are produced by the sun during large solar flares. These high energy particles are emitted by the sun during the solar cycles activity maximum. SCR are usually preceded by a corona mass-ejection [3]. Because of the high energy levels of the particles, they both travel and decay quickly. The most energized particles can reach earth’s orbit within an hour. Because of their quick decay, they are only present at the moon as radioactive particles for short periods of time. However, the radiation dosage from being exposed to SCR can be devastating to equipment and personnel. It is therefore important to shield equipment that needs to be active during solar maximum [2].
Galactic Cosmic Rays
Galactic cosmic rays (GCR) are similar in composition to SCR, but they originate from outside the solar system. The particles in this radiation can have accelerated within the
Milky Way’s magnetic field for several million years. During this time the electrons have been stripped from the atoms leaving only the nuclei in relativistic speeds traversing the galaxy [4]. High energy particles can be present in the solar system at any time, but since the sun’s magnetic field deflects some of the particles they are much more common during times of minimum solar activity [5].
Shielding from radiation
As shown by table 3.1 most particles will be stopped within the first centimeter of solid matter it will encounter. For the DALI rover, this will mean that the most exposed part to radiation will be the outer chassis, the Warm Electronics Box (WEB) and the solar panels on top of the rover. Among these three, the worst damage will be done to the solar panels. As stated in section 3.3.2, a solar cell is built up by a crystal structure on the atomic level. Long term exposure or exposure during solar maximum/minimum will reduce the effectiveness of the solar panel and reduce the time the rover can be operational. For this purpose, solar cells of a particle radiation resistant material and a mission planned at the middle of a solar cycle is adviced.
Solar Irradiance
Sunlight carries 1360 W/m2 to earth’s orbit. Before reaching the surface of our planet,
some of this energy is absorbed into the atmosphere, reducing the power to a mean value of 1000 W/m2. Since the moon orbits around the Earth, the moon mean solar irradiance
continues to be the same. The rover discussed in this thesis is to be placed on the far side of the moon. An argument can be made that this would increase the irradiance available according to the inverse-square law. A quick mathematical check shows that there is a difference of 0.5 per cent or 7 watts in mean power.
I · 1
r2 → 3.04 · 10
19· 1
(149600000 − 384400)2 = 1367 [W/m
2] (3.1)
WhereI is the intensity of the solar irradiance at the sun, andr is the distance between the sun and the moon when they are closest. Because of the variation of orbit, the irradiance is not constant. In this report 1367 W has been used as an average [6]. Traditionally, the amount of sunlight that reaches a celestial body’s surface is dependent on the thickness of the atmosphere and the latitude of the photovoltaic cell. Since the moon lacks an atmosphere, this affect can be ignored for calculations. The latitude of the solar cell will be briefly discussed in the next section, 3.1.2. The amount of power provided by solar power will therefore only be dependent on the inclination angle between the solar panels and the sun position in the sky. Since the rover will have flat solar panels, the azimuth angle (the angle between the solar cells and the horizon) does not come into affect. The moon’s day/night cycle is significantly longer than earth’s, 655 hours. The levels of solar irradiance will have a sinus form during this time, as shown by figure 3.1 [7]. The total amount of available energy will be the integral of this sine from dawn until dusk, which is equal to 283 kWh per square meter.
Figure 3.1 – Amount of solar power available at the moon’s surface per square meter.
3.1.2 Lunar Topography and Environment
Environment
The moon orbits the earth in 27.3 days and its own axis in 29.5 days. It also has a tilt to its rotational axis, giving it seasons like the Earth. This tilt is only 6.4, meaning that the effect is much less noticeable. Also, the rotational axis points 1.54 away from the northern elliptic pole. This gives a slow alteration of the moon’s day/night-cycle. During a period of 18 years, the night becomes longer until it reaches a maximum of 384 hours or 16 days. There is a severe difference in temperature between the day and the night. The solar radiated energy will heat up the day side to nearly 370 K, as shown by table 3.2. However, the lunar soil, called regolith, quickly radiates any stored energy when out of direct sunlight. This means that any shaded place will get cold very fast. At Apollo landing sites temperatures as low as 92 K have been recorded. Close to the polar regions there are areas never reached by the sun. These areas, usually impact craters, are kept at a constant temperature of approximately 40 K [8].
Table 3.2 – Measured temperatures and thermal conductivity of lunar surface and sub-surface. Source: [8], page 1986.
Tmax(K) Tmin(K) k(10−2W m−1K−1)
Subsolar point/midnight 374 120 ground based measurements (Pettit and Nicholson, 1930) Apollo 11 surface 395 93
Apollo 12 surface 389 88 0.115+ T on the moon calculated, using measured k of returned samples on the ground (Cremers et a., 1971). depth of 2 cem 320 145 +0.159·10−8
T3
depth of 10 cm 225 230 (Samples)
Apollo 15 surface 374 92 ∼ 0.15 measured with the HFE (Langseth et al., 1972, 1973); depth of 49 cm 255 248 1.40 (±0.14) value of k for the top 2 cm material
depth of 91 cm 251 1.70 (±0.17) (15·10−4W m−1K−1)deduced from diurnal surface
depth of 138 cm 253 2.50 (±0.25) temperature variation.
Apollo 17 surface 385 103 ∼ 0.15 measured with the HFE(Langseth et al., 1973); values of k depth of 130 cm 255 2.5/2.06 (±0.37) for probe 1 and 2, respectively.
depth of 233 cm 256 2.95/2.24 (±0.44)
Permanently shadowed 80 ground based measurements (Wendell and Low, 1970; polar crater ∼ 40 Burke, 1985; Keiken et aL., 1991).
