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Master's Thesis

Master's Thesis

2009:003 MASTER'S THESIS

Information Gathering Strategies for a Combined System of and Rovers

Yujie Li

Luleå University of Technology Master Thesis, Continuation Courses Space Science and Technology Department of Space Science, Kiruna

2009:003 - ISSN: 1653-0187 - ISRN: LTU-PB-EX--09/003--SE Luleå University of Technology & Cranfield University

Li, Yujie

Information Gathering Strategies for a Combined Martian System of Aerobots and Rovers

Department of Space Science & School of Engineering

MSc. Thesis Luleå University of Technology & Cranfield University

Department of Space Science & School of Engineering

MSc Thesis

Academic Year 2006-2007

Li, Yujie

Information Gathering Strategies for a Combined Martian System of Aerobots and Rovers

Supervisor: Dr. Priya Fernando Dr. Stephen Hobbs

Academic Year 2006 to 2007

This thesis is submitted in partial fulfilment of the requirements for the degree of MSc in Astronautics and Space Engineering

© Cranfield University, Luleå University of Technology,2007. All rights reserved. No part of this publication may be reproduced without the written permission of the copyright holder. ABSTRACT

Future missions will involve aerial and surface . Previous studies have examined the possibility of using a single robotic in the Martian to obtain meteorological and topological data. It may be advantageous, however, to use a combination of aerobots along with surface rovers to map and study Mars. This combined system will be much more effective if the different elements are working in tandem to gather complementary information. The aim of this thesis is to investigate cooperative strategies for exploring a given piece of Martian terrain using a system comprising several different elements. Various concepts and rovers are reviewed, the capabilities of different types of aerobots and rovers were compared, and a trade-off was made to select a preferred system. Collaboration strategies developed between aerobots and rovers enable them to navigate and explore the target terrain with greater precision and reduced intervention from mission control on . An + tethered depot + nanorovers architecture resulted as the best resolution. A suite of instruments were configured to gather useful information. It is concluded that heterogeneous robotics with a complementary suite of instruments work in concert will greatly increase the mission value.

Keywords: Mars, Airship, Heterogeneous, Cooperation, Nanorover

i ACKNOWLEDGEMENTS

I would like to express my thanks to my thesis supervisor Dr. Stephen Hobbs for his invaluable advice and guidance throughout the project.

I would also like to thank my friends Niall Dooley, Raza Rizvi for many interesting discussions on a variety of topics and their patience in proofreading my thesis, and thanks to all my friends who have given me support, advice and some well timed distraction when I’m feeling down.

Finally I would like to thank my family for their endless support and encouragement that has enabled me to reach the end of my studies, which I could not have done without them.

ii TABLE OF CONTENTS

ABSTRACT ...... i ACKNOWLEDGEMENTS ...... ii TABLE OF FIGURES...... v TABLE OF TABLES ...... vi TABLE OF EQUATIONS...... vii 1 Project Introduction...... 9 1.1 Introduction...... 9 1.2 Objectives ...... 9 2 Mars Exploration...... 10 2.1 Past Missions to Mars...... 10 2.1.1 4 & 6 & 7 – Nasa Mars (1964-1971)...... 10 2.1.2 & 3 – Soviet Mars Orbiter and (1971) ...... 11 2.1.3 – NASA Mars Orbiter (1971)...... 11 2.1.4 Viking Project – NASA orbiters/landers to Mars (1975)...... 11 2.1.5 – NASA attempted mission to Mars (1992)...... 12 2.1.6 - NASA lander and rover to Mars (1996) ...... 12 2.1.7 Mars Global - NASA Mars orbiter (1996)...... 13 2.1.8 - NASA attempted orbiter to Mars (1998)...... 13 2.1.9 - NASA attempted lander to Mars (1999)...... 14 2.2 Current Missions ...... 14 2.2.1 - NASA Orbiter Mission to Mars (2001) ...... 14 2.2.2 - ESA Mars Orbiter and Lander (2003)...... 14 2.2.3 Mars Exploration Rovers - Two NASA Rovers to Mars (2003)...... 15 2.2.4 Mars Reconnaissance Orbiter – NASA Orbiter (2005)...... 16 2.3 Future Mission ...... 17 2.3.1 Mission...... 17 2.3.2 NASA (US) Missions ...... 17 2.3.3 Future Mission Summary...... 18 3 Aerial - Aerobots...... 19 3.1 Advantages and Applications of Aerobot...... 19 3.1.1 Atmospheric science...... 19 3.1.2 Payload delivery ...... 19 3.1.3 Sample site selection...... 20 3.1.4 Composition & Subsurface investigation ...... 20 3.1.5 Surface imaging & mapping ...... 20 3.1.6 Magnetic environment ...... 21 3.2 Aerobot Concepts...... 21 3.2.1 Lighter than Air ...... 21 3.2.2 Heavier than Air ...... 25 3.2.3 Summary: ...... 30 4 Surface Robots - Rovers...... 32 4.1 Rover Missions...... 32 4.1.1 Lunar Roving Vehicle (LRV) ...... 32 4.1.2 Russian Lunokhod Rover...... 33 4.1.3 Mars Pathfinder ...... 33 4.1.4 ...... 34

iii 4.2 Rover Classification ...... 35 4.2.1 Macro rover...... 35 4.2.2 Mini rover ...... 35 4.2.3 Micro rover...... 35 4.2.4 Nano rover...... 36 4.3 Conclusion: ...... 37 5 Aerobots and Rovers Collaboration...... 38 5.1 Mission Objectives and Science Rationale:...... 38 5.2 Background...... 39 5.2.1 Terrestrial application...... 39 5.2.2 Planetary exploration ...... 40 5.3 Cooperation Strategies...... 40 5.3.1 Imaging and Mapping:...... 41 5.3.2 Localisation and Navigation/Route planning:...... 41 5.3.3 Communication: ...... 42 5.3.4 Transportation/Delivery/Retrieval:...... 42 5.4 Vehicle Selections ...... 43 5.5 Proposed Cooperation Architecture ...... 48 6 System Design Overview ...... 51 6.1 Environmental Conditions ...... 51 6.2 Airship Evaluation...... 51 6.2.1 Airship sizing ...... 52 6.2.2 Power and Propulsion System...... 54 6.2.3 Envelope and Fins ...... 60 6.3 Results and Analysis...... 62 7 Payload Selection...... 66 7.1 Aerobot Payload...... 66 7.1.1 Mass Spectrometer ...... 67 7.1.2 Mini Thermal Emission Spectrometer...... 67 7.1.3 Subsurface Sounding ...... 67 7.1.4 ...... 68 7.1.5 Ultra-Violet Sensor...... 68 7.1.6 Panoramic Camera (Pancam)...... 69 7.1.7 Meteorology Package ...... 69 7.2 Rover Payload and Selection ...... 70 7.2.1 Microrover...... 71 7.2.2 Nanorovers ...... 73 7.3 Intermediate Platform...... 74 8 Conclusion and Future Work...... 76 REFERENCES ...... 78 APPENDIX ...... 83

iv TABLE OF FIGURES

Figure 2-1 [1] ...... 10 Figure 2-2 Spacecraft & Lander [1]...... 11 Figure 2-3 Viking Lander [1] ...... 12 Figure 2-5 Mars Exploration Rover [1] ...... 16 Figure 2-6 Mars Reconnaissance Orbiter [1] ...... 17 Figure 3-1 Mars Solar Balloon Mission Scenario (Courtesy of NASA) ...... 22 Figure 3-2 Zero-pressure Balloon (Courtesy of IKI) ...... 23 Figure 3-3 DARE Mars Balloon Platform [12]...... 24 Figure 3-4 Airship Explorer [22]...... 25 Figure 3-5 ARES Mars Airplane (Courtesy of NASA) ...... 28 Figure 3-6 Martian Autonomous Rotary-wing Vehicle [27] ...... 29 Figure 4-1 (Courtesy of NASA)...... 32 Figure 4-2 Lunokhod Rover [3]...... 33 Figure 4-3 Pathfinder Sojourner Rover (Courtesy of NASA)...... 34 Figure 4-4 Mars Exploration Rover (Courtesy of NASA)...... 34 Figure 4-5 Muses-C Nanorover (Courtesy of NASA)...... 36 Figure 5-1 Aerobot-Rover Cooperation Strategy ...... 42 Figure 5-2 Aerobot-rover schematic diagram ...... 49 Figure 5-3 Cooperation Scenario & Deployment Phases ...... 50 Figure 6-1 Airship Configuration ...... 52 Figure 6-2 Propulsion System Drive Train [45]...... 57 Figure 6-3 Hydrogen & solar airship comparison...... 62 Figure 6-4 Solar airship payload mass and finesse ratio's influence ...... 63 Figure 6-5 RPS airship payload mass and finesse ratio's influence ...... 64 Figure 7-1 Microrover illustration...... 73 Figure 7-2 Cliff descent by cooperative rovers (Courtesy of JPL)...... 74 Figure 7-3 Intermediate platform...... 75

v TABLE OF TABLES

Table 2-1 ESA Arora Mission Milestones [1][16] ...... 17 Table 3-1 Requirement & Limitation in each category [6]...... 20 Table 4-1 Rover Comparisons [38][39][40]...... 35 Table 4-2 Rover Classification [37][38][40]...... 37 Table 5-1 Trade-off Weightings ...... 44 Table 5-2 Mass & size scores...... 44 Table 5-3 Payload capability scores ...... 44 Table 5-4 Manoeuvrability/Controllability scores...... 45 Table 5-5 Mobility/Terrain coverage scores ...... 45 Table 5-6 Flight scores...... 45 Table 5-7 Degree of surface interaction scores ...... 46 Table 5-8 Adverse weather capability scores...... 46 Table 5-9 Complexity/Reliability scores ...... 46 Table 5-10 Trade-off Table ...... 47 Table 6-1 Known parameters definition ...... 53 Table 6-2 Required parameter definition ...... 53 Table 6-3 Component Efficiency for Electric Powered Aircraft [45][47]...... 58 Table 6-4 Assumed Fixed System Power Level...... 60 Table 6-5 Fixed Component + Payload Masses...... 61 Table 6-6 Airship sizing comparison and mass breakdown...... 65 Table 7-1 Payload mass distribution...... 66 Table 7-2 Aerial Instruments...... 66 Table 7-3 Mass Spectrometer Specifications [54]...... 67 Table 7-4 Mini Thermal Emission Spectrometer Specification [55]...... 67 Table 7-5 Subsurface Sounding Radar Specifications [56] [57] ...... 68 Table 7-6 Magnetometer Specifications [58]...... 68 Table 7-7 UV Sensor Specification [59]...... 69 Table 7-8 Panoramic Camera Specification [60] ...... 69 Table 7-9 Meteorology Sensors Specification [61] ...... 70 Table 7-10 Summary of aerobot’s science payload...... 70 Table 7-11 Contact Experiment Instruments [2][4][16][50][51][52][53]...... 71 Table 7-12 Sojourner science payload summery [2] ...... 71 Table 7-13 Rover payload instruments ...... 72 Table 7-14 Muses-C rover characteristics [42] [44]...... 73 Table 7-15 Intermediate platform...... 75 Table A Software Input...... 83 Table B Software Output ...... 83 Table C Software Test Results ...... 85

vi TABLE OF EQUATIONS

1 G  G Equation 3-1...... 26 M 3 E 1    Equation 3-2...... 26 M 60 E L C  Equation 3-3 ...... 26 l 1 V 2 S 2 C mG / mG lM  M E Equation 3-4 ...... 26 C 1 2 1 2 lE (  V S ) /(  V S ) 2 M M M 2 E E E

ClM GM / GE =>  2 2 Equation 3-5 ...... 26 ClE ( M /  E )(VM /VE )(S M / S E )

ClM 2  20(VE /VM ) (S E / S M ) Equation 3-6...... 26 ClE

P0  hTh Ph  Equation 6-1 ...... 51 0T0 2 3 4 5 6 Th =238.74 - 34.488h + 35.133 h - 15.96 h + 3.7315 h - 0.47352 h + 0.030962 h - 0.000817 h 7 Equation 6-2 [45] 51 5 2 7 3  h =0.014694 - 0.001145h + 4.663810 h - 9.773710 h Equation 6-3 [45] ... 51

Ph   H 2 RH 2Th (R in J/kg.K) Equation 6-4...... 53

PhVH 2  nRH 2Th (R in J/kg.mol) Equation 6-5...... 53

M H 2  RH 2universal Th VH 2  Equation 6-6...... 53 RAM H 2  Ph

M tot = VH 2  h - VH 2  H 2 Equation 6-7 ...... 53

M tot RAM H 2 Ph M H 2 = Equation 6-8 ...... 53 RH 2universalTh ( h   H 2 ) 1 D  V 2C V 3/2 Equation 6-9 [46] ...... 54 2 dv a 4 V   (l  d 2 )  l  3 3V  f 2 4/  Equation 6-10 ...... 54 as 3 l f  Equation 6-11...... 54 d 2 3 3 4 4 Cdv = 0.23175 - 0.15757f + 0.04744 f - 7.041210 f + 5.153410 f - 1.4835105 f 5 Equation 6-12 [46] 54

M sa  Sa  Equation 6-13...... 55

vii  P  5.0 S a I om e  sc Equation 6-14...... 55 energystored  fuelhours M  Equation 6-15...... 56 fuelcell fuelcelldensity  efficiency

 p  mcem g prop Equation 6-16 [47]...... 57 P M  mc (minimum 0.5kg) Equation 6-17[45] ...... 58 em 1291 P M  (minimum 0.1kg) Equation 6-18[45] ...... 58 mc 6233 P  M  em mc (minimum 0.3kg) Equation 6-19[45] ...... 58 g 3278 P M  (minimum 0.2kg) Equation 6-20[45] ...... 58 pc 1000 a  RT Equation 6-21 ...... 58 D 2 d p  2 2 Equation 6-22...... 59 ct ((aM ) V ) 5 3 V prop  .9 25739 10  d p Equation 6-23...... 59

M prop   prop nb 1(  Fb )V prop Equation 6-24 ...... 59 2 3 4 ct   .0 012122  .0 14577J  .0 1408J  .0 05374J  .0 0068444J Equation 6-25 59 2 3 4 c p   .0 012752  .0 094954J  .0 053694J  .0 017534J  .0 0007872J Equation 6-26 59 V J  Equation 6-27 ...... 59 (aM ) 2 V 2

ct J  prop  Equation 6-28...... 59 c p DV P  150 Equation 6-29...... 60  p d   arccos( ) Equation 6-30 [62] ...... 60 l  S  2 (d 2  l 2 ) Equation 6-31 [62]...... 61 envelope tan( )

M envelope   envelope  Senvelope Equation 6-32...... 61

M f  2.1 R faVas e Equation 6-33 ...... 61

M total  M as  1.1  (M h2  M envelope  M f  M sys  M p  M sa  M rfcs  M payload )  1.1 Equation 6-34 61

viii 1 Project Introduction

1.1 Introduction

Planet Mars is currently the focus of extraterrestrial explorations. Because of its resemblance to the earth, Mars raises the fundamental question of the existence of past, present or future life on its surface. Recent missions to the red have proved to be a great success in terms of technology and scientific return. The two American Mars Exploration Rovers have provided and continue to provide valuable data through their mission. At the same time, the orbiter Mars Express has been a remarkable European space technology achievement. It has returned a great amount of data and discoveries including the existence of a frozen sea near the equator. These missions’ success and discoveries enhance the interest in the planet and more than ever justify future Mars exploration missions.

Traditional missions have performed local, ground reconnaissance through rovers and landers, or global mapping by orbiting spacecrafts. A gap in between them has brought up the aerobot concept for Mars. However, all the robots and missions considered so far work solitarily. It may be advantageous to combine the heterogeneous platforms for studies of Mars. Aerobots offer a longer traverse and greater surface coverage capability than rovers, while rovers have better capabilities in surface contact experiments than aerobots. The idea of aerobots and rovers collaborative exploration may bring a better way to investigate Mars in the future. In order to capture the potential benefits of multiple robot explorations, new system architectures need to be studied that join the unique capabilities of each asset of the heterogeneous robotic team.

1.2 Objectives

The aim of this thesis is to investigate collaborative strategies for exploring a given piece of Martian terrain.

This can be accomplished by firstly investigating coverage and capabilities of different types of aerobots and rovers, then evaluating and comparing possible systems leading to a small set of preferred configurations. The aim is to design a system comprising several different elements – aerobots and rovers.

9 2 Mars Exploration

Mars is the fourth planet from the in our . Its surface condition and the evidence of water existence in the past make it the most hospitable planet for life other than Earth. Mars is widely considered one of the most scientifically interesting planet and the most likely planet for human colonization. It has been the target of exploration since the 1960’s. Several missions have been sent, most of them are US programmes. This chapter reviews the past, current and planned future missions.

2.1 Past Missions to Mars

The missions reviewed in this section have now ceased operation.

2.1.1 Mariner 4 & 6 & 7 – Nasa (1964-1971)

The first successful flyby mission was Mariner-4 (Figure 2-1) which is the fourth in the series of US spacecraft used for planetary exploration. It returns the first images of the [1]. It was designed to conduct close-up scientific observations of Mars and send back to Earth. Mariner 4 was launched in November 28, 1964, of a total dry mass 260.68kg. The scientific payloads were a TV camera, magnetometer, dust detector, telescope, trapped radiation detector, solar probe, and ionisation chamber/. The mission was designed for eight months, but it operated for about over three years in .

Figure 2-1 Mariner 4 Spacecraft [1]

Following this success, flyby missions Mariner 6 and Mariner 7 were launched. The instruments carried were wide and narrow angle TV cameras, infrared spectroscope, infrared radiometer and an ultraviolet spectroscope, intended to analyse Martian atmosphere and surface at the equator and polar region. Mars mass, radius and surface

10 pressure were estimated from this mission, and the polar cap was identified of being composed of carbon dioxide primarily.

Flyby missions were designed to observe the planet through multi-spectral imagery, radiometry etc. from interplanetary trajectory. Such missions can provide only low resolution due to the large distances to the target planet, for example, the closest approach Mariner 4 reached was about 10,000 km above the surface [1].

2.1.2 Mars 2 & 3 – Soviet Mars Orbiter and Lander (1971)

Mars 2 & 3 missions consisted of identical spacecraft each with an orbiter module and a lander module. Both missions have successfully reached Mars, decent modules have successfully entered the Martian atmosphere and sent back a total of 60 pictures, but due to some unknown reasons both landers crashed [1]. Orbiter crafts continued to return data for some time.

Figure 2-2 Mars 3 Spacecraft & Lander [1]

2.1.3 Mariner 9 – NASA Mars Orbiter (1971)

Mariner 9 was the first spacecraft designed for orbiting Mars. It launched on 30 May, 1971. Mariner 9 was after the failure of which was destroyed during the launch, therefore the mission combined the objectives of both crafts which involved surveying planet surface and studying temporal changes in the atmosphere and on the surface. Mariner 9 gathered precious data on Mars atmospheric composition, temperature, density, pressure, surface composition, gravity and topography of Mars which revealed a detailed view of volcanoes, and the polar caps, along with the fist close-up pictures of Mars’ natural and [1].

2.1.4 Viking Project – NASA orbiters/landers to Mars (1975)

Viking one and Viking two were launched on 20th of August and 9th of September, 1975 respectively [1]. The mission included the first landers to return data from the surface for a significant period of time (Viking one lander lasted 7 years). The role of the orbiters in the mission was to transport the lander to the orbit, locate suitable landing

11 sites, and then relay the communication for landers as well as doing their own investigations of imaging, infrared thermal mapping and measuring atmospheric water vapour. Once the landing site selection was done, the landers were separated from the orbiters and soft landed at the destination. The landers were equipped with instruments allowing studying Martian chemical composition, biology, meteorology, seismology, magnetic properties, appearance and physical properties of the atmosphere and surface.

Viking mission gave the most complete view of Mars to date. Features such as canyons, craters areas, volcanoes, lava plains and evidence of surface water are apparent in the orbiter images. The landing sites and were characterised as iron-rich clay, with temperatures ranging from 150K to 250K, with a daily variation of 35 to 50K [1]. Seasonal dust , pressure changes were observed and biology experiment showed no evidence of life at either site.

Figure 2-3 Viking Lander [1]

2.1.5 Mars Observer – NASA attempted mission to Mars (1992)

Mars observer mission ended when the spacecraft was in the vicinity of Mars. It was conjectured that a fuel explosion led to the failure when the spacecraft began its manoeuvring sequence for Martian orbit insertion. The mission objective was to explore the Martian environment, study the geology, geophysics and [1].

2.1.6 Mars Pathfinder - NASA lander and rover to Mars (1996)

Mars Pathfinder mission arrived on Mars on 4 July, 1997; this opened a whole new chapter of Mars exploration. Pathfinder delivered a lander and a micro-rover “Sojourner”. The mission’s primary goal was to demonstrate low-cost exploration of the Martian surface. The science goals involved long-range and close-up imaging of the surface, developing an understanding of Mars atmosphere entry science, rock and soil composition and meteorology, and the characterization of the environment.

12 The lander used a series of parachutes and rocket burns to slow down its descent, and an airbag system deployed to cushion to the impact. The airbag technology helps to reduce the propellant mass for touchdown compared to a traditional landing method, it allowed the lander to hit the surface at 18m/s. The landing site selected was which is an ancient rocky flood plain in the northern hemisphere containing a wide variety of rocks deposited during a catastrophic flood [1].

The rover "Sojourner" is a six-wheel rocker-bogie vehicle (mass 10.6kg), controlled by an Earth-based operator, who uses images from both lander and rover systems. Note that the time delay was about 10 minutes, therefore requiring some degree of autonomous control by the rover. During its 83 sols operation, it analyzed the chemical properties of 16 locations, and sent back more than 500 photographs to Earth [1].

Figure 2-4 Martian Sojourner Rover [1]

2.1.7 - NASA Mars orbiter (1996)

The Mars Global Surveyor was a polar orbiter. It began its prime mapping mission in March 1999, and lost contact with NASA in November 2006. Mars Global Surveyor mission’s other objectives were to study Mars surface topography, atmosphere and interior. In order to complete its objective, the Mars Global Surveyor carried a camera, thermal emission spectrometer, laser altimeter and a magnetometer [1].

2.1.8 Mars Climate Orbiter - NASA attempted orbiter to Mars (1998)

As the name suggest, Mars Climate Orbiter was a designed for weather, climate study on Mars, and it also served as a data relay satellite for Mars Polar Lander. The most important two instruments on Mars Climate Orbiter were the Mars Climate Orbiter Colour Imager (MARCI) which acquires atmospheric weather images and high resolution surface images; Pressure Modulated Infrared Radiometer (PMIRR) which allows measurement of the atmospheric temperature, water vapour abundance, and dust concentration [1].

The Mars Climate Orbiter was destroyed when a navigation error (an unnoticed error in conversion between English and metric units) caused it to miss its target altitude, entering the Martian atmosphere during the manoeuvre [1].

13 2.1.9 Mars Polar Lander - NASA attempted lander to Mars (1999)

Mars Polar Lander was an ambitious attempt to land the craft at the edge of the south polar cap. The mission was to dig for water ice with a , study surface geology and atmosphere at site. Deep Space two were the two small probes onboard the lander, they failed during descent [1].

2.2 Current Missions

The missions described in this section are due to be operating at the time of writing.

2.2.1 2001 Mars Odyssey - NASA Orbiter Mission to Mars (2001)

Mars Odyssey has a similar design of the Mars Climate Orbiter and pursues the goals of the failed Mars Observer mission; it is also designed as a communication relay for future missions. The Orbiter has a period of 3h15min and an inclination of 93 2.  [4]. Scientific payloads are star cameras, Mars Radiation Environment Experiment (MARIE), Thermal Emission Imaging System (THEMIS), and Gamma-Ray Spectrometer (GRS). The instruments will conduct a mineralogical and chemical survey of the planet.

2.2.2 Mars Express - ESA Mars Orbiter and Lander (2003)

Mars Express was launched by European Space Agency (ESA). It has successfully entered the elliptical Mars orbit with a period of 7.5 hours and an inclination of86 . -2 was the lander. It was deployed several days prior to the arrival at Mars, but never sent a signal since entering the atmosphere. It was assumed the impact of the surface exceeded the braking capability of the design.

On Mars Express, the imaging of the surface is performed by a high resolution camera. This instrument enables three dimensional data to be derived from overlap. It has imaged the entire surface of Mars with a spatial resolution of 10 metres, and areas of special interest with a spatial resolution of 2 metres. The visible and infrared mineralogical mapping spectrometer maps mineral composition of the Martian surface with 100 metre resolution, and determines the composition and planetary circulation of the atmosphere. A sub-surface sounding radar altimeter examines the sub- surface to a depth of several kilometres.

Beagle-2 was a small lander intended to investigate the presence of water, determine the existence of carbonate minerals and organic residues, record the complexity and structure of organic material, and measure isotopic fractionation between organic and inorganic phases. The instruments used to fulfil the experiments were a mass spectrometer, Mossbauer spectrometer, microscope, imager, and a small package

14 containing all environmental sensors. Beagle-2 was a secondary payload of Mars Express and had very limited resources [16].

2.2.3 Mars Exploration Rovers - Two NASA Rovers to Mars (2003)

Mars Exploration Rovers (MER) are two identical rovers named “” and “”. They were sent successively in 2003 to explore the Martian surface and geology. The landing sites were selected based on the environmental conditions that may once have been wet and favourable for life. For Spirit, Crater was chosen, a basin that appears to have held a lake; For Opportunity, a broad plain named was chosen based on the mineral-mapping from Mars Global Surveyor that had identified a big size exposure of a mineral called gray hematite, a evidence for a possibly watery past, because gray hematite usually forms in the presence of liquid water on Earth.

MER mass is 174kg, more than 17 times as heavy as Sojourner rover. Spirit and Opportunity have identical suites of five scientific instruments and a (RAT) to expose fresh surfaces of rock [2]:

 A Panoramic Camera - Providing the geologic context and helping select rock and soil targets for further study

 A Miniature Thermal Emission Spectrometer – Identifying minerals at the site, particularly searching for distinctive minerals formed by water related process such as sedimentary cementation, evaporation or precipitation.

 A Moessbauer Spectrometer – Identifying iron-bearing minerals, it is placed against rock targets by the rover arm.

 An Alpha Particle X-Ray Spectrometer (APX) – Determining the composition of rocks, learning the elemental ingredients in rocks and soils.

 A Microscopic Imager – Looking at fine-scale features, which could provide essential clues of the forming of rocks and soils.

15 Figure 2-5 Mars Exploration Rover [1]

Now both rovers have worked over 3 years, twelve times longer than its three-month mission objective, traversed over 10 kilometres each [2], conducted experiments and analysis of different types, and transmitted an incredible volume of data back to Earth.

2.2.4 Mars Reconnaissance Orbiter – NASA Orbiter (2005)

Mars Reconnaissance Orbiter is designed to orbit Mars and conduct reconnaissance and exploration for one full Martian year with six scientific instruments, including a high resolution imager. The science objective is to look for evidence of past or present water, study weather and climate and identify landing sites for future missions.

After its one Martian year primary mission, MRO will serve as a highly capable relay satellite for future missions, allowing a maximum data rate of 6 megabits/sec [1].

16 Figure 2-6 Mars Reconnaissance Orbiter [1]

2.3 Future Mission

Missions in this section are planned at the time of writing, and subject to changes.

2.3.1 European Space Agency Mission

ESA has adopted a long term plan to explore Mars named “Aurora”, the intention is to increase ESA’s capability gradually through a series of progressive development and eventually enable a manned mission to Mars [1].

Mission Scheduled Launch Year Earth Re-entry Vehicle/Capsule Demonstrator 2007 ExoMars 2013 Human Mission Technologies Demonstrator 2014 Mars Sample Return mission with NASA 2016 Technology Demonstrator 2018 Human Lunar Mission 2024 Automatic Mars Mission 2026 First 2030/2033 Table 2-1 ESA Arora Mission Milestones [1][16]

2.3.2 NASA (US) Missions

NASA Mars exploration plan in the future includes science orbiter, rover and lander, and sample return missions. Mars Lander mission is scheduled for launch at the end of 2007. The lander is planned to land on the North Pole of Mars and investigate the evidence of past liquid water and demonstrate the energy sources available to sustain life [1].

17 is another proposed mission expected to be launched in 2009. This mission involves a long distance rover which contains a mobile laboratory. The scientific objectives are to assess biological potential of target areas and characterize geology, geochemistry and surface radiation with regard to habitability [1].

2.3.3 Future Mission Summary

Many proposed missions to Mars are currently under investigation. Only a small portion of those will be carried out in a foreseeable future. None of the proposed vehicles has the capability to carry out a number of surface samples over a significant area of the planet, or to accurately visit a specific site chosen prior to the mission and inspect it. The mobility limitation constrains the mission value.

Aerial Vehicles (Aerobots) are therefore proposed for their high mobility and their capability of imaging/atmospheric sampling over large areas, and the cooperation potential with surface robots.

18 3 Aerial Robots - Aerobots

Aerobots have significant advantages for exploring Mars compared with orbiters, landers and rovers. An aerobot offers a long traverse and extensive surface coverage capability while also being able to provide high resolution imagery.

However, there are many limitations for flight in the rarefied Martian atmosphere. A thin atmosphere means for Lighter-Than-Air (LTA) aerobots, in order to create the same buoyancy as on Earth, a much bigger volume is required; for Heavier-Than-Air (HTA) aerobots, a much higher forward velocity and a bigger wing area is required to keep in flight. An atmosphere comprising ninety five percents of carbon dioxide precludes the application of conventional air-breathing engine on Mars. In addition, the lower speed of sound places limitations on the propeller tip speed.

This chapter outlines the capabilities of aerobots, the applications they could be applied to, and different aerobot concepts.

3.1 Advantages and Applications of Aerobot

Aerobots have their special strength, just as aircrafts did on Earth. There are many jobs aerobots can fulfil on Mars. This section examines some key applications that make aerobots suitable for Mars exploration:

3.1.1 Atmospheric science

A better understanding of the Martian atmosphere and weather is important for future missions, the lack of such knowledge could easily jeopardise a mission. For instance, during the Mariner mission, a dust was detected that lasted three month and engulfed the entire planet. Aerobots give the ideal platform for atmospheric in-situ measurement, and it would provide high quality data to predict the weather, modify and validate the Mars Climate / Circulation Model. This is important for lander and sampling return missions, as they require precision on landings and lift-offs.

3.1.2 Payload delivery

Mars has a highly varying terrain. The bumpy surface brings forward a lot of challenges to the rover and lander missions. Many sites of scientific interest are unreachable by ground systems, such as heavily uneven areas, mountains, canyons, volcanoes. One example is what’s happening on Mars now, the Opportunity rover roamed on the rim of the Crate for months but couldn’t find a smooth entry in. While, aerobots would have the ability to reach this type of site and deliver payload instruments.

19 3.1.3 Sample site selection

Aerobots with their special perspective can provide critical information for sample sites selection. Aerobots would gather the high resolution topographic, mineralogical and morphological data of the Martian surface, this in turn can be used to identify scientifically interesting sites for future exploration. Observation sites are divided in three general categories, Regional, Local and Near Field.

Table 3-1 summarises the resolution requirements of mineralogical and morphological scans in each category [6].

Applicable Spatial Spatial Spatial Observational Range of Resolution Resolution Platform Limitation of the Scale interest Mineralogy Morphology Otherwise Platform require very large and 100km - sophisticated Regional 1000km 1m-5m 10cm Obiter instruments require very large and sophisticated Obiter instruments 100m - Lander limited time and Local 5km 1m-5m 10cm descent coverage Near Field 1m - 100m 1cm-1m 1mm-10cm Rover limited perspective Table 3-1 Requirement & Limitation in each category [6]

Near field data is useful for surface robots selecting science objects in view. Local data is required to guide surface robots to the sampling location. Regional scale observations are needed for selecting sites of interest for landing.

3.1.4 Composition & Subsurface investigation

The information of rock composition on and sub-surface is important for selecting in situ sample sites. Aerobots could be utilized to investigate rock and subsurface structure, identify water, ice, mineral and fossils using spectroscopy and radar sounding techniques. They can achieve a better spatial and depth resolution than any other means.

3.1.5 Surface imaging & mapping

Aerobots have the capability to provide high resolution imaging and cover a wide range of area, the collected information can be used to identify sites of interest and guide rover excursions. Surface mapping could provide invaluable information that could aid in understanding the history of Mars, the data can be used to study seismology, volcanology, hydrology and tectonic processes.

20 3.1.6 Magnetic environment

Mars does not have an intrinsic , however observations have shown parts of the Mars’ crust are magnetised, the study of paleomagnetics could unveil the history of Mars. Thanks to its altitude and spatial resolution, an aerobot will be able to detect stronger magnetic than a spacecraft and covers wider area than a rover, and thus provides an ideal platform to investigate the magnetosphere.