Topography
The terrain on the moon can be divided into two main groups, highlands and lowlands. The highlands reach peak altitudes of about 10 000 meters and the lowlands delve down
to a depth of 9000 meters from the mean height. The terrain has been formed during the millennia by two main factors, volcanism and impacts with asteroids and other space faring debris. Since there is no atmosphere, once an impact crater or a volcanic mountain has formed, the lack of erosion will preserve them indefinitely. This has given the moon a very rough surface. The Apollo 11 crew reported that the area around their landing site was filled with craters varying from 0.3 meter up to 20 meters in diameter. Later Apollo missions reported similar or even larger craters at their sites. The largest craters present at the moon can be over 300 kilometers in diameter. However, astronauts visiting the moon have reported that even fresh craters lacks a “sharp edge” [2].
The highlands can be identified as the lighter areas when looking at the moon. The lunar lowlands are darker in color and are often impact basins from large collisions which has unearthed the lava deep inside the moon. In the early days of astronomy the darker lowlands were mistaken for seas and given the name lunar marina. Modern telescopes and ultimately visits have debunked this misunderstanding. The name however, stayed, and is in use today. The lava in these marinas are usually relatively flat surfaces. Recent studies have found that the lava flows in the marinas are considerably younger than the craters, meaning that these surfaces are flatter than many other places on the moon. The moon’s surface consists of up to about 17 % of marina basins. However, most of these are located on the near side of the moon, leaving the far side with less than 1 % [2].
Figure 3.2 – Location of lunar marinas. Source: [2], page 96.
Previous student groups identified possible deployment sites based on the surface tem-perature, smoothness and how large smooth area existed. The ideal place for deployment would be a marina, but these are rare on the far side of the moon. Three suitable loca-tions were identified: the Mare Moscoviense, Aitken crater and Tsiolkovsky crater. An important factor that wasn’t discussed last year is the shading of the deployment site which affects the amount of sunlight available to the rover. Using a height map provided by NASA’s Lunar Reconnaissance Orbiter, an estimation of nearby mountains can be done [9].
Mare Moscoviense.
• Average crater basin to mountain peak height difference ≈ 4,400 m. • Surrounding mountain height ≈ 3,000 m.
Aitken crater.
• Central location on the far side. 16.8◦S, 173.4◦E
• Smallest proposed crater ≈ 15 km smooth space.
• Average crater basin to mountain peak height difference≈ 3,000 m. • Surrounding mountain height ≈ 4,000 m.
Tsiolkovsky crater.
• South-west location on the far side. 20.4◦S, 129.1◦E
• Largest smooth space ≈ 50 km.
• Average crater basin to mountain peak height difference≈ 3,900 m. • Surrounding mountain height ≈ 3,000 m.
An average height difference of 3,750 meters can be set for all these locations. However, equally important is the size of the crater. It would be possible for the rover to drive away from a deployment site near a mountain to be able to generate power faster after the night. For optimal solar conditions, a central location in all craters is recommended. To aid in selection of which deployment site is suitable, a calculation was performed for when the sun will climb over the surrounding mountains.
Figure 3.3 – Sunlight hitting the center of the crater.
With knowledge of the surrounding mountain height and the crater width, the angle and thus the amount of time with solar energy can be calculated.
θ = arctan(heightwidth
2 ) (3.3) tsun= Θ 360◦ = (180◦− 2θ) 360◦ (3.4)
For the suggested craters, the amount of sunlight per day is distributed according to table 3.3. It is clear that the most solar efficient location is the Tsiolkovsky crater which has sunlight for 45 % of a lunar day/night cycle.
All craters are within 30◦ of the equator, meaning that there will be an abundance of
sunlight available to the rover. The average energy provided at the craters is 92 % of the energy at the equator. The most distant crater, Mare Moscoviense, will still receive 89 % of the energy it would have been receiving at the equator. This can be calculated as the cosine of the latitude.
Table 3.3 – Sunlight percentage.
Aitken crater Mare Moscovinse Tsiolkovsky crater
Theta, θ
21.8
◦16.3
◦8.9
◦Sun time, t
38 %
41 %
45 %
Latitude, Cos(ϕ)
95.7 %
88.9 %
93.7 %
3.2 Previous Rovers
In this section previous rovers used on celestial bodies other than earth are described.
3.2.1 Lunokhod 1
Lunokhod 1, which in Russian means moon walker, was the first unmanned rover to drive on a celestial body other than the earth. On November 15, 1970 the spacecraft Luna 17 carrying Lunokhod 1 soft-landed on the moon in the Sea of Rains. The Lunokhod 1 then descended to the moon’s surface [10].
The rover was designed in the Soviet Union and it had a mass of 756 kg [11]. The chassis was shaped like a tube-like compartment with a large convex lid. The rover moved on eight independently powered woven-wired wheels, with a maximum speed of 1 km/h [12]. A pair of TV-cameras in the front were used for the rover’s operators, while four corner cameras created high-resolution panoramic images [11], [10]. Lunokhod was also equipped with antennas plus different instruments for research.
On the underside of the convex lid a solar cell array was mounted [13]. It was of the type Gallium arsenide solar cells (GaAs) [14]. During the lunar day, the lid was unfolded and powered the drive-wheels and instruments inside the compartment [10], [13]. To protect the rover from the cold during the two-week lunar night the lid would close. As heating, to withstand temperatures of -150◦C [11], a Polonium-210 radioactive power
source was installed into the back of the Lunokhod. The radioactive power source powered a generator which heated the gas inside the insulated and pressurized compartment [13]. On the lunar day the temperature could reach 130◦C [14]. To avoid heating issues, the
vehicle was equipped with a ventilation system. A gas circulated in the compartment and discharged the heat outside by a thermal exchange unit [13].