3.2 Aerobot Concepts

Aerobots that have been considered for planetary exploration under different concepts include , , airplanes, gliders and rotorcrafts. They can be divided into two basic categories by the way they keep airborne:

Lighter-Than-Air(LTA) Aerobot – Staying aloft by static lift generated by lighter gas Heavier-Than-Air(HTA) Aerobot – Staying in flight by dynamic lift generated by aerofoil

Aerobots have very unique capabilities and wide areas of application. They bridge the world of orbiters and rovers, and many proposals for future exploration missions have included the use of aerobots.

3.2.1 Lighter than Air

LTA – balloons, airships

3.2.1.1 Balloons

There’re three types of balloons: Montgolfiers, Zero-pressure balloons and Super- pressure balloons.

3.2.1.1.1 Montgolfier

This type of balloon uses heated atmospheric air as the lift gas. It is also well known as the hot-air balloon. To keep the internal air less dense, the air has to be regularly heated. The traditional way is using a burner, but some balloons utilize the solar flux and infrared flux to transfer heat inside the envelope.

A solar Montgolfier absorbs solar energy as the heat source, it rises to the ceiling altitude during the time of daylight, cools down and descends at night. A guide-rope can be used and deposited on the ground at night to allow the balloon maintain its buoyancy. The guide-rope can be instrumented for surface measurement or soft landing payload.

21 A Mars solar Montgolfier balloon has been proposed to explore Mars Polar Region by JPL [7]. During Martian Polar summer, the continuous period of daylight could last for about an Earth year, thus solar power becomes a feasible option for a long period mission on Mars. The balloon’s altitude control is accomplished by keeping and venting out the internal hot carbon dioxide. When the vent is closed, hot air is trapped and the balloon rises, when open, hot air escapes and the balloon descends. The vent opening and closing are radio controlled. Calculations in the proposed system have shown that an envelope of 9.25kg with a diameter of 15 metres, using 13g/ m 2 areal density material, could carry a 2kg gondola; in comparison, using a less dense material - 9g/ m 2 Mylar, carrying the same weight, the balloon is reduced to 4.39kg and 12.46 metre diameter.

Figure 3-1 Mars Solar Balloon Mission Scenario (Courtesy of NASA) Solar Montgolfiers have a big advantage in that small gas leaks will not impair the mission performance, since leaking air will be quickly replaced and heated by the sun.

3.2.1.1.2 Zero-pressure

Zero-pressure balloon envelopes are made of thin plastic material, usually polyethylene [18]. Hydrogen and Helium are the common options for lift gas. On Earth, Helium is mostly preferred over hydrogen due to safety reasons. The balloon envelope is partially inflated and sealed before launch. As the balloon rises up, the envelope expends to maintain the pressure balance till fully inflated, where the highest altitude is reached. To maintain the same altitude at night, ballast must be dropped to compensate for the cooler gas and smaller volume. Because it is zero pressure, there’s less stress on the envelope than other types of balloon, which makes zero-pressure balloon a very good tool for very high altitude missions.

There have been extensive studies of this type of balloon on Earth. NASA has flown balloons of payload capacity of four tons, to altitude of 40km with volumes of 2,000,000 m3 . The Long Duration Balloon flight set a record of 31 days in 2002, flying a zero-pressure balloon over Antarctica at an altitude of 38km. [17]

Advanced studies on zero-pressure balloons were investigated for Mars exploration. A Mars 96 aerostat was designed and planned to fly in the Mars-96 mission. The concept

22 featured a cylindrical Mylar envelope sealed with helium gas and equipped with a guide-rope. The balloon was planned to fly during the day and sink at night as the volume decreases. Part of the guide-rope would lie on the ground at night and carry out in situ analysis. The totally mass of the balloon was 65kg, with a 15kg gondola and a 13.5kg instrumented guide-rope [13]. Unfortunately, this Russian programme was cancelled at the end due to some financial difficulties.

Figure 3-2 Mars 96 Zero-pressure Balloon (Courtesy of IKI)

3.2.1.1.3 Superpressure

Another type of balloon is the so called superpressure balloon, it is made of a tough and inelastic envelope (usually Mylar) and filled with the high pressure lift gas and then sealed [18]. Superpressure balloons maintain a constant volume and fly at a constant density altitude. This type of balloon can stay aloft for weeks or months and is great for long atmospheric studies.

The first and the only planetary balloon mission so far – VEGA, used this type of balloon for the exploration of . It was a joint Russian French mission in 1985 [11]. Two balloons performed for two days measuring of temperature, cloud density and wind speed on Venus until the batteries ran out. The design of the aerobot was a spherical superpressure balloon filled with helium gas, with a small diameter of 4 metres (due to the very dense atmosphere on Venus), the payload mass was 7 kg, and the total mass was 21 kg. Although the mission duration was short, it proved the feasibility of the planetary aerobot concept.

With the improved material technology, superpressure balloons have become more and more favoured in Mars balloon studies, as they would drift in an almost constant altitude. Study [17] has simulated that at a height of 5km (high enough to avoid most topography), a superpressure balloon could circumnavigate Mars in 7 days. Different from other balloons, superpressure balloon will float at a higher altitude at night,

23 because at night the atmosphere cools down and atmospheric density increases. However, seasonal changes could result in a 2km’s flight altitude variation [17].

A unique design “DARE” was proposed for Mars exploration using a guided superpressure balloon [12]. A special trajectory control system was introduced for the balloon. The idea was to use a wing hanging below the balloon through a long tether, due to the difference in winds between the altitudes, the wing would create a sideways lifting force and drag the balloon to avoid topographical obstacles [12]. Mars General Circulation Model was utilized to simulate the controlled and uncontrolled trajectories. Results showed over a 5-day period, the trajectory control system altered the balloon trajectory significantly. However, over a longer period, the controlled and free drifting trajectories did not present any significant difference. Note that the DARE control system can not achieve manoeuvring the balloon over a predetermined site.

Figure 3-3 DARE Mars Balloon Platform [12]

3.2.1.2 Airships(Blimps)

“Airship is fundamentally a balloon with a means of propulsion.[15]” In order to reduce air resistance and enhance controllability, the balloon shape has to be modified to streamline form. Airships were widely used as cargo carriers for transportations in early days before their capabilities were surpassed by terrestrial airplanes.

There are a few feasibility investigations on Martian airships were found during the research. Girerd [19] presented a 199kg propeller driven hydrogen Martian airship, powered by a reciprocating engine. The solar power concept was ruled out due to the difficulty of packing solar cells and the risk of damage during deployment and inflation. Gundlash [20] presented an innovative hybrid airship design, which uses

24 largely dynamic lift with vectored thrust supplied by photovoltaic power. Only 4% of the lift is generated from the atmospheric buoyancy. The airship has to keep a 24m/s speed to stay in flight. At night it must land using its vectored propellers because there’s no power storage onboard. It could carry 10kg of payload with a total mass of 50kg.

Besides Martian airships, a Titan airship was developed by Jeffery L. Hall, et al [22]. It was a propeller driven superpressure airship using helium as the lift gas. Due to the fact that Titan is remote from the Sun, nuclear power is considered the only practical power source. Stirling-cycle Radioisotope Thermoelectric Generator (RTG) was introduced, it would produce continuous power, and the waste heat from RTG provides the thermal energy for other subsystems. The airship flies at 10km altitude, and it is approximately 10 metres long, 2.5 metres in diameter, 100kg total mass with 25kg science payload. The high density of the atmosphere on titan gave the helium balloon significant buoyancy with very little mass. The steering and manoeuvring are achieved by three motored propellers. For vertical control, a ballonet is used, it is a secondary internal gas bag filled with ambient gas by a blower. The change of density inside the envelope would cause the airship to move up and down. On the bottom of the airship an inflatable wheel is attached which could cushion the landing on rugged surfaces and provide flotation on the anticipated liquid methane ocean.

Figure 3-4 Titan Airship Explorer [22]

3.2.2 Heavier than Air

HTA – Fixed wing aerobots: Airplanes, Gliders (unpowered airplane) Moving wing aerobots: Rotorcrafts, VTOL aircraft They generate lift by the high relative velocity of wings through air.

3.2.2.1 Fixed Wing Aerobots –Airplanes & Gliders

Mars airplane is one of the most studied aerobot platform. No matter which planet it is, aerodynamics governs the design of all aircrafts. For an aircraft in flight, it has to balance its weight with lift, drag with thrust.

25 The advantage for wings on Mars is that the gravitational attraction on the Mars surface is one third of that on Earth. However, the disadvantage is that the atmospheric density is 1/60 that of Earth. 1 G  G Equation 3-1 M 3 E 1    Equation 3-2 M 60 E

We can compare the required lift coefficient for airplanes designed to operate on Earth and Mars. Subscripts M and E refer to the values for Mars and Earth, respectively.

Lift Coefficient L C  Equation 3-3 l 1 V 2 S 2 where L - lift force  - atmospheric density V - true airspeed S - wing area

C mG / mG lM  M E Equation 3-4 C 1 2 1 2 lE (  V S ) /(  V S ) 2 M M M 2 E E E C G / G => lM M E Equation 3-5  2 2 ClE ( M /  E )(VM /VE )(S M / S E )

Inserting Equation 3-1 and Equation 3-2 into Equation 3-3, it is obtained the following relation

ClM 2  20(VE /VM ) (S E / S M ) Equation 3-6 ClE

Hence, all things being equal, Equation 3-6 indicates that a Martian aircraft wing must have a lift coefficient 20 times greater than a corresponding aircraft on Earth. If it is allowed to double the wing area and double the velocity, ClM still need to be 2.5 greater thanClE . If increase in Cl is not possible, this deficit must be made up for by increasing wing area and flight speed. However, the dilemma is that: - The airplane size is constrained by the launcher capacity and the difficulty of packing large wings. - The airplane’s speed is limited because the speed of sound on Mars is lower than on Earth (roughly 2/3), hence gives the airplane a higher cruising Mach number.

26 Several Martian airplanes have been investigated and range significantly in scale and objectives. One concept was the NASA Mini Sniffer. It was a high altitude UAV demonstrator for flight on Mars using a non-airbreathing hydrazine engine. It has demonstrated both flight at ultra high altitude on Earth and the feasibility of flight in the Martian atmosphere.

The Canyon Flyer was a NASA proposed micro-mission to fly an aeroplane on Mars. It is a propeller driven subsonic airplane designed to conduct scientific exploration on Mars. This Canyon Flyer [23] has a folded 2.2 metre wingspan which will be deployed during the . Two concepts for the propulsion system were considered, one being a battery powered , and the other a hydrazine internal combustion engine. The mission concept was to fit a Mars mission within Ariane 5 launcher which had a spare capacity of 200kg into geosynchronous transfer orbit. It would fly at 500m altitude above Valles Marineris with 8kg payload, taking high- resolution images at a range of less than 1km. It was desired to transmit the acquired data for the full duration of the mission via micro mission spacecraft, but the design was constrained by the requirement which leaves the mission endurance for image relay only 20 minutes.

Solar powered airplane concepts have also been proposed for longer endurance Martian flight. On Earth, the AeroVironment’s Helio solar powered UAV had flown a record altitude of 29km where the atmosphere condition is similar to Mars.

The investigation done by Colozza [24] presented a comparison of preliminary design for Mars airplanes with solar power and RTG power sources to carry a 100kg payload. This report concluded that both power sources result in a feasible long endurance Mars airplane. The RTG powered airplane ranges in total mass from 322 to 520kg; the solar powered airplane dependent on efficiencies range from 438kg with 47m wingspan to 1149kg with a 128m wingspan. In addition, the solar airplane is limited to low latitude exploration.

Another airplane design was investigated by the NASA ARES team [25]. The mission will return scientific data about the planet’s atmosphere, surface and interior, using the payload instruments carried onboard: two , a Neutron Spectrometer, a Video Camera, a Context Camera, and a Mass Spectrometer. The airplane has a wing area of 7 square metres, total wet mass 175kg of which 48kg is propellant. It will survey a 500km pre-planned route within 60minutes, at a cruise Mach number of 0.68 and an altitude of 2km. The plane is powered by Li  SO2 batteries and driven by a bi- propellant rocket propulsion system with MMH fuel and MON-3 oxidizer. Science and engineering data are continuously transmitted to the carrier spacecraft. Navigation uses inputs from IMU (inertial measurement unit), barometer, air data system (airspeed, angle-of-attack, and angle-of-sideslip) and a radar altimeter.

27 Figure 3-5 ARES Mars Airplane (Courtesy of NASA)

3.2.2.2 Moving Wing Aerobots – Rotorcrafts, VTOL aircraft

Rotorcraft is a category of flying machines that include helicopters, autogyros, gyrodynes and tiltrotors. The common feature of them is that they generate lift by the use of wings that revolve [4].

On Earth, the biggest terrestrial rotorcrafts have payload capacity on the order of 20 tons, or 120 passengers for commercial variant [26]. However, for Mars rotorcrafts, restricted by the rotor limitation and enormous power consumption, most work has been focused on under 50kg class craft design.

For rotorcrafts work on Mars, large blade surface area and rotor radii are required to provide adequate lift in the Mars’ thin atmosphere. To fulfil this, light-weighted blade construction is essential. Calibration by Larry A. Yong, et al [26] implies that Mars rotor blades can only weigh roughly 10-14% of their equivalent sized terrestrial rotorcraft blades.

The MARV (Martian Autonomous Rotary-wing Vehicle) design by University of [27] was a twin bladed, electrically powered coaxial rotorcraft. It was designed to fly on Mars for 30 minutes at altitudes up to 100 metres, capable of hover and soft landing. The mission scenario is that after the autonomous deployment, the rotorcraft will take off, hover above the ground, and then accelerate to its horizontal maximum speed of 11.5m/s in 60s; at this maximum speed, it will climb to its 100m maximum altitude, and starts doing scientific investigations; during the 30 minutes mission lifetime, MARV will cover a range of 25km distance; finally, the rotorcraft will descent, decelerate, and then soft land on the ground.

28 Figure 3-6 Martian Autonomous Rotary-wing Vehicle [27]

The rotorcraft has a total mass of 50kg including 10.8kg of payload. To be the lowest blade mass possible, Mylar was used as the main component of the skin structure. PEM ( exchange membrane) fuel cells were selected for the power system due to their minimum mass. Rotorcraft’s propulsion system is the most power and energy demanding system on any Martian aerobot. During the hovering phase, PEM will have to deliver a maximum power of 6.43kW, and on average 4.6kW.

Other than the aerobot concepts mentioned above, VTOL aircraft concept attracted attentions to the Mars aerobot missions. A study was found on Martian VTOL aerobot. A tilt-wing VTOL aircraft concept was developed by S.A. Ali [9]. This is a novel aircraft concept in between of a helicopter and an airplane. VTOL aircraft combines the vertical lift capability of a helicopter with the speed and range of an airplane. It uses proprotors, which can be angled to vertical or horizontal for lift and propulsion. The power system being used in his study was based on an assumed high performance regenerative fuel cell which could provide 10kW of power and have an energy density of 800Wh/kg. The aircraft has a gross take-off mass of 70kg where 20kg are allocated to fuel cells that will provide a total energy of 16,000Wh. With in this energy constraint, the tilted-wing aircraft could cruise at an altitude of 5km with a maximum forward velocity of 110m/s and cover a range of 808km in 2 hours and 23 minutes duration.

29 3.2.3 Summary:

Aerobots have presented a brand new stage of Mars exploration. Each of them has its strengths. The fundamental problem and the biggest obstacle however for both LTA and HTA aerobots is the Martian low density atmosphere. Aerobots of any kind must have a considerable volume or size to carry significant payloads.

The option of deployment and inflation during the descent in the atmosphere presents a big advantage, for it would eliminate the heavy and expensive rocket landing system, thus be a very effective way to improve the payload mass fraction for a mission.

Balloon is considered the simplest system. It is a cost effective method for Martian exploration. Balloons are lightweight relative to other concepts and would provide the longest durability, and they have the technology readiness for an immediate mission. However, the disadvantage is that balloon has no means to control its flight path. Balloon flight control on Earth relies on variation in wind at different altitudes. On Mars, control using wind direction would require a better understanding of the weather than we currently have. Essentially, balloons are carried by wind to wherever it blows. Robotic balloons are an attractive concept for free flying but not well suited for missions with predetermined targets.

Airship uses engine or motors to turn propellers which control the craft. By providing the envelope with control, airship can go where it is programmed to go rather than passively floating with the wind. Airships combine the long term airborne capability of balloons and the controllability of HTA vehicles. The stream lined shape is designed to reduce drag in the forward direction but also to maximise buoyancy. However, the disadvantage compared to the spherical balloon is that the airship is less efficient in terms of envelope material mass (around 15-20% greater) [16], and it requires a certain amount of power to push through the air.

The fabrication of a strong, lightweight, leak-proof material is a challenge for LTA aerobots. Gas leakage causes balloons/airships to descend and reduces their endurance. Modern materials can minimise the gas leakage and the necessary envelope mass, thus improve the useful payload as a fraction of total mass.

Airplane offers the ability to cover large areas in a short amount of time, as well as the ability to keep on a predetermined flight path with high reliability. It is a very attractive aerobot concept. Unfortunately, unless aircraft are capable of VTOL then take-off / landing is a major problem, the limited endurance (current concepts have a flight duration of a few hours) constrains airplane’s overall information collection capability.

Glider – an unpowered airplane depends heavily on wind and updraft patterns, which drastically limit its performance.

Rotorcraft presents some unique abilities, which are vertical take-off, hover, and landing. It exhibits potentials to incorporate aerial robot function and some surface robot functions such as conducting contact experiments and sample collection missions when it lands. However, due to the enormous power requirement which also leads to

30 short flight endurance and coverage, the ability to bring large valuable science return is very limited compared to other aerobots concepts. Rotorcraft would be more attractive if it can be refuelled by a lander or Martian base.

VTOL aerobot combines the advantages of both airplanes and helicopters, it is able to hover, vertical take-off and landing, fly longer time and cover wider areas than rotorcrafts. However, VTOL aircrafts is also a heavy and power hungry concept for Mars exploration.

31 4 Surface Robots - Rovers

Rovers have a long heritage in extraterrestrial missions. They are favoured throughout the history. Compared to aerobots, surface rovers have a high degree of technology maturity and a better position in providing in-situ experiment. They were one of most important means for Mars exploration and still will be in the future.

This chapter reviews the past rover missions and categorizes the different types of rovers.

4.1 Rover Missions

Eight successful rovers have been deployed on the and Mars over the last 37 years, including crewed vehicles as well as remote controlled and autonomous robots. The United States Lunar Roving Vehicle used on the Apollo missions and the ’s two Lunokhod lunar rovers explored the Moon. The Mars Pathfinder performed the first experiment, and the Mars Excursion Rovers Spirit and Opportunity continue to perform well.

4.1.1 Lunar Roving Vehicle (LRV)

“The Lunar Roving Vehicle (LRV), wasan designed to traverse the lunar surface, allowing the Apollo to extend the range of their surface extravehicular activities.” [14] Three LRVs were driven on the Moon on , 16, and 17. Each rover was used on three traverses, one per day over the three day course of each mission. The longest traverse was 20.1 km and the greatest range from the Lunar Module was 7.6 km, both on the mission.

The LRV had a mass of 210 kg and was designed to hold a payload of an additional 490 kg on the lunar surface. The frame was 3.1 metres long with a wheelbase of 2.3 metres and maximum height 1.15 metres [38][39].

Figure 4-1 Apollo Lunar Rover (Courtesy of NASA)

32 4.1.2 Russian Lunokhod Rover

Lunokhod 1 and 2 were two unmanned lunar rovers landed on the Moon by the Soviet Union in 1970 and 1973, respectively. The Lunokhod missions were primarily designed to explore the moon surface and return pictures.

Lunokhod 1 had a mass of 900 kg and was designed to operate for 90 days being remote controlled by a 5-person team from Earth. Lunokhod 1 explored the are for 11 months, travelling 11 km while relaying television pictures and scientific data.

Lunokhod 2 was an improved version of Lunokhod 1. It was faster and carried an additional television camera. The rover traversed 37 km in 8 weeks. It was 1.35m high, 1.7m long and 1.6m wide, with a mass of 840 kg. The 8 wheels each had an independent suspension, motor and brake. The rover had two speeds, 1km/hr and 2km/hr. Via the cameras mounted on the vehicle, a five-man team of controllers on Earth sent driving commands to the rover in real time. Power was supplied by batteries charged by a solar panel. A Polonium-210 isotopic heat source was used to keep the rover warm during the nights [3][38]

Figure 4-2 Lunokhod Rover [3]

4.1.3 Mars Pathfinder Sojourner

Mars Pathfinder was designed as a technology demonstration of delivering an instrumented lander and a free-ranging robotic rover to the surface of Mars. Pathfinder successfully accomplished its mission goal. In its four months of operation, the rover Sojourner, traversed a total distance of about 100 m. It had a mass of 10.6 kg (~1kg payload), and was 654830cm in size. The wheels and suspension used a rocker- bogie system. The rocker-bogie mobility system allowed the rover to traverse obstacles, providing exceptional stability on steeply sloped surface, and permitted each wheel to conform to the rugged terrain independently. Three motion sensors along Sojourner’s frame detect excessive tilt in order to stop the rover before it could tip over. Sojourner rover carried two finger-sized black and white cameras in front, a colour camera in the back, and an APXS for determining rocks and soil composition. It was capable of scaling a rock on Mars more than 20 cm. [1]

33 Figure 4-3 Pathfinder Sojourner Rover (Courtesy of NASA)

4.1.4 Mars Exploration Rover

The Mars Exploration Rovers (MER) - “Spirit” and “Opportunity” are two twin rovers landed on the opposite sides of Mars to search for rocks and soils that hold clues to past water activities on Mars. “The landing sites are at Gusev Crater, a possible formerlake in a giant , and Meridiani Planum, where mineral deposits suggest Mars had a wet past.” [2]Each MER has a mass of 174kg (~22kg payload), and speed of up to 40m/. To date, both rovers have worked for over 3 years, far exceeding the 3 month and 1km primary mission objective. The design of the suspension system for the wheels is similar to the “rocker-bogie” system on the Sojourner rover.

Figure 4-4 Mars Exploration Rover (Courtesy of NASA)

Future rover missions currently being planned include NASA's Mars Science Laboratory (launch in 2009) and the ESA's ExoMars mission (launch in 2013).

The family of extraterrestrial rovers has evolved in capability and application. Table 4-1 compares various performance parameters of the Sojoumer, MER, and Apollo rovers.

34 Rover Lunar Roving Comparisons Sojourner(micro) MER(mini) Vehicle(macro) Launch Year 1996 2003 1971 Rover Mass 10.6kg 174kg 700kg Payload Mass 1kg 22kg 500kg(incl. 2 astronauts) Rover Power 50Wh/sol 600Wh/sol 2800Wh Compute Power 0.25MIP 20MIP Human Control Scheme CARD+Behaviours CARD+Hazard Human Rover Life >90sols(actually) 90sols Hours Rover Range ~0.1km ~1km >25km Average Speed 0.3cm/s 5cm/s 6mph Mobility Low moderate high Table 4-1 Rover Comparisons [38][39][40]

4.2 Rover Classification

Although there’s no accepted uniform term for different type of rovers, they can be classified according to the vehicle’s size and capabilities.

4.2.1 Macro rover

Large Macro rovers are with a system mass around 1 tonne and an operating range of tens to hundreds of kilometres, typified by the Russian Lunokhod. This class now is considered obsolete since no mission can afford to transport such mass to the Moon or anymore [37], though this may change with possible human , but the cost is formidable.

4.2.2 Mini rover

Unmanned medium-sized rovers (Minirovers) are typified by the MER class rovers and are generally with a system mass around 200kg and a high degree of autonomy (operating range of 1-10s of km). This is the system of choice for regional exploration and scouting, able to conduct long term investigation, return in situ measurements and samples from the field.

4.2.3 Micro rover

Microrovers with a system mass of the order of 10kg, 100m range, tend to have moderate degree of autonomy in control and communications. These simple rovers offer great applications and potentials for regional explorations, with minimal own resource consumption and the small mass, they can be deployed and extend the reach of measurement instruments to the actual sites of interest (e.g. features in view of the

35 lander or aerobot), thus enhance the return of scientific missions. The Sojourner rover of NASA’s Mars Pathfinder mission is an example of this type.

4.2.4 Nano rover

Tiny nanorovers are of masses less than 5 kg with limited range of less than 10 m and limited capabilities.

Nanokhod, LAMalice and Muses-C are three examples found of this type. Nanokhod is a tethered, tracked rover designed to perform geochemistry investigation on the Martian surface. Thanks to the use of a tether through which power is provided, Nanokhod achieves a very high payload mass (1.1kg) to rover mass ratio of 50% [5]. While, LAMalice is a 30gram mass nanorover, 11cm6cm4cm in dimension, 1cm/s speed, developed by Swiss Federal Institute of Technology Lausanne. It has only the basic navigation sensors and a communication module which communicates with the lander. Using the on-board batteries it can achieve continuous operation of 20 hours.

Muses-C nanorover had been selected as a technology experiment on the Japanese sample-return mission MUSES-C, scheduled for launch in 2002, but was cancelled due to budget constraints. For this mission, the nanorover would have been dropped onto the surface of the asteroid to gather close-up imagery and spectral data, then relay the data through the MUSES-C spacecraft back to Earth. The Muses-C rover with a mass of 1.3 kg is 14146cm which is about 30% the size of the microrover Sojourner. It would have powered by solar energy and carry three scientific instruments: a multi-band imaging camera, a near-infrared point spectrometer, and an Alpha/X-ray spectrometer (AXS). The payload is capable of making texture, composition and morphology measurements of the surface layer at scales smaller than 1 cm, investigating lateral and vertical structure and to measure mechanical and thermal properties of the surface layer [42][44]. A variety of sensors were accommodated for navigation and mobility control, including Hall effect sensors for wheel odometry, and infrared proximity sensors for hazard detection [42]. The prototype consists of a four- wheel mobility chassis that is capable of rolling, climbing, or hopping around on the surface of the asteroid.

Figure 4-5 Muses-C Nanorover (Courtesy of NASA)

36 Recent advances in micro-technology and mobile robotics have made it feasible to create very small automated nanorovers. There has not been any successful nanorover mission so far, however nanorovers are under intensive development.

Table 4-2 generalises the different classes of rovers in terms of mass and mobility in comparison.

Planetary Rovers Classification Mass(kg) Mobility Macro Rovers ~ 1000 ~ 10 - 100km Mini Rovers ~ 30-250 ~ 1 - 10km Micro Rovers ~ 5-30 ~ 100m Nano Rovers <5 ~ 10m Table 4-2 Rover Classification [37][38][40]

4.3 Conclusion:

Although rovers are only capable of travelling relatively small distances and much of the Mars uneven surface is inaccessible to them, rovers have a long heritage and are clearly competent at conducting surface science and contact experiments. Whereas, aerobot offers the type of mobility and perspective that rovers are in lack of. It is thought therefore through collaboration, rovers and aerobots could both benefit from each other maximising the science return.

37 5 Aerobots and Rovers Collaboration

5.1 Mission Objectives and Science Rationale:

Different mission objectives and requirements bring different solutions. For this research, a mission is defined to investigate a given piece of terrain - Meridiani Planum.

Meridiani Planum is a plain of diameter of 1100 km, located 2 south of the equator. Evidence has shown that this plain hosts a rare occurrence of gray hematite - a mineral often formed in hot springs on earth which scientists believe is an indication of water past. And investigation of such a piece of terrain has a great potential to yield valuable science return.

Science Objectives:

 Understand and Characterise Mars Geology  Understand Mars Environment  Look for life trail on Mars / Biology

Mission Objectives / Requirements are as follows:

1. Multiple sampling

Multiple samples across a thousand kilometres of a location believed to have held water will enable the acquisition of valuable cross section of data. Craters are scientifically interesting for they enable access to the layers beneath the soil and a means of analyzing the geological history of Mars. There are 12 craters distributed within this area. Primary targets: Crater (the biggest crater in the plane with an overlapped small crater – Airy 0 in the centre, 41km diameter) and Victoria Crater (750m diameter), 380 km separated from each other. Mission should have the ability to achieve in-situ contact experiments at Airy crater and Victoria Crater. (NB. Victoria crater is known to be inaccessible by surface rover from its rim) Taking at least 2 sampling sites per crater: rim area + centre area

2. High resolution surface survey & mapping

High resolution images of the terrain surface. Look for potential landing/sampling sites and potential human base locations. Spatial resolution less than 1m.

3. Terrain compositional & subsurface investigation: penetration depth more than 5km (deeper than Mars orbiters have probed so far)

4. Atmospheric vertical profile

38 Rovers are mobility constrained, unable to explore multiple sites that are long distant from each other, and are unlikely to explore hazardous sites. On the other hand, aerobots have great mobility, but when it comes to surface contact experiments, they become restricted.

Future Mars exploration missions will incorporate multiple cooperative robot work together for maximum science return. The potential advantages of heterogeneous robots include the potential to share sensors and capabilities, to help one another simultaneously explore large areas in a more efficient way. For this mission, a solitary rover is not capable of exploring such a vast area neither is an aerobot feasible to achieve the required high degree of surface interaction. Aerobot and rover combined exploration is favoured.

5.2 Background

"The combination of the ground-level and aerial view is much more powerful than either alone," said Steve Squyres, the principal investigator for Opportunity and Spirit. "If you were a geologist driving up to the edge of a crater in your jeep, the first thing you would do would be to pick up the aerial photo you brought with you and use it to understand what you're seeing from ground level."[2]

The combination of rovers and orbiter team when Opportunity reached the cliff of Victoria crater was a tremendous example of how Mars missions are reinforced through higher level vehicle to expand lower level vehicle’s ability to explore and discover. Using the high-resolution camera on the Mars Reconnaissance Orbiter, it enabled the operators choose which way to send Opportunity around the rim, and guide the rover's exploration of Victoria.

Heterogeneous robots cooperation has been a hot topic lately, many examples of research have been found in several papers:

5.2.1 Terrestrial application

One study of aerial and ground vehicles cooperation for terrestrial application was carried out by the U.S Navy. It has shown that autonomous UAVs, unmanned ground sensors, and unmanned ground vehicles, can perform a complex mission autonomously. “Using the Navy’s low cost Silver Fox UAV (resembles a model aerobot with an 8 foot wingspan), the Navy conducted a completely autonomous intercept of a car travelling down a road. The car triggered ground sensors that formed as part of the network. The ground sensors, detecting the car, called two UAVs over to have a look. The sensors provided an approximate GPS location, and an onboard UAV visual system identified the object as a car. UAV then called in ground vehicles, which surrounded the car— all without human intervention”. [30]

39 5.2.2 Planetary exploration

Heterogeneous Robots Cooperation Studies

J. L. Hall et al [22] have described a concept of a sensor herd controlled by an overhead blimp for the exploration of Titan. The blimp is inserted into the atmosphere of Titan and self-inflated, then it descends to the surface and deploys several mobile sensors (called sondes), these sondes are guided to different areas of scientific interest by the blimp using the aid of overhead images. A simulation study was investigated of three sondes moving in formation across the surface using sonde-to-sonde and sonde-to- blimp communication for full state estimation of the four node network (three sondes + one blimp). J. Hyams et al [36] developed a cooperative navigation strategy for microrovers, and experimented with a little robot controlled by a big robot with more sensors, the big robot used colour segmentation methods to estimate the position of the small robot relative to the landmark. A. Elfes et al [31] described a concept for air- ground teams in which a blimp transmits overhead pictures to the ground rover, the ground rover uses this information as well as its own data collected for the path planning. The proposed path planning process and algorithms was demonstrated in laboratory with several robots and an overhead camera.

Map Building Studies

Map building through a team of ground robots was studied by M. Lopez-Sanchez et al [34]. Each robot in the team has similar capabilities, each robot’s control system commanded a partially random behaviour that allowed each of them to map different areas. Each robot recorded its actions and perceptions, and shared the data it collected with a central host computer and other robots. R. Grabowski et al [35] also built a team of robots to cooperatively build maps. The concept was similar to Lopez-Sanchez et al in that multiple robots explore unknown territory and build local maps. These maps are then transmitted to a team leader for map merging.

Localisation Cooperation Studies

L. Chaimowicz, et al [32] described a set of experiments comparing methods of localization of ground robots using aerobots. They used a 9m overhead blimp as the aerobot equipped with GPS, 3-axis IMU, and camera. For ground robot, an in-house autonomous truck robot was used. Study showed that rover localization based on overhead images was very accurate with 0.1m precision, compared with the 1m accuracy using GPS/IMU, and 3m accuracy using GPS only. R. Vaughan, et al [33] tackled the reverse problem of localising an aerobot using ground rover GPS. In the experiment, a ground robot sent rover location via GPS measurements to an overhead flying helicopter, which then used images of the ground robot to localise its own position. Using this method, 5m location accuracy for the helicopter was achieved.