Figure 3.4 – Lunokhod 1. Source: [15].
3.2.2 Lunokhod 2
Over two years after the first successful lunar rover mission, the spacecraft Luna 21 carrying Lunokhod 2 was launched. On January 15, 1973 Luna 21 landed on the Moon inside the Le Monnier crater [16].
Lunokhod 2 was an upgraded version of the Lunokhod with better and more equipment. The rover had a mass of 840 kg [16]. The speed had been improved to a maximum of 2 km/h.
Just like the previous Lunokhod, the second edition was controlled remotely from a team on Earth. The mission was to explore the moon and perform a variety of experiments [16].
Figure 3.5 – Lunokhod 2. Source: [16].
3.2.3 Lunar Roving Vehicle
The Lunar Roving Vehicle (LRV) was a rover intended to transport astronauts around the moon. Three LRVs were sent to the moon during different Apollo missions [17]. The total mass of the Lunar Roving Vehicle was about 210 kg. It had been designed to carry an extra mass of 517 kg on the moon, which included two astronauts plus equipment etc. [18]. The vehicle had a length of 3.5 m, a wheelbase of 2.3 m and a maximum height of 1.14 m [17].
The LRV had four wheels, each powered by its own electric motor. The motors were of the type series wound DC Motor, with the specifications 0.25 hp and 10 000 rpm. To prevent overheating, each motor was thermally monitored over a thermistor on the stator field. If the temperature reached 204 ◦C a thermal switch, placed on each motor, would stop
the vehicle [19]. Two 36 V non-rechargeable batteries, type silver-zink, with a capacity of 121 Ah powered the LRV. The batteries provided redundancy as each battery could power the whole vehicle [17]. Speed control was made possible through a pulse-width modulation frequency converter [18]. Depending on the surface and terrain, the vehicle’s top speed varied between 9 to 13 km/h [19].
Rover 15 was driven by the astronauts David Scott and Jim Irwin. The total operational time was about three hours, which transported the LRV 27.8 km [17].
Rover 16 was the second LRV on the moon. It was brought by the Apollo 16 lunar module, Orion. On April 21, 1972 the Lunar Module landed on the moon [21].
The astronauts John W. Young and Charles M. Duke Jr. drove Rover 16 during 3 hours and 26 minutes. It traveled a total distance of 4.5 km [17].
Rover 17 was the third and last LRV on the moon. On December 11, 1972 the Apollo 17 lunar module landed on the moon [22].
The rover operated for 4 hours and 26 minutes and was driven 35.9 km. Rover 17 was driven by the astronauts Gene Cernan and Harrison Schmitt [17].
Figure 3.6 – Apollo 15 LRV. Source: [17].
3.2.4 Sojourner
Sojourner, alternatively Mars Pathfinder Rover [23], was the first autonomous rover to be sent to the planet Mars [24]. Because of its small size, Sojourner has often been described as a micro-rover.
The micro-rover was included in the Mars Pathfinder mission which was part of the Discovery program. The primary objectives were exploration of the Martian surface and demonstrating the feasibility of low-cost landings on Mars [23]. The Mars Pathfinder lander (MPF) landed on Mars on July 4, 1997 [25].
The rover had the mass of 11 kg. It had six wheels and the top speed of 0.036 km/h (0.6 m/min). The six-wheeled rocker bogie design provided Sojourner with great stability, and allowed the micro-rover to tip as much as 45◦ on one side without tipping over [24].
Sojourner was powered by a 0.22 m2 GaAs solar panel. Nine D sized non-rechargeable
lithium batteries backed up the solar panel, and provided up to 150 Wh [25]. Primary batteries had been chosen to minimize the weight and the cost. Once the batteries were used, Sojourner could only operate at times of the martian day when the solar panel generation exceeded the driver power requirement [26].
The micro-rover’s driver power requirement was about 10 W. The combination of a solar panel and a battery system allowed the micro-rover system to draw up to 30 W during zenith (mid-solar) [25].
During the martian night, temperatures of - 110◦C are common. A warm electronics box,
abbreviated WEB, contained and protected Sojourner’s components from the ambient Mars temperatures [25].
Figure 3.7 – Sojourner on Mars. Source: [23].
3.2.5 Spirit and Opportunity
The Mars Exploration Rovers, Spirit and Opportunity, were two twin rovers in the Mars Exploration Rover (MER) mission. On June 10, 2003, the rover Spirit launched towards Mars. Almost a month later on July 7, 2003, Opportunity followed. The twin rovers landed on Mars January 4 respectively January 25, 2004. Today, Opportunity is still
active [27]. Spirit’s mission ended on May 25, 2011, due to lost contact [28]. Previously, Spirit had been trapped in soft ground [29].
The scientific objectives of MER were to search for evidence of past water activity and examine the geology [30].
Each Mars Exploration Rover had a mass of 174 kg and was 1.6 meters long and 1.5 meters high [31].
Partially, the designs of the twin rovers were based on the previous Sojourner rover. Similar to Sojourner, the rovers had six wheels, a rocker-bogie suspension [31] for driving over rough terrain and a warm electronic box (WEB) containing important components [32]. Each wheel was attached to its own electric motor. The two front and two rear wheels were steerable, independently [32]. The top speed of the twin rovers was about 3.75 cm per second (0.135 km/h) [32].