5.3 Cooperation Strategies

Aerobots and rovers’ cooperative methods are considered in the following ways:

40  Imaging and Mapping  Localisation and Navigation/Route planning  Communication  Transportation/Delivery/Retrieval

5.3.1 Imaging and Mapping:

An aerobot serves the role of creating/merging maps from different sources of images from multiple robots: aerobot images, ground rover images. An aerobot will be able to capture broad regional view, middle resolution images of ground features, while rovers capture local high resolution images and subtle ground features that may be hidden from the aerobot’s perspective. Rovers then identify these features and transmit the complementary images back to the aerobot for the mapping.

The aerobot’s images will be the main source of information for mapping and planning for rovers. Orbiting spacecraft take low resolution/wider field of view images, overhead spacecraft images can be used to plan navigation for the aerobot. Air and ground vehicle information can be then fused onboard the aerobot to create a map of the terrain that is under investigation. Methods of map fusing have been well-established including occupancy grids and landmark based topological mapping. [33]

5.3.2 Localisation and Navigation/Route planning:

As the mast-mounted stereo cameras on Pathfinder Lander used in conjunction with the Sojourner rover to localize the 3-D positions of targets, hazards, and other objects. An aerobot serves a similar function in localisation.

Rover localisation can be realised in two ways: rovers use their own sensors to estimate its location and orientation (self-localization), or rovers get the location and orientation from overhead image (image-based localisation). In collaboration, an aerobot localises rovers using image-based localisation, in the meantime, rovers performs their own self localization and send this information to the aerobot. The aerobot performs data fusion to combine the estimates into a better estimate.

The current navigation and route planning technique for Mars rovers require operators on Earth to point out the waypoints. While in cooperative navigation and route planning, using the aerobot’s map, computing resources and rover localization, the aerobot will compute an optimal path for rovers to follow to reach their destination while avoiding obstacles, minimising human intervention. There has been a large body of work on motion planning and cooperative control of many robots. Common methods of path planning in robotics include graph search techniques, artificial potential field approaches, and heuristic rule-based approaches [28]. In the cases where smaller obstacles or subtle ground features that hamper the rover route are hidden from the aerobot’s view, the rover itself should detect the unseen obstacles and report back to the aerobot to calculate another path plan.

41 5.3.3 Communication:

Rovers communicating via an overhead aerobot instead of directly with an orbiter will have several benefits. Firstly, having a closer distance makes a simpler communication subsystem on a rover possible, which in turn will spare more room for scientific payload; Secondly, longer communication time could result in a reduced data storage subsystem on a rover and higher data return possibility from surface rovers. Thirdly, there always exists a fundamental trade-off between computation and communication. The more information can be transmitted, the less computation has to be done onboard, therefore, the rover to aerobot communication will ease the computation demands on rovers.

5.3.4 Transportation/Delivery/Retrieval:

Taking advantage of the aerial platform’s mobility, an aerobot can transport and deploy small rovers/science probes at different geographically separated sites, and hazardous locations that are unreachable conventionally. For instance, Canyons or the interior wall of craters (eg. Victoria Crater) are inaccessible to a rover from outside. An aerobot taking the role of a carrier platform can facilitate rovers into such sites and take unprecedented measurements. Retrieval and re-deployment of rovers/science probes virtually multiply the payload usage.

Figure 5-1 illustrates the collaboration methods:

Transportation/Delivery/Retrieval Figure 5-1 Aerobot-Rover Cooperation Strategy

42 Strategically, aerobots possess a better position and therefore have a more important role in the aerobot-rover team. It is natural that aerobots are physically bigger than rovers and hence have greater ability to generate power, accommodate computing resources, and have more sensors and higher communication bandwidth. It is also essential that the aerobot have those abilities for the realising of cooperation strategies. It is to be noted that the implementation of these capabilities will require achieving a high degree of aerobot autonomy.

5.4 Vehicle Selections

Aerobot selection:

The following parameters are identified as key in assessing aerobots’ suitability to fly and cooperate with rovers on Mars:

 Vehicle Mass and Size  Payload Capability  Manoeuvrability/Controllability  Mobility/Terrain coverage  Flight Endurance  Degree of surface interaction  Adverse weather capability  Complexity/Reliability

Trade-Off Weightings

For interplanetary missions, the major cost driver and the key metric that determines cost is mass. While payload is the core and the meaning of a mission, payload capacity essentially decides the science capability of a mission. Therefore these two parameters are placed at the highest weighting. The mission objective of a wide area investigation and the need to take surface contact samplings in specific terrain targets were the significant issues driving the selection criteria towards aerobots. To fulfil such requirements, aerobots’ controllability, mobility, endurance, and surface interaction are crucial.

In the scale of 1 to 10, 10 being the highest importance, key parameters were weighted as in Table 5-1:

43 Vehicle Mass and Size 9 Payload Capability 9 Manoeuvrability/Controllability 8 Mobility/Terrain coverage 7 Flight Endurance 7 Degree of surface interaction 6 Adverse weather capability 4 Complexity/Reliability 3 Table 5-1 Trade-off Weightings

Trade-Off Scores

In the evaluation of each type of aerial vehicles, scores are giving from 1 to 5, the higher score indicates higher compliance in that regard.

Vehicle Mass and Size, Payload Capability, Manoeuvrability/Controllability

Robotic balloons typically only have control of flight altitude. Their motion is passive, provided by winds which are not well studied on Mars. Attractive features are that they can potentially cover very long distances, long endurance (global travel) with very high mass and energy efficiency. Airships are like balloons but with a propulsion system to provide manoeuvrability. In terms of mass and complexity, they are between balloons and HTA aircraft. Aircraft of both the fixed-wing and rotary-wing types, at the expense of higher structure mass and complexity, could provide the maximum controllability of motion.

Vehicle mass and size Balloon 5 Airship 4 Glider 3 Aircraft 2 Rotorcraft 1 VTOL Aircraft 1 Table 5-2 Mass & size scores

Payload Capability Balloon 5 Airship 4 Glider 3 Aircraft 3 Rotorcraft 2 VTOL Aircraft 2 Table 5-3 Payload capability scores

44 Manoeuvrability/Controllability Balloon X Airship 3 Glider 3 Aircraft 4 Rotorcraft 5 VTOL Aircraft 5 Table 5-4 Manoeuvrability/Controllability scores

Mobility/Terrain coverage, Flight Endurance, Degree of surface interaction

Balloons and airships can stay buoyant without propellant. They fly in high altitude, and have the longest duration and the greatest terrain coverage. For an airship, propulsion failure will not claim the end of mission but reduces it to a balloon that flows without active flight control. Thanks to LTA vehicles’ high payload to weight ratio, they have great potential to carry and release science probes to conduct surface experiment. In particular, an airship with its station-keeping ability has the possibility of more surface interaction.

In the contrary, HTA aerobots expend significant energy simply staying airborne. The huge energy demand determines limited mission duration and coverage. For fixed wing aircraft, it is the end of mission once vehicle lands. Rotorcraft and VTOL aircraft are capable of temporary station keeping, soft landing and take-off, and in-situ soil/rock sampling; they consume the most energy but have the highest degree of surface interaction amongst the other aircraft.

Mobility/Terrain coverage Balloon 5 Airship 4 Glider 2 Aircraft 3 Rotorcraft 2 VTOL Aircraft 3 Table 5-5 Mobility/Terrain coverage scores

Flight Endurance Balloon 5 Airship 5 Glider 2 Aircraft 3 Rotorcraft 1 VTOL Aircraft 3 Table 5-6 Flight endurance scores

45 Degree of surface interaction Balloon 3 Airship 4 Glider 2 Aircraft 3 Rotorcraft 5 VTOL Aircraft 5 Table 5-7 Degree of surface interaction scores

Adverse weather capability

On the whole, LTA aerobots are less affected by the Martian atmosphere than HTA aerobots. However none of the vehicles are able to cope with the severe and versatile weather on Mars easily. Airship and rotorcraft are affected most by the windy weather in terms of control and manoeuvrability.

Balloon 2 Airship 1 Glider 2 Aircraft 2 Rotorcraft 1 VTOL Aircraft 2 Table 5-8 Adverse weather capability scores

Complexity/Reliability

Major deficiencies of HTA vehicles are the complexity of the vehicles themselves (especially rotorcraft & VTOL aircraft) and the complexity of deployment. In contrast, LTA vehicles are relatively simple in structure, light weight and mid-air deployable.

Balloon 5 Airship 4 Glider 3 Aircraft 3 Rotorcraft 1 VTOL Aircraft 1 Table 5-9 Complexity/Reliability scores

46 Trade-Off Result

Table 5-10 Trade-off Table

As can be inferred from the trade-off table, airship outstandingly got the highest score, and on most accounts, better suited to be the aerial platform than other vehicles. Balloons although with several desirable characteristics are excluded here, because their lack of manoeuvrability, which is crucial to the mission of investigating a predetermined terrain.

A robotic airship has unique capabilities and features that made it an ideal aerial platform for the mission. Fundamentally, an airship derives its lift from aerostatic forces, rather than aerodynamic forces, therefore, it is not required to spend a significant amount of energy to stay aloft, but only to move between locations or to counteract the drift caused by wind when station-keeping. Combining the long-term airborne capability of balloons with the manoeuvrability of airplanes, an airship allows flight path control for surveying, long range and close ground observations, station-keeping for long term monitoring, transportation and deployment of rover/scientific probes across vast distances (particularly key science sites that are inaccessible to rovers). Also an airship is more suitable for remote sensing, because the aerodynamic characteristic of an airship makes it stable and low vibration. Hence, it has fewer noise and disturbance on electronics than airplanes or rotorcrafts.

From those, station keeping ability is a significant advantage in executing rover localisation and route planning, where the aerobot needs to stay above the rovers observing. To communicate and command with the rovers, aerobots also have to be able to stay in sight of the rovers’ view since there’s no ionosphere to bounce long distance radio signals on Mars.

Rover selection:

Due to the mass constrain, only small rovers (microrover or nanorover) are applicable as deliverables. Moreover, a small rover is less self-sufficient and therefore has the potential to benefit more through cooperation with an aerobot in every sense. Further selection between microrover and nanorovers is discussed in Chapter 7.

47 5.5 Proposed Cooperation Architecture

Configurations

Airship falls into three structural types, rigid, non-rigid, and hybrid.

The rigid type has an internal load bearing structure with a textile covering. The is contained in non-pressure gas bags in separate compartments. This type of airship (eg. Zeppelin) is predominant during early years of airships. They are outdated with most airships operating today being non-rigid.

The non-rigid airship (also known as blimp) has no rigid internal structure and obtains its hull shape only by internal overpressure. The only solid parts are the tail fins and the gondola. The advantage is obvious that the airship is many times lighter than a comparable rigid airship.

Hybrid airships most notably are the lifting body type where dynamic lift is more significant over the overall lift than conventional ellipsoidal airships, and the envelope is shaped to maximise this. However there has not been a successful demonstration of a lifting body airship.

Traditionally, an airship achieves altitude control and landing via a ballonet system, in which atmospheric gas is brought within the envelope to increase the mass within the airship’s volume, thereby reducing the buoyancy until equilibrium is achieved at a higher atmospheric density at lower altitude. The atmospheric gas is brought in using a compressor or a blower and the rate of collecting or venting gas determines the rate of climb and descent of an airship. [15]

Different from the terrestrial airship, the proposed airship doesn’t have a ballonet system. It would stay in constant pressure altitude, for that simplifies the structure and reduce the mass of the envelope (a ballonet system accounts for around 10-15% of the total airship mass). [15] In supplement, a tether system is applied where rovers can be lowered to the surface for sample collection and measurements. More to the point, airship landing is an extremely difficult and dangerous operation. In terrestrial applications, instead of landing, airships dock at the mooring mast when loading and unloading cargos. On Mars, such task would be more difficult, besides deploying rovers to inaccessible hazardous sites put airship landing at risk as well. The tethered intermediate platform renders a promising solution.

The proposed architecture is composed of one airship, a retrievable tethered platform (depot) containing rovers. The aerobot will explore the target terrain in concert with surface rovers. Throughout the flight, the aerobot is approaching target as well as searching for terrain features of interest, as it finds the target feature, it hovers and deploys the rovers over the site.

48 The architecture expands the planetary exploration capabilities allowing high-resolution targeted observation, and augmenting observations at atmospheric altitudes with in-situ surface observations. The rover platform enhances the aerial system’s surface interaction, the aerobot in turn, provides a greater deployment accuracy (MERs landing error has an order of tens of kilometres) and counteracts the crucial drawback of the ground system which is their limited coverage. Furthermore, the deployment of rovers from an aerial platform simplify the atmospheric entry and deceleration hardware thus reducing overall mass and permitting more scientific payload.

Figure 5-2 Aerobot-rover schematic diagram

Team investigating a piece of terrain will follow the following phases:

Cruising & Terrain Searching Phase: Science measurement Sensing & generating terrain map with imager and radar Computing landing area away from hazard Deployment Phase: Winching down surface platform from the station-keeping airship Depot deployment Rovers driving out Station-keeping Phase:

49 Aerobot station-keeping or hovering above in line of sight Science measurement Site imaging, map generating Rover communication, path planning, navigation Rover recovery Phase: Rover driving in Depot reeling up

Figure 5-3 depicts a scenario of terrain investigation and sequences, one or more rovers deploy from the intermediate platform, and can be recovered.

rover recovery hovering leaving deployment cruising

surface experiment

Figure 5-3 Cooperation Scenario & Deployment Phases

The aerobot either is instructed by operators on Earth or chooses autonomously (using its acquired information, such as unique rock detections) the potential sites of interest. When at these sites, it can perform in-situ measurement and ground sampling by rover deployment. Rovers thereby gather data that complement the remote sensing data obtained by the aerobot, resulting in a more complete picture of the surface under examination.

Auxiliary, if budget permits, microprobes/penetrators can be equipped and air-deployed from the aerobot to enhance the mission. Microprobes will give preliminary readings on the surface composition or presence of signs of life to determine if the passing terrain is worth the delivery of rovers.

50 6 System Design Overview

To establish the feasibility of using the proposed system on Mars, an airship calculation and sizing is necessary.

To start the analysis, some basic mission parameters are defined. These are as follows:

Flight Altitude: 5km. At this height, most topography is low-lying enough to be avoided.

Mobility Requirement: 5m/s forward speed wrt. assumed windspeed of 10m/s. Forward speed of 5m/s is determined considering both the need of low speed high stability and the time taken in scanning the target area –it will take two and a half sols to scan the diameter distance once. As with the Earth’s environment, the wind speed on Mars is highly variable with seasons and locations. At the Viking landing sites, the wind speed were in the range of 2 to 7m/s, during a dust storm, it could reach to 50m/s. It is assumed an average Vwind =10m/s (conservative wind speed value) for calculation.

6.1 Environmental Conditions

The Mars atmosphere condition is firstly required for airship evaluation.

According to the Gas Law, atmosphere pressure P is proportional to the product of atmosphere density and temperature P  T , the pressure at altitude of h km is:

P0  hTh Ph  Equation 6-1  0T0

Equation 6-2 and Equation 6-3 calculate the atmosphere temperature and density at a given altitude (apply from above the surface to nearly 10km in low latitude).

2 3 4 5 6 7 Th =238.74 - 34.488h + 35.133 h - 15.96 h + 3.7315 h - 0.47352 h + 0.030962 h -0.000817 h Equation 6-2 [45]

5 2 7 3  h =0.014694 - 0.001145h + 4.663810 h - 9.773710 h Equation 6-3 [45]

6.2 Airship Evaluation

Based on the environmental conditions specified previously, an evaluation using an airship for exploration was performed. The requirements for airship were that it would need to be capable of station-keeping at a given location as well as operating for a long enough period of time

51 The final airship configuration is a standard cylindrical shape with three conventional tail fins for stability and a motored propeller system for propulsion and control.

This configuration is shown in Figure 6-1.

Solar Array

Control Fins

Gondola Propeller

Figure 6-1 Airship Configuration

6.2.1 Airship sizing

An airship calculator was programmed in MATLAB for the development. The methodology is described by the following.