Power was generated by solar arrays. During full sun conditions the solar arrays could generate up to 140 W. Two rechargeable batteries backed up the power [32]. The batteries were of the type Li-ion and each contained eight prismatic, 10 Ah, Li-ion cells [33].
Figure 3.8 – Opportunity. Source: [31], page 2.
3.2.6 Curiosity
Curiosity, also called Mars Science Laboratory (MSL), is a large rover that was sent to Mars on November 26, 2011. The rover landed on the martian surface on August 6 the year after launch and is still active at the time of writing. The objective of MSL is to search for evidence of current or past habitat of life on Mars [34]. MSL’s mission is part of the NASA’s Mars Exploration Program [35].
MSL’s size is the largest of all active rovers today. It has the length of approximately 2.8 meters and the mass of 750 kg. Six wheels transport the rover, where the front
and rear wheels have individual steering motors. The wheels are mounted on a rocker-bogie suspension. The body consists of a WEB, which is necessary for protecting the vital components. The rover is also equipped with cameras, a mast and other scientific instruments [34].
Curiosity is powered by a radioisotope thermoelectric generator (RTG). At the start of the mission the RTG provided 125 W. A heat exchanger is used for the thermal control [34].
Figure 3.9 – Curiosity. Source: [35].
3.2.7 Yutu
On December 2, 2013, the spacecraft Chang’e-3 was launched towards the moon. [36]. The spacecraft carried a rover named Yutu, which in Chinese can be translated as “Jade Rabbit”. The spacecraft landed on the moon on December 14, 2013 [36],[37].
Chang’e-3 lunar mission is part of China’s lunar program on behalf of the Chinese Na-tional Space Agency (CNSA). The objectives of Chang’e-3 are to explore the moon, examine the lunar soil and practice soft-landings [36].
Yutu has a mass of 120 kg and a payload of 20 kg [37]. The rover is equipped with cameras and instruments. The power is provided by solar panels and backed up by batteries [36]. In order to save energy during the 14-days long lunar night, Yutu hibernates. Radioisotope heater units (RHU) plus two-loop heat pipes are used for heating [38]. During the mission Yutu has had operational problems, mainly temperature related [39].
Figure 3.10 – Yutu. Source: [37].
3.3 Halmstad University Rover Concept
As stated in section 1.1, the purpose of the rover is to deploy a radio array consisting of six dipole arms, 100 m long. The rover will then park in the middle of the array and act as a transmission hub, sending packets of collected data to earth. The rover itself can not transmit to earth, but will instead transmit to another ground station or an orbiting satellite. The students of 2013 created a flow chart for the rover’s mission. Since the deployment phase will require much more energy than the later transmit phase, the solar cells will need to be dimensioned accordingly.
Figure 3.11 – Initial process schematic.
Previous students have created concepts for the rover’s deployment mechanism, driving mechanism and the chassis materials. However, no work has been done on actually dimensioning these parts. The total mass of the rover was estimated to 160 kg. There is no data on important energy dependent systems such as the Warm Electronics Box (WEB), computer, transmitters or electric drive motors. Therefore, many components will have to be estimated to see how much energy they will need. These estimates will be based on previous rovers’ power consumption and advice from Dr. Ashitey Trebi-Ollenu. There are five drive modes that the rover needs to accomplish to be able to deploy the antennas:
1. Communicating with earth, for new instructions. 2. Drive to different sites on the moon.
3. Pick up rolls and lift them.
4. Roll out rolls in the correct place. 5. Hibernation to survive lunar night.
These five modes can be linked to specific energy needs, as seen by appendix A. To ensure an uninterrupted deployment, the solar cells and/or batteries must always provide enough energy.
The rover’s main computer will be a RAD 750 single board computer. It usually uses 15 W of power. However, it can power down to one tenth of that during hibernation [40].
Figure 3.12 – Concept picture of the Halmstad University Rover.
3.3.1 Warm Electronics Box
Electronics and other power using components are often sensitive to heat and cold. Due to the moon’s temperature variations, a Warm Electronics Box (WEB) is necessary to protect the electronic components from the extreme cold and heat. The WEB is striving to keep the temperatures rather constant, at levels all the components can survive. It is basically an insulated box which houses all temperature sensitive equipment. For the Halmstad University rover, a temperature span between -20 to +30 C is required. To achieve this, the WEB is heated during the night and cooled during the day. Components inside the WEB will aid in heating during the night. Additional heat can be provided by a heat generating resistor or by a radioisotope heating unit (RHU) such as the one used by Lunokhod 1 & 2. The amount of heating/cooling required is dependent on the thickness and efficiency of the insulation. Previous student group Johan Winberg & Tomislav Stanimirovic concluded in their report that aerogel can be used as insulation for the WEB. Their conclusion builds on NASA’s own research and the fact that aerogel has been used as insulation on space missions since the Mars pathfinder mission. Aerogel is a gel where the water has been dried out. The result is a light weight material with exceptionally good insulating properties [41]. The inside of the WEB will need to have a volume capable
of housing batteries and various electronics. Aerogel’s thermal conductivity varies with the material chosen and the outside temperature.
3.3.2 Photovoltaic Power Generation
A photovoltaic cell (usually called a solar cell) utilises the photoelectric effect to create electricity from sunlight. This effect works on the atomic level by exciting the electrons which then can be captured and fed into an electric circuit. A photon hits the material, usually a semiconductor such as silicon doped in phosphorous and boron. The amount of energy collected by the cell is dependent of the energy level of the semiconductors band gap. Photons that have an energy level equal to the band gap will excite an electron to leave the crystalline structure. When granted a way to the positive pole the electrons will form a current. This current can then be used to power a load such as an electric motor or a lightbulb.