6.2.1.1 Mass and volume of the lift gas:

Equations for the volume of the envelope and mass of the internal gas are derived for an altitude of h km. the derivations are as follows:

Known Parameters Symbols(unit) 3 Atmospheric Density  h ( kg / m )

Atmospheric Pressure Ph (Pa) 3 Relative Atomic Mass of Helium RAM He = 410 kg / mol 3 Relative Molecular Mass of Hydrogen RAM H 2 = 210 kg / mol

Temperature Th (K)

RHe  2077J / K  kg

Molar Gas Constant for Helium RHeuniversal  .8 308J / K  mol

52 RH 2  4124J / K  kg

Molar Gas Constant for Hydrogen RH 2universal  .8 248J / K  mol

Total Balloon Mass M tot (kg) Table 6-1 Known parameters definition

Parameters Needed Symbols(unit) 3 Density of the gas  H 2 or  He ( kg / m )

Mass of internal gas M H 2 or M He (kg) 3 Volume of airship Vas =VH 2 or VHe ( m ) Table 6-2 Required parameter definition

The principles of calculations and deviations of the airship volume and the mass of internal gas are as follows:

First, according to the Gas Law, the relationship of atmospheric density, pressure and temperature is given in Equation 6-4 and Equation 6-5:

Ph   H 2 RH 2Th (R in J/kg.K) Equation 6-4

PhVH 2  nRH 2Th (R in J/kg.mol) Equation 6-5

Thus the internal gas volume can be expressed in Equation 6-6

M H 2  RH 2universal Th VH 2  Equation 6-6 RAM H 2  Ph From Archimedes’ principle, lift equals the weight difference of gas/fluid displaced. If the lift is equal to the airship weight for an equilibrium condition:

M tot = VH 2  h - VH 2  H 2 Equation 6-7

From Equation 6-6 and Equation 6-7, the mass of lift gas can be derived as in Equation 6-8.

M tot RAM H 2 Ph M H 2 = Equation 6-8 RH 2universalTh ( h   H 2 )

Using the result from Equation 6-8, substitute back to Equation 6-6, the volume of the airship can also be known.

53 6.2.1.2 Drag & Balloon size

The power required by the airship is given by the power needed to operate the onboard systems, payload and the power needed to overcome the drag to keep the forward speed. The airship drag (D), given by Equation 6-9 is based on a volumetric drag coefficient

(Cdv ) and airship volume (Va ). The volume is determined from length and finesse ratio (f) of the airship. This relationship is given in Equation 6-10. For this analysis, the airship has to be able to station-keep as well as manoeuvre at a speed of 5m/s. Therefore, the velocity (V) at which it is operating is the wind speed plus 5m/s.

1 3/2 D  V 2C V Equation 6-9 [46] 2 dv a 4 V   (l  d 2 )  l  3 3V  f 2 4/  Equation 6-10 as 3

The drag coefficient is based on the airship finesse ratio (f) which is the ratio of the length (l) of the airship to its diameter (d) as given by Equation 6-11, and the volumetric drag coefficient is dependent on the finesse ratio given by Equation 6-12: l f  Equation 6-11 d

2 3 3 4 4 5 5 Cdv = 0.23175 - 0.15757f + 0.04744 f - 7.041210 f + 5.153410 f - 1.483510 f Equation 6-12 [46]

6.2.2 Power and Propulsion System

Since the Martian atmosphere is made up of carbon dioxide and , air-breathing engines were ruled out. For existing engine technologies, this leaves mono propellants (eg. Hydrazine). Mono propellant Hydrazine was used to power the propeller driven Mini-Sniffer and was proposed by Girerd [19] for a Martian airship. However the use of hydrazine would significantly limit the mission endurance. Girerd’s airship design had an endurance of one month. However it was a much smaller airship than what is considered here.

The only other means of fuel burning propulsion is rocket propulsion and this is neither a long endurance nor a low speed solution. Propulsion by propeller would seem the only feasible option for a vehicle operating in this very low speed range of 5m/s.

Because of the desire for long duration flight, the airship must use long-lasting power resource. Two types of propulsion/power systems were considered, namely solar photovoltaic system and nuclear radioisotope system.

54 6.2.2.1 Power system

Currently, the most mature and cost-effective power generation technologies for planetary applications are solar cells and radioisotope thermoelectric generator which obtains its power from radioactive decay.

The power system is a key component to airship operation. Power will be used for propulsion system, onboard electronics and payloads.

6.2.2.1.1 Solar photovoltaic system

Solar array:

The design case select spring equinox where the solar power will be an average value and at the equator where the day is a fixed length of 12.3 hours of daylight and 12.3 hours of darkness.

It is designed that the array is placed on the upper half of the airship envelope in the same way as many designs of terrestrial high altitude airships. A critical factor in the airship design is minimising mass. Because of this desire, a thin film array was selected. Thin film arrays have a number of characteristics that are desirable for an airship application. Primarily, thin film arrays can be made very lightweight. The material that makes up the array is only 1 to 2 microns thick, whereas the typical thickness of a rigid silicon solar cell is approximately 250 to 350 microns [46]. In addition to being lightweight, thin film array are also flexible, it could be stored and deployed easily without damaging solar cells, this is a major benefit for the curved surface of the airship.

The typical mass of an amorphous silicon thin film array is approximately 0.12kg/ m 2 , efficiency sc about 8%. [45] The solar array mass ( M sa ) is given by the specific mass of the array multiplied by solar array area ( Sa ).

M sa  Sa  Equation 6-13

The total output of solar array is calculated based on the incident solar radiation on the array considering attenuation due to the atmosphere. The array output or power available (P) is given in Equation 6-14. It is based on the mean solar intensity at the I planet’s orbit (589W / m 2 ) and the attenuation due to the atmosphere ( = e  =0.427) I o [45]. Taking account of the airship envelope's curvature and the variation of local solar elevation angle, calculation simplifies by taking a factor of 0.5.

 P  5.0 Sa I ome  sc Equation 6-14

55 Regenerative Fuel Cells (RFCS)

In order not to drift away, the airship has to be fully controllable, that is to be able to operate during the night. When solar energy is used in applications that require power at night some form of energy storage is necessary.

The hydrogen-oxygen RFCS has been promoted as one of the most favored energy storage technologies for solar electric power in aerospace applications, mainly due to its high specific energy (up to 1000Whr/kg) and power delivery (4kW electric power in discharge) [49]. A RFCS could provide much higher specific energies than any advanced battery system and potentially the highest storage capacity and lowest weight of any non-nuclear device [49].

ESA believes the immediate future of space and planetary exploration will be based upon electric propulsive technology. NASA is investing heavily in the development of fuel cell technology, as they expect high potential from regenerative fuel cells. Many concepts were ruled out due to the unavailability of high energy storing batteries. Fuel cells could provide a solution to this problem in a few years time. It is then decided that using regenerative fuel cell for continuous operation.

energystored  fuelhours M  Equation 6-15 fuelcell fuelcelldensity  efficiency

6.2.2.1.2 Nuclear radioisotope system

RPS

Radioisotope Power Systems (RPS) is considered the future power source for deep space missions and is currently under intensive development. The potential of radioisotope power was demonstrated by Viking Lander, which was powered by an RTG and operated on Mars for over six years. RPS has multiple benefits, it is relatively compact, inherently reliable, and can not only provide a long-lasting continuous power source, but also the waste heat produced can be used for thermal control. The principal disadvantage of RPS is its comparatively high cost and public concerns regarding nuclear pollution.

For this design, an Advanced Stirling Radioisotope Generator is selected. The Advanced Stirling Radioisotope Generator (ASRG) is now under NASA’s development as a next generation of Radioisotope Power Systems, which will produce power more efficiently, reducing the Plutonium-238 fuel requirements and increasing specific power.

ASRG will produce electric power approximately 7 W/kg ( Ps ), and will use only one fourth of the Pu-238 that would be required for a comparable RTG. Engineering unit generator is currently being fabricated, and projected to be available for mission use as early as 2012.

56 6.2.2.2 Propulsion System

The design and optimisation of a propeller for efficient flight is a crucial aspect of the overall success of an aerial vehicle. However a design of a Martian propeller is beyond the scope of this thesis, some preliminary thoughts are required on the size and mass of this key component.

The electric propulsion system is composed of 5 components:  power conditioning system  motor controller  electric motor  gearbox  propeller

The power train is shown in Figure 6-2

Figure 6-2 Propulsion System Drive Train [45]

The operational efficiency of the power system must be taken into account for sizing the airship. For a basic electric propulsion system the efficiency is given by the following.

 p  mcem g prop Equation 6-16 [47] Equation 6-16 represents the efficiencies of the main power system components, which include the electronic controls, drive motor, gearbox and propeller. The operational efficiency associated with each of these components is given in Table 6-3. These efficiencies are representative approximations, which are subject to change based on a more detailed system and component design.

57 Component Efficiency Control Electronics 0.98 Motor 0.9 Gearbox 0.9 Propeller TBD Total System Efficiency TBD Table 6-3 Component Efficiency for Electric Powered Aircraft [45][47]

The electric motor mass ( M em ), motor controller mass ( M mc ), gearbox mass ( M g ), and power conditioning system mass ( M pc ) calculations are based on a linear scaling with power output.

P M  mc (minimum 0.5kg) Equation 6-17[45] em 1291 P M  (minimum 0.1kg) Equation 6-18[45] mc 6233 P  M  em mc (minimum 0.3kg) Equation 6-19[45] g 3278 P M  (minimum 0.2kg) Equation 6-20[45] pc 1000

Propeller

Propeller’s tip Mach number must be less than one to avoid shock waves. Shock waves produce significant increases in drag on the blade and reduce its efficiency, they can also cause propeller blade to flutter and potentially destroy the blade [46]. Therefore maintaining a subsonic tip speed of propeller is necessary for stability as well as the efficiency of the propeller.

-The speed of sound

To calculate the propeller mass and size, the speed of sound on Mars is required, and is given by Equation 6-21

a  RT Equation 6-21 Where,  = specific heat ratio (dimensionless) R = specific gas constant ( J / K kg ) T = atmospheric temperature (K)

58 -Propeller sizing

The propeller mass ( M prop ) is based on the blade diameter, volume, material density

(  prop ), and the number of blades ( nb ). Briefly, following the method presented by Colozza [46], the propeller diameter is given by Equation 6-22

D 2 Equation 6-22 d p  2 2 ct ((aM ) V ) The volume of the propeller blade can be calculated using Equation 1-23, the relationship is based on geometry characteristic of the SD8000-PT low number airfoil which provides good lift-to-drag characteristics at low Reynolds numbers.

5 3 V prop  .9 2573910  d p Equation 6-23

The total propeller mass is given by Equation 6-24.

M prop   prop nb 1(  Fb )V prop Equation 6-24

-Propeller efficiency

It was assumed that a 2 bladed propeller was utilized. Following Colozza [47]’s methodology, propeller efficiency is calculated as follows:

2 3 4 ct   .0 012122  .0 14577J  .0 1408J  .0 05374J  .0 0068444J Equation 6-25

2 3 4 c p   .0 012752  .0 094954J  .0 053694J  .0 017534J  .0 0007872J Equation 6-26 V J  Equation 6-27 (aM ) 2 V 2

ct J  prop  Equation 6-28 c p

Where, ct = thrust coefficient

cp = power coefficient J = advance ratio

 prop = propeller coefficient a = speed of sound M = Mach number V = flight velocity

59 Now, with propeller efficiency, the total efficiency  p of the power system can be derived from Equation 6-16.

Total power required can be determined by Equation 6-29. Total power is composed of two parts, one part is to be used to push airship against drag through the atmosphere whose magnitude depends on the airship’s shape and volume; the other part is to be used to provide onboard subsystems in operation.

The preliminary power requirements for the fixed systems are assumed in Table 6-4.

System Power Level Communications 50W Control and Operations 50W Payload 50W Table 6-4 Assumed Fixed System Power Level

DV P  150 Equation 6-29  p Where: D = Airship drag V = Flight and wind velocity

 p = Propulsion system efficiency

6.2.3 Envelope and Fins

Airship balloons are broken down into two parts: the envelope, and fins.

6.2.3.1 Envelope

Maximizing the useful mass of scientific payload requires reducing the masses of other components. Advanced lightweight and strong materials for envelop can lead to substantial increases in useful payload mass. The material selected was the Planetary Aerobot Composite Material developed by NASA JPL which was used in Girerd’s aerobot study [19]. This material has a strong yield stress property, it is made of Mylar glued to polyethylene with a loose scrim of Kevlar. The material has an areal density 2 2 ( envelope ) of 0.016 kg / m to 0.02 kg / m . The envelope mass is based on the volume of the airship, and the density of the envelope material (  e ).

The surface area of the envelope is given in Equation 6-31

d   arccos( ) Equation 6-30 [62] l

60  S  2 (d 2  l 2 ) Equation 6-31 [62] envelope tan( ) Mass of the envelope:

M envelope   envelope  S envelope Equation 6-32

6.2.3.2 Fins

The three conventional tail fins are for control and stability purpose. The fin size is based on the volume of the airship. To determine the fin area, the generalised fin area to 2 3 airship volume ratio ( R f ) 0.021 m / m was taken from High Altitude Long Endurance Airship [46]. The fin mass is decided in Equation 6-33.

M f  2.1 R faVas e Equation 6-33 Where a 20% increase in the in mass was used to account for the internal structure of the fins.

To estimate the total mass of the airship, it was broken down into a number of subsystem components, some components were assumed to be fixed masses.

Component Mass (kg) Flight Control&Navigation System 10 Communications Equipment 10 System/Gondola Structure 20 Payloads 40 Table 6-5 Fixed Component + Payload Masses

The airship’s total mass is the sum of all components’. A margin is used to account for various miscellaneous items and is assumed to be 10% of the stringent airship mass

M total  M as  1.1  (M h2  M envelope  M f  M sys  M p  M sa  M rfcs  M payload )  1.1 Equation 6-34

Where: M h2 = internal gas mass

M envelope = envelope material mass

M f = fins mass

M sys = fixed components mass

M p = electric propulsion system mass

M sa = solar array mass

M rfcs = regenerative fuel cells mass

61 M payload = payload mass

M as = total airship mass without margin

M total = total airship mass with margin

6.3 Results and Analysis

Gas Selection

Helium and hydrogen are considered as the option gases. Calculation shows that a hydrogen airship can carry the same payload with a smaller overall mass and volume than a helium airship. This is because hydrogen has a lower mass density than helium.

Hydrogen & Helium airship comparison 150

100 g k

- 50

s s a M

d a o l 0 y a P

-50

-100 0 100 200 300 400 500 600 Airship Total Mass - kg blue=hydrogen, red=helium

Figure 6-3 Hydrogen & Helium solar airship comparison

Hydrogen is a molecule and is larger in comparison to atomic helium. That means hydrogen gas is less likely to penetrate through the airship envelope than helium gas. Too much gas leakage would result in the airship losing lift and termination of the mission.

Another factor need to be considered is the safety of the gas. Helium is an inert gas, while hydrogen is active and flammable in an oxygen rich environment, which is why on Earth helium is preferred over hydrogen. On Mars, the atmosphere consists of

62 mainly carbon dioxide (95.32%), with only 0.13% oxygen (compared to 16% on Earth).

As hydrogen is not reactive withCO2 , it would be safe to use as the internal gas on Mars.

Therefore hydrogen is selected as the superpressure airship internal gas.

Airship Selection

Both the solar powered and radioisotope generator powered airship design were developed. Effect of Payload Mass VS Airship Mass & Effect of Finesse ratio 150

100 g k

- 50

s s a M

d a o l 0 y a P

-50

-100 0 100 200 300 400 500 600 Airship Total Mass - kg Finesse ratio: black=2.5 red=3 blue=4 yellow=4.5 =5 cyan=6 magenta=7

Figure 6-4 Solar airship payload mass and finesse ratio's influence

Figure 6-4 illustrate the effect of finesse ratio on the solar powered airship, the envelope shape has tricky but great influences on the airship’s capacity. This is due to the complex relationship between envelope shape/mass, corresponding drag force and the power required to overcome that drag. The smaller the finesse ratio is, the more round is the envelope, and thus the less material is used, envelope mass is smaller. However, rounder shape means more drag, consequently greater power is required, thus increases the power system mass. The results indicate the airship with a finesse ratio of 4 would be the best configuration. When finesse ratio is smaller than 4, the increase of power system mass outweighs the advantage of small envelope mass. Vice versa, when finesse ratio is greater than 4, the increase of the envelope mass outbalanced the decrease of power system mass. Also note from Figure 6-4, as said in previous chapters, the airship indeed has higher payload to mass ratio as the total mass increases. Airship will be a

63 heavy mission, even in the optimum shape at finesse ratio of 4, it must be bigger than 242kg, or else there will be no margin left for the payload. Effect of Payload Mass on Airship Mass 250

200

150 g k

-

s 100 s a M

d

a 50 o l y a P 0

-50

-100 0 100 200 300 400 500 600 700 800 900 1000 Airship Mass - kg Finesse ratio: red=3 blue=4 yellow=4.4 green=5 cyan=6 magenta=7

Figure 6-5 RPS airship payload mass and finesse ratio's influence

Similar to the solar powered airship, Figure 6-5 illustrate the case of Sterling Radioisotope Generator powered airship. The best performance is achieved when finesse ratio is 4.4.