A single solar cell will only provide a small voltage, <1V. The voltage can be increased by serial coupling the cells into a solar module. Through adapting the amount of cells in the module, different voltage levels can be acquired. The current generated is dependent on the amount of solar irradiance. The amount of power provided by a solar cell depends on several losses from quantum theory, optical losses, recombination and resistance losses [42]. It was during the “space race,” when the need for a reliable source of power in space that the development of solar power became interesting. Since then solar power has been used on almost every space flight within the inner solar system. The International Space Station (ISS) features eight solar arrays, each 73 meters long. Together they generate 84 kW of power at 14 % efficiency [43]. For space operations, the key features of a solar cell is power per kilo and efficiency. A high efficiency solar cell may get a lot of power, but it will also have a greater mass, increasing the amount of fuel needed to send it into orbit.
There are several different techniques available and in research at the time of writing. Since the launch date is several years into the future there is no point in selecting a specific cell to be used. However, since the physical requirements of a solar cell to survive on the moon will not change, it is possible to list potential candidates to help focus research effort and determine how much energy will be available to the rover.
There are three main techniques used as layouts for solar cells:
1. Single junction cell. This is the most common and oldest technique. Usually a semiconductor doped to create a p-n junction that can absorb some of the sun’s energy. However, all photons with energies not matching the band gap will be lost. Single junction is the most common technique on sale, but does not produce as much power as multi junction cells.
2. Multi junction cells, which uses several layers of different semiconductors to absorb a larger portion of the photons emitted by the sun. Multi junction cells are very commonly used in space since they can achieve much higher efficiency than other solar cells. The Mars exploration rovers used triple junction cells [45].
3. Thin film solar cells are made by spreading a thin layer of semiconductors over a material. The result is a highly flexible solar cell that can easily be several hundred meters long. It can also be rolled up for storage. Since the layer can be so thin, a thin layer solar cell can and are used in builing panels and windows. It does however have low efficiency. While thin film would be unpractical to use on the rover, it is possible to be used later in the mission when the DALI radio array has been deployed. An additional roll of thin film solar cells could be deployed to generate extra power [42].
Difficulties posed by the Lunar Environment
As have been shown in chapter 3.1, the moon is a very hostile environment to Man and equipment. For solar cell’s, there are several dangers that can cause damage to them or otherwise prevent them from functioning as intended:
• Temperature. The high difference between day and night temperatures can seriously damage the solar cells.
• Radiation. The cosmic and solar radiation may destroy the solar cells structure. • Dust. The dust from the moon could potentially cover the solar cells.
The high temperature will cause the semiconductor’s band gap to decrease, causing a loss in efficiency [46]. The band gaps given for different materials are measured at 300 K. On the moon the band gap will be lower in accordance with equation:
EG(T ) = E0−
α · T2
T − β (3.5)
Where EG is band gap dependent on temperature, E0 is the band gap at absolute zero,
α is eV/K, β is a constant and T is temperature in Kelvin. An example can be seen for Gallium Arsenide in figure 3.14.
250 300 350 400 Temperature 1.38 1.40 1.42 1.44 1.46 eV
Figure 3.14 – Gallium Arsenide band gap dependency on temperature.
The high temperature will also cause the open voltage circuit (VOC) to decrease, giving a
lower power output. When estimating the amount of power available to the rover, these factors will need to be addressed [47].
High energy particles from different sources clearly have the potential to damage the solar cell’s crystalline structure. This will in time decrease the amount of power the solar cells can produce and might be what finally ends the rover’s life. To increase the life expectancy of the rover, it will be important to select a semiconductor that is resistant to radiation.
The lunar dust can be read about in great detail in 2013’s student rover reports. Since the solar cells will be covered during landing and the rover will not be travelling at any great speeds, dust build up on the solar cells will be neglible.
Silicon
Silicon was among the first materials used in solar cells and in space. Silicon solar cells have therefore been the focus of an abundance of research and are very dominant in the private solar cell market on earth. The band gap of silicon is 1.12 eV. Silicon is a crystal material which can be used in all three solar cell layouts. The material can be used in three ways; as monocrystalline, polycrystalline, or amorphous silicon. Mono- and polycrystalline can be used in single junction cells, and monocrystalline could also be used in a multi junction cell. The efficiency of a single junction silicon cell is usually low with a laboratory measured efficiency of about 28 %. As shown by figure 3.13, polycrystalline is less efficient than monocrystalline. Amorphous silicon can be used to make thin film solar cells [48].
Copper Indium selenide
Copper indium selenide (CIS) is a semiconducting material with a low band gap of only 1.1 eV. It can not form an effective n-p junction on its own, but can be combined with another semiconductor such as cadmium sulphide (CdS). It can also be infused with gallium to form copper indium/gallium (di)selenide (CIGS). This material has a higher band gap, 1.7 eV. The VOC is also raised to 0.5 V, on par wih silicon. CIS can be used
when making multi junction cells, while CIGS is used for thin film solar cells with an efficiency of around 20 %. The main benefit of CIS is that it requires a very thin material, reducing the mass of the solar cells needed [42].
Gallium Arsenide
Gallium arsenide (GaAs) is a very promising material for space exploration. It is more resistant to damage from radiation when compared to silicon. The band gap of 1.42 eV also allows for absorption of more photons. GaAs is also a good material for multi junction cells, usually combined with germanium and gallium indium phosphide. Triple junction cells are among the most effective, capable of reaching efficiencies of 39 % as shown by figure 3.13. Thanks to the high efficiency, the mass of the solar cells are comparatively low [42].
3.3.3 Energy Storage
Electric batteries are devices consisting of one or more electrochemical cells connected together, and packaged into a finished product. The basic idea of batteries and electro-chemical cells is to convert electro-chemical energy into electrical energy [49].
Each electrochemical cell consists of three active components: anode, cathode and elec-trolyte. Both the anode and cathode are electrodes (electrical conductors), but depending on the electrical potentials they are named differently.
During discharge the anode is the negative side of an electrode, where electrons are released from the cell. The cathode is the positive side of an electrode at which electrons enter the cell during discharge. The tip of the ion electrode is immersed in an electrolyte solution. The electrolyte contains mobile ions which makes the substance electrically conductive. Positively charged ions, named anions, move from the cathode to the anode. Negative ions, named cations, move the opposite direction. A separator divides the anions and cations and let them through in a pace not too fast. This is done to obtain as much energy from the electrochemical reaction as possible during as long time as possible. As the anions and cations move through the separator, the anode builds up negative charge and the cathode builds up positive charge. This results in the cell voltage V(t) as shown in figure 3.15. Through the circuit, negatively charged electrons flow from the anode to the cathode. This creates a current over the resistance in the opposite direction [49], [50].
Figure 3.15 – Image of a simple electrochemical cell.
State of charge (SOC) measures the remaining capacity or energy remaining in a cell or battery.
Depth of discharge (DOD) is the opposite to SOC (DOD = 1 − SOC). DOD describes how much capacity or energy that is removed from a fully charged battery [49].
Difficulties
As described in chapter 3.1, the lunar environment can be very strenuous. For electric batteries this is not an exception, two parameters excel:
• Temperature. At temperatures of -150 to 150◦C there are few batteries that survive.
A normal li-ion battery has a temperature range of -20 to 40 ◦C.
• Radiation. As described previously, radiation can harm electric components. The solution is to place the batteries in a warm electronics box (WEB). The WEB protects the components from radiation, and controls the temperature between given limits [51]. Apart from problems due to the moon’s climate, there are other problems as well. A lithium battery has the following age-related problems:
• Solid–electrolyte interface (SEI) layer growth. The impedance rises because the SEI layer grows on the negative side of the electrode.
• Lithium corrosion. Over time, lithium can corrode. This results in capacity fade due to loss of mobile lithium.
• Contact loss. Contact loss occurs as SEI layer disconnects from the negative elec-trode. This also leads to an increased cell impedance.
• Lithium plating (deposition). At low temperatures, high charge rates, and low cell voltages, the lithium metal can plate on the negative electrode. This causes irreversible loss of cyclable lithium.
Redundancy
Redundancy in engineering refers to duplicating or multiplying critical components in order to increase the reliability of the system [52], [53]. In the battery system, this means multiple batteries should be coupled in parallel where each of the batteries should be able to provide power without the other batteries functioning.
At NASA the term “Design for Minimum Risk” (DFMR) is used for describing and designing a redundant and failure-tolerant system [54].
Batteries in Space Applications
Batteries are important components in space applications. In the beginning, nickel-cadmium batteries (NiCd) were used for satellites. Later on, in the 1980s, nickel-hydrogen (NiH2) batteries began to be used in satellites because of their long lifetime [49].
As described in chapter 3.2, the battery types used in previous rovers have been either silver-zinc or lithium-ion.
Silver-zinc
Silver-zinc batteries exists as both primary and secondary batteries. An advantage of this battery type is a high energy density at a relative low cost. It is also more environmental friendly than batteries made of materials like cadmium, lead, and mercury [55]. The battery discharge reaction for a silver-zinc oxide battery is: Zn + Ag2O → ZnO + 2 Ag
[49].
Lithium-ion
A lithium-ion battery is a rechargeable battery type, known for its high energy density. In space missions volume and mass tends to be very valuable. Li-ion batteries also have a low discharge rate and a long cycle life [50]. The chemical battery reactions are discharge from left, and charge from right: C6+ LiM O2
→
← LixC6+ Li1−xM O2 [49].
Table 3.4 – Comparison of NiCd, NiH2, Silver-Zinc, and Li-ion performance.
Sources: [51], page 2; [56], page 3.
Nickel-Cadmium Nickel-Hydrogen Silver-Zinc Lithium-Ion Specific energy (Wh/kg) 25 30 ∼100 >100 Energy density (Wh/L) 100 50 ∼150 >250 Battery Mass for 300 Wh MER (kg)† 33 28 11 6
Battery Volume for 300 Wh MER (lit)† 9 17 6 2.2 Cycle life (50% DOD) >1000 >1000 <100 >1000 Wet life (Storeageability) Excellent Excellent Poor Good Self-discharge rate (per month) 15% 30% 15-20% <5% Low temperature performance Moderate Moderate Moderate Excellent Temperature range,◦C -10 to +30 -10 to +30 -10 to +30 -20 to +40 Charge efficiency, % 80% 80% 70% ∼100%
4 Results
4.1 Rover Power Usage
Warm Electronics Box
In order to be able to perform a simple estimation of the thermal losses Φ, the material Spaceloft® Subsea from Aspen Aerogel will be used [57]. At 4T = 150 C between hot and cold sides, the thermal conductivity k of this material is 16.5 mW/m-K. The cross section area S will be the inner area of a rectangular box, 1 × 1 × 0.5 m. The thickness 4x of the insulation will be the unknown in equation 4.1. In table 4.1 the amount of heat transfer at different thicknesses is shown.
Φ = k · S · 4T
4x (4.1)
Table 4.1 – Insulation depending on thickness.