To have a 40kg payload capacity, the solar array and regenerative fuel cell (density 400Wh/kg) combination airship achieves a total mass of 332 kg with balloon volume of .3 15104 m3 , propeller blade diameter 0.86m, the continuous operation consumes power 1.23kW. In comparison the radioisotope generator powered airship design is much heavier (598kg) and bigger ( .5 68104 m 3 ) resulting in a 1.71kW power requirement.

64 Solar Powered RPS Operation Continuous Continuous Components Mass(kg) % Mass(kg) % Internal Gas 6.9 2.3% 12.5 2.3% Envelope 123.5 40.9% 187 34.3% Fins 9.2 3% 16.5 3% Propulsion System 3.2 1% 3.9 0.7% Payload 40 13.4% 40 7.4% Systems 40 13.2% 40 7.4% Solar Cells 25.1 8.3% Fuel Cells 54.1 17.9% Radioisotope Engine 244.1 44.9% Total 302 100 544 100 Total + 10% margin 332.2 598.4 Table 6-6 Airship sizing comparison and mass breakdown

Table 6-6 compares in detail the two aerobots, evidently the solar plus regenerative fuel cell combination exhibits a more competent option, thus is preferred.

65 7 Payload Selection

Scientific payload should be best distributed to make both aerial and surface platforms capable and complementary, hence maximizing the scientific goal.

The 40kg total payload budget is composed of aerobot scientific payload, surface rovers and a depot platform (tether+depot). The initial mass distribution is as in Table 7-7

Payload Mass Aerobot science payload ~15kg Rovers ~10kg Intermediate Tether Platform Depot ~15kg Table 7-1 Payload mass distribution

7.1 Aerobot Payload

Scientific payload should fulfill the mission’s scientific goals as well as letting the aerobot exert its advantages as an aerial platform.

Many of the instrumentation approaches devised for orbital spacecrafts and landers can be applied to aerial platforms. The gondola underneath the hydrogen envelope should provide temperature control and good protection from radiation.

A set of aerobot instrument were chosen to meet the mission goal.

Objectives Aerial Instruments Sub-surface structure Subsurface Sounding Radar Panoramic Camera Survey & Mapping Thermal Emission Spectrometer Magnetometer Thermal Emission Spectrometer Geology Geological history Panoramic Camera Atmospheric composition Mass Spectrometer Temperature Pressure Wind Meteorology sensors Environment Radiation UV sensor Thermal Emission Spectrometer Subsurface Sounding Radar Look for water Mass Spectrometer Life/biology Bio-chemicals Mass Spectrometer Table 7-2 Aerial Instruments

66 7.1.1 Mass Spectrometer

Mass spectrometer is a powerful tool for analyzing molecular compositions. The mass spectrometer on the International Space Station was chosen, for its proved high performance and small mass relative to other mass spectrometers.

A slightly modified version could be used on Mars to analyze the atmospheric composition and relative abundance of each molecule detected. It works by drawing in a small atmospheric sample, ionizing the molecules with an electron beam then measuring the mass-to-charge ratio of the constituents using radio frequency filtering. Therefore different species and the relative abundance can be identified. [54]

Mass spectrometer has the potential of indicating the presence of water, or bio- chemicals such as methane which has been detected by Mars Express. Rovers could follow up being released for further investigation in areas where such molecules are concentrated.

Mass Spectrometer Specifications Mass/(kg) 2.3 Power/(W) 15 Data rate/(bits/s) TBD Table 7-3 Mass Spectrometer Specifications [54]

7.1.2 Mini Thermal Emission Spectrometer

Currently operating on the MER mission rovers. The Mini Thermal Emission Spectrometer was chosen due to its ability to identify minerals from a distance, it analyses the infra-red radiation emitted from the ground, recognizing organic molecules and minerals formed in water. It has a distinctive ability to penetrate through Martian surface that obstruct readings from other instrumentation.

This instrument can therefore search for evidence of water, perform a geological analysis, and collect useful data for landing site selection.

Mini Thermal Emission Spectrometer Mass/(kg) 2.4 Power/(W) 5.4 Data rate/(bits/s) 240 Resolution/(m) 100 Table 7-4 Mini Thermal Emission Spectrometer Specification [55]

7.1.3 Subsurface Sounding Radar

A subsurface radar sounder emits radio frequency waves which would be reflected at boundaries of changing density at, and beneath the surface, allowing views below the

67 surface deposits into possible reservoirs of liquid and geologic layers. The subsurface structure and composition can then be revealed.

Subsurface sounding with long-range radar similar to those used on Mars Reconnaissance Orbiter – SHARAD and Mars Express – MARSIS promises to be more capable on an aerobot than on an orbiter since sounding distance is less. Besides, operation of a radar sounder from aerial platform is much easier than from orbit as the clutter signal and noise is reduced, therefore finer resolution and higher data rate can be expected.

Subsurface Sounding Radar SHARAD MARSIS Mass/(kg) 15 7.5 Power/(W) 10-60 10 Frequency/(MHz) 15-25 1.3-5.5 Depth/(km) 1 5 Resolution/(m) 7 10-100 Data rate/(bits/s) TBD 75000 Table 7-5 Subsurface Sounding Radar Specifications [56] [57]

MARSIS is selected as an aerobot instrument. This would complement the results of SHARAD and MARSIS on orbit.

7.1.4 Magnetometer

Magnetometers were proposed for on CNES Netlander Mars mission and NASA ARES Mars airplane mission. Although Mars has no intrinsic magnetic field like earth, NASA Mars Global Surveyor satellite measured local magnetic fields in some regions which are believed to be magnetized in ancient time. By measuring the magnetic field preserved in crust and rocks, the evolution of Mars and the composition of the core can be studied.

Magnetometer Mass/(kg) 1.15 Power/(W) 0.6 Table 7-6 Magnetometer Specifications [58]

7.1.5 Ultra-Violet Sensor

UV sensors tell how much ultra-violet radiation penetrates the Mars atmosphere, giving an indication as to weather life could survive on the surface. It is also worth knowing how much protection would be required for a future manned mission. The UV sensor selected was originally designed for Mars lander, it is now being proposed for ESA’s ExoMars mission. It consists of six photodiodex which measure a wavelength range of 200-400nm.

68 Ultra-Violet(UV) Sensors Mass/(kg) 0.05 Power/(W) 0.026 Table 7-7 UV Sensor Specification [59]

7.1.6 Panoramic Camera (Pancam)

A stereo high resolution camera will support geological investigations, provide scientists searching for water and life with images of surface features such as water erosion gullies, and produce broad view map for rover navigation.

The PanCam selected is the one designed for the ESA ExoMars mission. It consist of two Wide Angle Stereo Cameras (WAC) and one High Resolution Camera (HRC). The WACs provide wide angle (65° field of view) multi-spectral stereo imaging. It is composed of two identical CCD cameras and each with a filter wheel unit containing 12 filters. The HRC optics have a focal length of 130mm, yielding a field-of-view of 5°. The spatial resolution of HRC is 13 times that of the WACs at about 85 microradians per pixel.

PanCam WACs HRC Mass/(kg) 0.4 0.3 Power/(W) 1.5 0.9 Data rate/(bits/pic) 1M 1M Resolution/(pixel) 1024*1024 1024*1024 Table 7-8 Panoramic Camera Specification [60]

This PanCam could fulfill the digital terrain mapping requirements of the mission as well as providing multispectral imaging, colour and stereo panoramic images.

7.1.7 Meteorology Package

The meteorology package consists of a set of sensors which collect atmospheric information during the flight. Data acquired will permit the reconstruction of atmospheric temperature, pressure and wind profile from flight altitudes. The package consists of three instruments taken from Beagle-2 Lander:

- Pressure Sensor Instrument (PSI) - Atmospheric Temperature Sensor (TEMP) - Wind speed and Direction sensor

69 Pressure Sensor Instument Mass/(kg) 0.222 Power/(W) 0.05 Data rate/(bits/s) 16 Range(bar) 0 ~ 2 Resolution(mbar) 0.005 Temeprature Sensor Mass/(kg) 0.086 Power/(W) 0.05 Data rate/(bits/s) 67 Range(°C) -200 ~ 50 Resolution(°C) 0.02 Wind Speed and Direction Sensor Mass/(kg) 0.167 Power/(W) 0.3 Data rate/(bits/s) 670 Range(m/s) 0 ~ 80 Resolution(m/s) 0.02 Table 7-9 Meteorology Sensors Specification [61]

The Meteorology package may look insignificant in total mass, however it is not less important than others. It helps to understand the Mars environment, in preparation for human exploration.

Two meteorology sensor packages are to be place on gondola and depot respectively. This enables atmospheric measurements in different altitude to be performed simultaneously.

Instruments Mass(kg) Power(W) Mass Spectrometer 2.3 15 Mini - Thermal Emission Spectrometer 2.4 5.4 Subsurface sounding radar 7.5 10 Magnetometer 0.575 0.3 Ultra-Violet(UV)Sensor 0.05 0.026 Panorama Camera(PANCAM) 0.7 2.4 MET Package 0.475 0.4 TOTAL 14 33.526 Table 7-10 Summary of aerobot’s science payload

7.2 Rover Payload and Selection

The rover’s strength is in providing surface contact experiment in small scale coverage. Surface contact experiments vary in different kinds, each instrument has its own benefits, however they are all effectively used to determine the presence of chemicals, either on the surface or beneath it. Some of the main chemical detection experiments are listed in Table 7-11:

70 Experiment Instrument Capability Mass/(kg) Power/(W) Alpha Particle X-ray Spectrometer Elements detection 0.64 1.5 Iron content and oxidation Mossbauer Spectrometer ratio 0.5 1.6 Optical Microscope/Microscopic Imager Micron resolution 0.25 0.5 Integrated DNA sampling Organic marker compounds 0.1 1 Raman Spectrometer Molecular analysis 1.5 1.1 Infra-red Spectrometer Molecular analysis 1 3.5 Acoustic sonar system Sub-surface structure 1 1 Pyrolytic gas chromatograph and Organic/Inorganic mass spectrometer compounds 0.5 1 Laser induced breakdown spectrometer Common elements detection 1.5 2.4 Table 7-11 Contact Experiment Instruments [2][4][16][50][51][52][53]

Some of the instruments’ functions overlap, however most of them can be utilised restricted only by total mass. The power requirement of each of the instruments listed is less than a few watts.

7.2.1 Microrover

Under the current technology, a 10kg class microrover is capable of carrying payload up to 20% of the total mass. Taking sojourner rover as an example:

Sojourner science payload characteristic Instument Mass Power Atmospheric Structure Instrument/Meteorology Experiment(ASI/MET) 2.04kg 3.2W Imager for Mars Pathfinder(IMP) on the Lander 5.20kg 2.6W Alpha-Proton X-Ray Spectrometer(APXS) 0.74kg 0.8W Table 7-12 Sojourner science payload summery [2]

Thus a 2kg mass instrument collection is first selected for a microrover concept.

 Integrated DNA sampling device  APXS  Mossbauer Spectrometer  Optical microscope

Integrated DNA sampling devices have been proposed although these are still in a state of development. A sample taken from the target is introduced onto a glass slide where numerous segregated reactive chemicals react with the sample, each displaying the presence of an individual marker of some sort. The fluorescence of each of the reacted substances is then measured using an imager. Many biomarkers can tested for in this

71 way. [50] The integrated DNA sampling allows the detection of specific biological markers; the low mass device offers a powerful tool in search of life.

The APXS works by bombarding the target rock or soil to a stream of energetic alpha particles and X-rays from its radioactive element curium-244 onboard the instrument. The target material is determined from the backscatter response. The values quoted are based on the instrument used on MER missions, A high-quality analysis requires about ten hours of operation which amounts to 15WHrs of energy consumption [51]. This can be carried out while the rover is stationary and may be done at anytime of day or night. The APXS capability to detect virtually any element (except hydrogen and helium) and its low associated mass makes it a very valuable instrument. The APXS will enable chemical analysis to be performed at sites and coupled with the abrasive tool will allow invasive measurements of rocky material.

The Mossbauer Spectrometer works in a similar way to the APXS, but measures the Mossbauer Doppler shift effect which reflects the relative amount of iron in various valence states [52]. The Mossbauer spectrometer has the capability to measure iron isotope content which the APXS cannot do alone. This will characterise different types of rock and improve understanding of Mars’ development and formation.

An optical microscope works by observing reflected or emitted light from a target material. The Beagle-2 microscope had a spatial resolution of 1.5  m in the spectral band of 435-865nm (filtered for several bands from visible to NIR) [53]. The optical microscope will allow surface features to be analysed optically.

Instruments Mass Comments Integrated DNA 250g Low mass, great potential sampling Mossbauer 500g Mineral classification Spectrometer Optical microscope 250g High resolution imagery APXS 640g Low mass (heritage from MER) Total 1.64kg Great capability Table 7-13 Rover payload instruments

72 Figure 7-1 Microrover illustration

7.2.2 Nanorovers

Alternatively, using multiple nanorovers instead of a single microrover, these robots can work together with scientific equipment distributed between them.

Muses-C type rover (introduced in chapter 4) promises great potential. Although it was to be used on an asteroid, application of Muses-C to Mars is straightforward, because the JPL nanorover research program was created with Mars as the original destination. With minor modifications, the Muses-C type rover can be adapted to operate on Mars. The original prototype has the following characteristics:

Rover Characteristic Value Mass 1.3 kg Size 14*14*6 cm (30% of Sojourner) Power 2.9W Panoramic and Near-Imaging Camera Near-Infrared Spectrometer Instruments Alpha X-ray Spectrometer(AXS) 130g Data rate up to 38.4kbits/sec Table 7-14 Muses-C rover characteristics [42] [44]

It is devised that 4 nanorovers carry the same payload selection that was selected for microrover. Due to the payloads’ low mass and power consumption characteristics, it is

73 estimated that the modification should not be too great to accommodate the instruments, each should be confined within the mass budget (2.5kg each). Four surface contact instruments: Integrated DNA sampling device, APXS, Mossbauer Spectrometer and Optical microscope carried separately on four nanorovers, each one is capable of one type of experiment independently, while collectively, they can achieve not only the same objective as the single microrover does, but also several distinct advantages:

 First of all, robustness is improved through multiple rovers, single rover failure is not critical to the mission, though the surface experiment capabilities would be restricted, the surface platform function will continue on the whole.

 Surface experiments would be more efficient with distributed investigations. The APXS instrument must be operated for several hours to achieve a full sample of data, during which rover has to stay motionless. In the case of a single microrover, that means the ceasing of other operations. Whereas in the case of nanorovers, one operation doesn’t constrain the other experiments from being conducted. Every instrument can be used fully during the surface investigation period.

 These rovers can act either on their own analysing samples or work collaboratively examining or imaging a specific site. Figure 7-3 envisages one scenario of nanorovers cooperative investigation.

Figure 7-2 Cliff descent by cooperative rovers (Courtesy of JPL)

Therefore, the nanorover scheme is chosen as the rover platform.

7.3 Intermediate Platform

The depot is attached beneath the gondola through a tractable tether. It contains rovers and protects them from crashing during landing. The depot is simply an aluminium box with styrofoam type material underneath to cushion and absorb shocks while landing. In order to construct the atmospheric vertical profile, a suite of meteorological sensors and an altimeter is devised on the platform to complement the airship’s measurement at different altitude. The altimeter not only works in conjunction with sensor readings, but also the soft-landing require precise altitude monitoring. Existing commercial radio-

74 altimeters, commonly used in aviation are available as extremely lightweight and low power instruments.

tether sensors & solar array on altimeter upper side

depot

Figure 7-3 Intermediate platform

The airship will float up to an altitude approximately 6.5km when the depot platform is deployed on the surface. An 8km tether has been baselined, the tether needs to be made very thin to minimise drag yet strong enough to bear the platform’s weight. Steel or Kevlar are possible material options.

Intermediate Platform Mass Tether ~8kg (8km) Depot ~7kg Table 7-15 Intermediate platform

The depot platform is an important asset to the aerobot-rover team, it is the bridge that brings the aerobot and rovers together and makes it possible to realise the delivery and retrieval strategy.

Overall, the scientific instrumentation chosen is capable of satisfying the scientific requirements. It will study surface as well as atmospheric compositions, meteorology, geology, mineralogy, topography, subsurface structure, water trail and biological chemicals.

75 8 Conclusion and Future Work

The study suggested that a robotic airship and nanorovers’ collaborating will provide a strategic alternative for Mars planetary exploration.

A proposed architecture (airship + tethered depot + nanorovers) has combined the aerial platform and surface platform together. This combination satisfies the objectives of investigating a wide area and multiple in-situ samplings in a given piece of Martian terrain.

In terms of cooperative behaviours between the aerobot and rovers, the aerobot has a principal role in the team, its higher intelligence and ability is essential to the realization of the collaboration. The team utilises the aerobot for navigation and transport of the rovers and in turn make use of the rovers to complement the aerobot’s remote sensing data to get the most value out of the terrain target.

One significant area where substantial technology development is needed is in autonomy. The navigation, station keeping tasks require a formidable capability in computing and data fusion. However, the principal development lies in software rather than hardware.

This exploring approach by airship-nanorover team is suitable but not limited by this specific mission requirement or target. In general, it has offered an excellent opportunity for distributed scientific data collection, especially to locations characterized hazardous to rovers. Due to the reason that the airship does not require any energy to remain aloft, but only for active manoeuvrings, it has got great potential for extended mission durations. If taking the advantage of wind-powered long-distance passive flight combined with self-powered active manoeuvring, a global remote and in-situ investigation is possible. In the team, the airship has the highest degree of independence, its performance is less restricted by the rover platform than the other way round. A failure of the aerobot virtually means the end of the mission. While, a failure of any rover will reduce the scientific return but not terminate the whole mission. Life limitation on the mission is most likely due to gas leakage, or envelope or propulsion failure of the airship.

The airship design eliminated the ballonet system considering the complexity in airship landing. Employing the final configuration, given a 40kg payload provision, at a finesse ratio of 4, the solar powered airship resulted in a .3 15104 m3 volume and a gross mass of 332kg. It can cruise at an altitude of 5km with a forward velocity of 5m/s, and cover a range of 1100km in the duration of two and a half sols with power consumption of 1.23kW. The collaboration architecture is best suited to large missions, where an airship mass below 200kg is deemed unfeasible.

The instruments baselined are capable of gathering complementary information from air and ground, which enable the greatest amount of information to be gathered from any one target. The distributed instrumentation between rovers increases each one’s independence and reduces the likelihood of the whole surface platform from failing.

76 With complementary capabilities, the airship and rovers working in concert enhance each other’s ability and mission value as a whole.

There are many issues that could not be covered by this thesis, due to the scope and objective of the study. However, there is much work that could be stemmed from this investigation. They are listed below: 1. Investigating a refined design of each element and its subsystems. 2. Developing the software to realise the autonomous collaboration strategies. 3. Simulating the airship-rover team performances in the Martian atmosphere by the use of circulation models. …

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82 APPENDIX

Software Validation

Functionality: software performs airship sizing and parameters calculation:

Input: Mtot total mass (kg) h flight altitude (km) f finesse ratio V flight velocity (m/s) Table A Software Input Output: 3 Rho1 H altitude atmospheric density ( kg / m ) T1 H altitude temperature (K) p1 H altitude pressure (Pa) Mh2 internal gas mass (kg) 3 Vas airship envelope volume ( m ) Menvelope envelope material mass (kg) Mfin fin mass (kg) P power required (W) 2 Sa solar array area ( m ) Msa solar array mass (kg) Mrfcs regenerative fuel cells mass (kg) Mp electric propulsion system mass (kg) Mpayload payload capacity (kg) Table B Software Output Test Results Inputs Mtot = Mtot = Mtot = Mtot = 280 440 280 280 h= h= h= h= 5 5 6 7 f= f= f= f= 4 4 4 4 V= V= V= V= 15 15 15 15 Results Rho1 = Rho1 = Rho1 = Rho1 = 0.01 0.01 0.0093 0.0086 T1 = T1 = T1 = T1 = 222.0156 222.0156 219.0278 215.2556 p1 = p1 = p1 = p1 = 403.0232 403.0232 368.9722 336.7503

83 Mh2 = Mh2 = Mh2 = Mh2 = 5.8524 9.1966 5.8524 5.8524 Vas = Vas = Vas = Vas = 2.66E+04 4.18E+04 2.87E+04 3.09E+04 Menvelope = Menvelope = Menvelope = Menvelope = 110.2003 148.9518 115.8291 121.6871 Mfin = Mfin = Mfin = Mfin = 7.722 12.1346 8.3212 8.9603 P = P = P = P = 1.11E+03 1.45E+03 1.09E+03 1.07E+03 Sa = Sa = Sa = Sa = 189.2644 246.8632 185.5402 181.9812 Msa = Msa = Msa = Msa = 22.7117 29.6236 22.2648 21.8377 Mrfcs = Mrfcs = Mrfcs = Mrfcs = 48.9701 63.8732 48.0065 47.0857 Mp = Mp = Mp = Mp = 2.5084 3.29 2.4644 2.4225 Mpayload = Mpayload = Mpayload = Mpayload = 16.5805 92.9302 11.8071 6.6997 Inputs Mtot = Mtot = Mtot = Mtot = 440 440 440 440 h= h= h= h= 5 5 5 5 f= f= f= f= 3 5 4 4 V= V= V= V= 15 15 14 16 Results Rho1 = Rho1 = Rho1 = Rho1 = 0.01 0.01 0.01 0.01 T1 = T1 = T1 = T1 = 222.0156 222.0156 222.0156 222.0156 p1 = p1 = p1 = p1 = 403.0232 403.0232 403.0232 403.0232 Mh2 = Mh2 = Mh2 = Mh2 = 9.1966 9.1966 9.1966 9.1966 Vas = Vas = Vas = Vas = 4.18E+04 4.18E+04 4.18E+04 4.18E+04 Menvelope = Menvelope = Menvelope = Menvelope = 137.6513 159.0926 148.9518 148.9518 Mfin = Mfin = Mfin = Mfin = 12.1346 12.1346 12.1346 12.1346 P = P = P = P = 1.82E+03 1.40E+03 1.18E+03 1.75E+03 Sa = Sa = Sa = Sa = 308.5667 237.4523 199.8774 297.2584 Msa = Msa = Msa = Msa = 37.028 28.4943 23.9853 35.671 Mrfcs = Mrfcs = Mrfcs = Mrfcs = 79.8383 61.4382 51.7161 76.9124

84 Mp = Mp = Mp = Mp = 4.1333 3.1619 2.676 3.9499 Mpayload = Mpayload = Mpayload = Mpayload = 80.0178 86.4818 111.3395 73.1836 Table C Software Test Results

Tests showed that software had been implemented correctly, function satisfied the requirements. With changing input parameters, results change as expected.

Source Code

Solar powered airship

% Solar powered airship calculator % Hydrogen gas as internal gas format compact n=400; % cycle times M_Mtot=zeros(n); % total mass M_Mpayload=zeros(n);% payload mass Vas_Vas=zeros(n); for i=1:n % loop

%Input

Mtot=100+i % total mass with 10% margin h=5; % altitude in km f=4; % finesse ratio V=10+5; % airship forward speed 5m/s, average wind velocity 10m/s

%Initialization p0=6.36*10^2; % pressure at ground level Rho0=0.014694; % atmosphere density at ground level T0=238.74; % temperature at ground level

% parameters at H altitude level Rho1=0.014694-0.001145*h+4.6638*10^(-5)*h^2-9.7737*10^(-7)*h^3; % H altitude density T1=238.74-34.488*h+35.133*h^2-15.96*h^3+3.7315*h^4- 0.47352*h^5+0.030962*h^6-0.000817*h^7; % H altitude temperature p1=p0*Rho1*T1/(Rho0*T0); % H altitude pressure

% hydrogen gas characteristic RAMh2=1*10^-3; % relative atomic mass

85 Rh2=4124; % Rh2 spicific molar gas constant Rh2u=8.248; % Rh2 universal molar gas constant

% airship internal gas mass and volume calculation Mas=Mtot/1.1; % stringent total mass with no margin Rhoh2=p1/(Rh2*T1); % internal gas density at H altitude Mh2=Mas*RAMh2*p1/(Rh2u*T1*(Rho1-Rhoh2)) % mass of internal gas Vh2=Mh2*Rh2u*T1/(RAMh2*p1) % volume of internal gas Vas=Vh2 % volume of airship envelope L=(3*Vas*f^2/(4*pi))^(1/3); % length of the elliptical envelope - according to ellipsoid formula W=L/f; % width of the elliptical envelope

% airship envelope and fin calculation ArealDensity=0.02; % planetary aerobot composite material areal density 20g/m2 oc=acos(W/L); % variable for the calculation of ellipsoid area Area=2*pi*(W^2+L^2*(oc/tan(oc))) % area of envelope Menvelope=ArealDensity*Area % mass of envelope

Rf=0.0121; % fin area to airship volume coefficient Mfin=Rf*Vas*ArealDensity*1.2 % fin mass

% speed of sound and propeller tip Mach number gamma=1.29; % Martian atmosphere specific heat ratio Rair=191.8; % Martian atmosphere specific gas constant J/kg/K a=(gamma*Rair*T1)^(1/2); % speed of sound at H altitude M=0.9; % propeller tip Mach number

%electric propulsion system efficiencies J=V*pi/((a*M)^2-V^2)^(1/2) % advance ratio ct=-0.012122+0.14577*J-0.1408*J^2+0.05374*J^3-0.0068444*J^4; % thrust coefficient cp=-0.012752+0.094954*J-0.053694*J^2+0.017534*J^3-0.0007872*J^4; % power coefficient Etaprop=ct*J/cp % propeller efficiency Etamc=0.98; % motor controller efficiency Etaem=0.90; % electric motor efficiency Etag=0.90; % gearbox efficiency Etap=Etamc*Etaem*Etag*Etaprop; % propulsion drive train efficiency

% Drag and Power Cdv=0.23175-0.15757*f+0.04744*f^2-7.0412E-3*f^3+5.1534E-4*f^4-1.4835E-5*f^5; % volumetric grag coefficient D=0.5*Rho1*V^2*Cdv*Vas^(2/3); % Drag P=D*V/Etap+150 % communication 50W, control operation 50W, payload 50W

% solar panel sizing

86 Rhosa=0.12; % thin film density Iam=589; % mean solar intensity tau=0.427; % solar attenuation factor Etasc=0.08; % solar cell efficiency Sa=2*P/(0.5*Iam*tau*Etasc) % double the power for half sol day and half sol night Msa=Rhosa*Sa % mass of solar array % battery %DoD=0.4; % DoD 40% %Etab=0.9; % efficiency %Cr=150*12/(DoD*Etab); % battery capacity %Energydensity=90; % energy density of Lithium-Ion battery 90Whr/kg %Mb=Cr/Energydensity % mass of battery % RFCS Mrfcs=P*12.3/(400*0.70) % regenerative fuel cells mass

% propulsion system sizing Mem=P*Etamc/1291; % mass of electric motor if Mem<0.5 Mem=0.5; end Mmc=P/6223; % mass of motor cotroller if Mmc<0.1 Mmc=0.1; end Mg=P*Etaem*Etamc/3278; % mass of gearbox if Mg<0.3 Mg=0.3; end Mpc=P/1000; % mass of power conditioning system if Mpc<0.2 Mpc=0.2; end % propeller calculation dp=(D*pi^2/(ct*(a*M)^2-V^2))^0.5; % propeller blade diametre Vprop=9.25739E-5*dp^3; % propeller blade volume Mprop=1380*2*(1-0.5)*Vprop; % propeller mass

Mp=Mem+Mmc+Mg+Mpc+Mprop % mass of electric propulsion system

%------Msys=40; Mpayload=Mas-Mh2-Menvelope-Mfin-Mp-Msa-Mrfcs-Msys

% save into matrix for plotting M_Mtot(i)=Mtot; M_Mpayload(i)=Mpayload; Vas_Vas(i)=Vas; end

87 % plot diagram plot(M_Mtot,M_Mpayload,'b'), xlabel('Airship Total Mass - kg'), ylabel('Payload Mass - kg'), title('Effect of Payload Mass on Airship Mass') hold on

RTG powered airship

% RPS powered airship calculator % Hydrogen gas as internal gas format compact n=600; % cycle times M_Mtot=zeros(n); % total mass M_Mpayload=zeros(n);% payload mass Vas_Vas=zeros(n); for i=1:n % loop

%Input

Mtot=300+i % total mass with 10% margin h=5; % altitude in km f=4.4; % finesse ratio V=10+5; % airship forward speed 5m/s, average wind velocity 10m/s

%Initialization p0=6.36*10^2; % pressure at ground level Rho0=0.014694; % atmosphere density at ground level T0=238.74; % temperature at ground level

% parameters at H altitude level Rho1=0.014694-0.001145*h+4.6638*10^(-5)*h^2-9.7737*10^(-7)*h^3; % H altitude density T1=238.74-34.488*h+35.133*h^2-15.96*h^3+3.7315*h^4- 0.47352*h^5+0.030962*h^6-0.000817*h^7; % H altitude temperature p1=p0*Rho1*T1/(Rho0*T0); % H altitude pressure

% hydrogen gas characteristic RAMh2=1*10^-3; % relative atomic mass Rh2=4124; % Rh2 spicific molar gas constant Rh2u=8.248; % Rh2 universal molar gas constant

% airship internal gas mass and volume calculation Mas=Mtot/1.1; % stringent total mass with no margin Rhoh2=p1/(Rh2*T1); % internal gas density at H altitude

88 Mh2=Mas*RAMh2*p1/(Rh2u*T1*(Rho1-Rhoh2)) % mass of internal gas Vh2=Mh2*Rh2u*T1/(RAMh2*p1) % volume of internal gas Vas=Vh2 % volume of airship envelope L=(3*Vas*f^2/(4*pi))^(1/3); % length of the elliptical envelope - according to ellipsoid formula W=L/f; % width of the elliptical envelope

% airship envelope and fin calculation ArealDensity=0.02; % planetary aerobot composite material areal density 20g/m2 oc=acos(W/L); % variable for the calculation of ellipsoid area Area=2*pi*(W^2+L^2*(oc/tan(oc))); % area of envelope Menvelope=ArealDensity*Area; % mass of envelope

Rf=0.0121; % fin area to airship volume coefficient Mfin=Rf*Vas*ArealDensity*1.2 % fin mass

% speed of sound and propeller tip Mach number gamma=1.29; % Martian atmosphere specific heat ratio Rair=191.8; % Martian atmosphere specific gas constant J/kg/K a=(gamma*Rair*T1)^(1/2); % speed of sound at H altitude M=0.9; % propeller tip Mach number

%electric propulsion system efficiencies J=V*pi/((a*M)^2-V^2)^(1/2) % advance ratio ct=-0.012122+0.14577*J-0.1408*J^2+0.05374*J^3-0.0068444*J^4 % thrust coefficient cp=-0.012752+0.094954*J-0.053694*J^2+0.017534*J^3-0.0007872*J^4 % power coefficient Etaprop=ct*J/cp % propeller efficiency Etamc=0.98; % motor controller efficiency Etaem=0.90; % electric motor efficiency Etag=0.90; % gearbox efficiency Etap=Etamc*Etaem*Etag*Etaprop; % propulsion drive train efficiency

% Drag and Power Cdv=0.23175-0.15757*f+0.04744*f^2-7.0412E-3*f^3+5.1534E-4*f^4-1.4835E-5*f^5; % volumetric grag coefficient D=0.5*Rho1*V^2*Cdv*Vas^(2/3); % Drag P=D*V/Etap+150 % communication 50W, control operation 50W, payload 50W Prps=7 % specific power of Stirling Radioisotope Generator 7W/kg Mrps=P/Prps % RPS mass

% propulsion system sizing Mem=P*Etamc/1291 % mass of electric motor if Mem<0.5 Mem=0.5; end Mmc=P/6223 % mass of motor cotroller

89 if Mmc<0.1 Mmc=0.1; end Mg=P*Etaem*Etamc/3278 % mass of gearbox if Mg<0.3 Mg=0.3; end Mpc=P/1000 % mass of power conditioning system if Mpc<0.2 Mpc=0.2; end % propeller calculation dp=(D*pi^2/(ct*(a*M)^2-V^2))^0.5; % propeller blade diametre Vprop=9.25739E-5*dp^3; % propeller blade volume Mprop=1380*2*(1-0.5)*Vprop; % propeller mass

Mp=Mem+Mmc+Mg+Mpc+Mprop % mass of electric propulsion system

%------Msys=40; Mpayload=Mas-Mh2-Menvelope-Mfin-Mp-Mrps-Msys

% save into matrix for plotting M_Mtot(i)=Mtot; M_Mpayload(i)=Mpayload; Vas_Vas(i)=Vas; end

% plot diagram plot(M_Mtot,M_Mpayload,'b'), xlabel('Airship Mass - kg'), ylabel('Payload Mass - kg'), title('Effect of Payload Mass on Airship Mass') hold on

90

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