4x 10 mm 50 mm 100 mm 150 mm 200 mm 300 mm
Φ 990 198 99 66 49.5 33
With 300 mm of aerogel, only 33 watts will leak through the WEB. Thus, 33 W of heating will need to be provided during the coldest part of the night. Similarly, 33 W of cooling must be provided during the day. The heating of the rover during the night can be accomplished using a heat generating resistor (HGR) or a radioisotope heating unit (RHU). An advantage the HGR holds over the RHU is that the former can provide the exact amount of power needed at any given moment. It can also be turned off during the day, mitigating the need for cooling away the extra heat generated by a RHU during the day.
Power Usage
Table 4.2 estimates the operational time and power usage for each mode. For a detailed list of energy intensive components see appendix A. Since there is a significant difference in the rover’s power need depending on if it utilises a HGR or a RHU, both of these are listed. Furthermore, the rover will need to charge its batteries. This will result in charging losses.
Table 4.2 – Time estimation.
Drive cases Communication Drive to site Pick up roll Roll out roll Hibernation Total Heating RHU HGR RHU HGR RHU HGR RHU HGR RHU HGR RHU HGR Total energy need (W) 74.5 41.5 169 136 122 89 199 166 5 38
% of time 3 % 21 % 6 % 15 % 55 % 100 %
Hours of a moon day 20.2 141.2 40.3 100.9 369.6 672 Energy need per day (kWh) 1.50 0.84 23.9 19.2 4.92 3.59 20.08 16.75 1.85 14.04 64.44 42.23
A Simulink model was used to determine how long the rover would be able to support different drive cases. The model consists of a solar array linked to an active power load and a rechargeable battery. The simulation’s purpose was to determine if 1. The rover could survive the night using a RHU/HGR. 2. Would the rover be able to run its drive cases on solar power, or would the batteries be needed to provide enough energy? A second simulation was also performed to see how the rover would perform under four different loads. The results of the different loads are shown below in table 4.3. The simulation ran for 2.4 million seconds, or 672 hours. The solar cell used for modeling was a Sanyo HIP-225HDE1. The battery used was a Lithium ion, 77 Ah. Voltage for the model was 34 V. The different power loads relates to the drive cases.
Table 4.3 – Simulation results.
Power load (W) Solar power time (h) Battery time (h)
5 334 447
50 316 61
100 296 23
200 255 11
As shown in tables 4.2 & 4.3, it is clearly possible to have an operational rover on the moon powered by solar energy. Furthermore, figure 4.1 shows that the rover will be able to perform all different drive cases. However, the rover can not use a HGR for heating during the night, since the batteries are too small. There would be a loss of time due to the rover not being able to run energy intensive drive cases (P>100 W) at night or during the early morning/late evening.
4.2 Power Generation
The photovoltaic panel used in the simulation, Sanyo HIP-225HDE1, is a mono-crystalline solar panel chosen for its VOC. The simulation model did not allow for individual input
of material abilities. Therefore the simulation does not exactly model the capability of energy generation of the rover, but serve as a ’proof of concept’ that solar energy is a viable power option. As shown by figures 4.2(a) & 4.2(b), the rover can perform its mission without having to use its batteries during the middle of the day.
The solar panels eventually used on the rover will need to supply a large amount of power with great efficiency regarding space and mass. As table 4.4 shows, the best alternative for this is a multi junction cell.
Figure 4.1 – Rover simulation at varying loads.
Table 4.4 – Materials for solar cells.
Technique Material eV† V
OC Wpeak/kg Wpeak/m2 η
Multi junction Gallium Arsenide 1.35 2.7 1607 518 0.379
Single junction Monocrystalline Silicon 1.08 0.5 242 342 0.25 Thin film Copper Indium/Gallium selenide 1.5 eV 0.5 119 284 0.208 † Calculated at 450 Kelvin using eq. 3.5
0 0.5 1 1.5 2 2.5 x 106 -200 0 200 400 600 800 1000 1200 1400 Irradiance (W/m^2) -5 0 5 10 15 20 25 30 35 40 45 <Voltage (V)> 0 10 20 30 40 50 60 70 80 90 100 <SOC (%)> Time offset: 0 (a) 100 W load. 0 0.5 1 1.5 2 2.5 x 106 -200 0 200 400 600 800 1000 1200 1400 Irradiance (W/m^2) -10 0 10 20 30 40 50 <Voltage (V)> 0 10 20 30 40 50 60 70 80 90 100 <SOC (%)> Time offset: 0 (b) 200 W load.
Figure 4.2 – Solar power simulations. The x-axises show the time in seconds and the y-axises show the SOC percentage, voltage and irradiance. Higher resolutions are available in appendix D.
4.3 Energy Storage
The battery size is based on the power consumption during hibernation. As a safety factor, the power requirement has been multiplied by 1.25, giving 1.85 · 1.25 = 2.31 kWh for RHU respectively 14.04·1.25 = 17.55 kWh for heat generating resistor (HGR) useage. Combining this with table 3.4, assumptions of the weight and volume can be made. The values are proportional to corresponding values in rows in table 3.4. The results are presented in table 4.5 and take into account both efficiencies and two parallel coupled batteries. The whole battery system will therefore have twice as much energy as the need, but one battery will only be used at a time.
Table 4.5 – Volumetric and mass properties of batteries.
Nickel-Cadmium Nickel-Hydrogen Silver-Zinc Lithium-Ion 2.31 kWh
(RHU) 17.55 kWh(HGR) 2.31 kWh(RHU) 17.55 kWh(HGR) 2.31 kWh(RHU) 17.55 kWh(HGR) 2.31 kWh(RHU) 17.55 kWh(HGR) Battery Mass (kg) 254.1 1930.5 215.6 1638.0 84.7 643.5 46.2 351.0 Battery Volume (lit) 69.3 525.5 130.9 994.5 46.2 351.0 19.94 128.7
a nominal voltage of 30 V. The blue line corresponds to a load of 5 W and the green line corresponds to a load of 38 W. The discharge was also confirmed by the simulation as shown by figure 4.4 below. Finaly, figure 4.5 shows a simulation result determining the charge time of the batteries, which can be necessary during the early morning. The Simulation shows that the batteries will be fully charged after 28 hours.
0 2 4 6 8 10 12 x 105 0.156 0.158 0.16 0.162 0.164 0.166 0.168 0.17 0.172 36 36.5 37 37.5 38 38.5 39 39.5 40 <Voltage (V)> 20 30 40 50 60 70 80 90 100 <SOC (%)> Time offset: 0 (a) 5 W load. 0 2 4 6 8 10 12 x 105 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 <Current (A)> 0 5 10 15 20 25 30 35 40 <Voltage (V)> 0 10 20 30 40 50 60 70 80 90 100 <SOC (%)> Time offset: 0 (b) 50 W load.
Figure 4.4 – Hibernation simulations. The x-axises show the time in seconds, and the y-axises show SOC percentage, voltage and current. Higher resolutions are avail-able in appendix E.
0 2 4 6 8 10 12 x 104 0 50 100 150 200 250 300 350 400 31.5 32 32.5 33 33.5 34 34.5 35 35.5 <Voltage (V)> 10 20 30 40 50 60 70 80 90 100 <SOC (%)> <SOC (%)> <Voltage (V)> Irradiance Time offset: 0
Figure 4.5 – Battery charging with a load of 5 W. Battery: 30 V, 77 Ah, SOC 15% from begining. The x-axis shows the time in seconds and the y-axises show the SOC percentage, voltage and current.
5 Discussion
5.1 Power Usage
Two options were considered for heating the rover, using a radioisotope heating unit (RHU) which generates heat through radioactive decay or a heat generating resistor (HGR) which generates heat using electricity. There are two total energy needs shown in table 4.2 depending on which option is being used. The rover’s total energy need could therefore be either 64.4 kWh for a RHU or 42 kWh for a HGR. This would make it seem that the ideal heating option would be the HGR, since it requires less energy. However, the RHU needs energy for cooling during the day, when energy is abundant and easy to acquire. The HGR would require a solution to bring it energy during the night, when the only source of energy will be the batteries of the rover. This would require nearly six times more batteries than the RHU option and greatly increase the rover’s mass. A heavier rover would require more energy to run and heat, meaning that even more batteries would be needed. As it stands today, it is simpler to use a RHU for heating than a HGR.
5.2 Power Generation
This report has opted for a solar powered rover with a supply of batteries to survive the night on the moon. This is not the only option for powering a rover. The Mars Science Laboratory rover Curiosity uses a radioisotope thermoelectric generator (RTG). The RTG provided 125 W when it left earth and the decay of the radioactive medium will lower the power output continuously. Curiosity is expected to power itself for the duration of its mission, 2 years. However, the power source using plutonium-238 could still provide enough energy for 10-15 earth years of operation. The plutonium-238 provides 2000 W of thermal energy, which could also heat a rover on the moon during the night. The solar cells will decay from damage from radiation and possibly the extreme temperatures on the moon. This report lacks the theoretical framework for examining the survivability of the solar cells. This should be examined in later student projects. However with a particle radiation resistant material, the time frame could easily stretch into a decade or more. The most promising material to resist the bombardment of particles is gallium arsenide combined with other materials in a multi junction cell.
Curiosity was equipped with an RTG because the solar power on Mars was to weak to provide enough energy for a rover that size. Since the moon is much closer to the sun than Mars, solar power can provide enough energy. Solar cells are better for deployment near the earth since they have a high power to mass ratio. However, the 14 day long lunar night means that solar cells will require a sizable amount of batteries to survive
during the night, which might help heat the motors and other parts of the rover during the night.
The solar powered option’s main advantage is that it is easier to keep the solar cells cool during the day. The RTG will heat the rover internally with 2000 W. At the same time the sunlight will heat the outside with 1360 W/m2. Since an insulated WEB will need
very little energy to heat the inside during the night, but would require a lot of energy to cool during the day, there is a clear advantage to the solar cells.
Ultimately, this report is about examining the feasibility of a solar powered rover on the moon. Due to the short time the project has had to examine this, the RTG option of a rover has not yet been examined. This is an important component that must be examined closely in upcoming student projects.
In the most energy intensive drive case, the rover requires 199 W of power to sufficiently operate. During peak hours of sun, a multi junction solar panel of 0.5 m2 could provide
enough energy. But since the sun rises and falls in the sky, so does the amount of power that can be provided by the solar panel. Since the rover will need to have a full battery charge to survive the night, this poses an interesting question: How long each day should the rover to be operational? As can be seen in table 3.3, the amount of sunlight available each day is limited. If the solar cells are designed to provide 199 W at peak power, the rover would quickly have so low available energy that it could not continue to operate. Instead, it would have to wait until the next day, remaining stationary and using the power from the sun to hibernate. This would delay the mission and waste valuable time. Therefore, a full m2 of solar panels is recommended.
Figure 5.1 – Solar power provided by different amount of solar cells.
5.3 Energy Storage
For a rover placed on the moon with a combined photovoltaics and battery energy system, there are two governing constraints for the battery size selections; the lunar night and the
