Comparative Study of Aerial Platforms for Mars Exploration

Thesis by

Nasreen Dhanji

Department of Mechanical Engineering McGill University Montreal, Canada

October 2007

A Thesis submitted to McGill University in partial fulfillment of the requirements for the degree of Master of Engineering

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While these forms may be included Bien que ces formulaires in the document page count, aient inclus dans la pagination, their removal does not represent il n'y aura aucun contenu manquant. any loss of content from the thesis. Canada Abstract

The primary objective of this thesis is to develop a framework to compare the performance of fixed-wing aircraft, airships, and the rotary-wing aircraft in the Martian environment and through that framework, determine which of these platforms is best suited to conduct a series of scientific investigations on Mars. Three Mars mission scenarios provide the context within which the performance of the platforms is evaluated. The mission scenarios are used to derive the performance requirements including the range and flight path to be covered, the altitude at which the platforms fly, and the scientific investigations to be performed along with the associated scientific instrumentation to be carried as payload. Existing platform designs are used for the purpose of this comparative study and are modified depending on specific mission requirements. A set of weighted performance metrics, including the gross takeoff mass, power required, manoeuvrability, and complexity, serves as a common basis for comparing the performance of the three aerial platforms. The results of this comparative study indicate that the airship is best suited for all mission scenarios considered due to its simplicity and high degree of manoeuvrability. However, it is important to note that a series of subjective design choices with respect to platform speed and available power were made that significantly impact the overall performance of the platforms. Altering these design choices as well as the mission requirements could result in a different platform being best suited for each Mars mission. For instance, increasing the cruising velocity of the fixed-wing aircraft may allow its dimensions to be scaled down thereby reducing the complexity and making it a more competitive platform for long-range missions. In addition, for short-range missions that do not require a high degree of manoeuvrability but where the gross takeoff mass and complexity are more important factors, the rotary-wing aircraft becomes the best option.

1 Resume

La presente these a pour objectif principal de concevoir un cadre de comparaison portant sur la performance de l'aeronef a voilure fixe, le dirigeable et de l'aeronef a voilure tournante dans Penvironnement martien, grace auquel il sera possible de determiner laquelle de ces plateformes aeriennes convient le mieux pour effectuer une serie de recherches scientifiques sur Mars. Trois scenarios de mission fournissent le contexte dans lequel la performance des plateformes est evaluee. Les scenarios servent a determiner les criteres de performance y compris la distance franchissable et la trajectoire de vol a parcourir, 1'altitude de vol des plateformes et les etudes scientifiques a mener ainsi que le materiel scientifique a transporter comme charge utile. Des modeles existants de plateforme sont utilises aux fins de la presente etude comparative et sont modifies selon les exigences specifiques de chaque mission. Une serie de parametres ponderes, comprenant la masse au decollage, la puissance necessaire, la manoeuvrabilite et le niveau de complexite, sert de commune mesure pour la comparaison de la performance des trois plateformes aeriennes. Les resultats de la presente etude comparative demontrent que, compte tenu de sa simplicity et de sa grande manoeuvrabilite, le dirigeable convient le mieux a tous les scenarios de mission considered. Toutefois, il est important de noter que des choix subjectifs de modeles, ayant une incidence considerable sur la performance globale des plateformes, ont ete effectues relativement a la vitesse des plateformes et a la puissance disponible. Le fait de modifier ces choix de modeles ainsi que les exigences de mission reviendrait a opter pour une plateforme differente convenant le mieux a chaque mission sur Mars. Par exemple, l'augmentation de la vitesse de croisiere de l'aeronef a voilure fixe permettrait la diminution de ses dimensions, reduisant ainsi le niveau de complexite et rendant la plateforme plus appropriee pour les missions lointaines. En outre, en ce qui concerne les missions a courte distance n'exigeant pas une grande manoeuvrabilite, mais pour lesquelles la masse au decollage et le niveau de complexite sont des facteurs plus importants, l'aeronef a voilure tournante devient la meilleure option.

u Acknowledgements

I would like to extend my sincere gratitude to my thesis advisor, Professor Meyer Nahon, for his support and guidance. He always provided good direction when required and was a great source of information. In addition, I would like to extend my sincere gratitude to my thesis co-advisor, Dr. Erick Dupuis of the Canadian Space Agency. Erick played an instrumental role in determining the scientific investigations that could be performed on Mars using aerial platforms. He also provided helpful referrals to experts in the field of Mars exploration at the Canadian Space Agency.

I would like to thank Alain Berinstain of the Canadian Space Agency for his help in identifying regions on Mars that are of scientific interest with respect to the Mars missions developed in this study. I would also like to thank Victoria Hipkin of the Canadian Space Agency for her help in identifying the type of scientific instruments that can be used to perform the specific scientific investigations on Mars.

Finally I would like to thank Yuwen Li, a fellow Master's student at McGill University, for his help in determining the yaw rate associated with the airship as well as the stable airship configuration.

in Table of Contents

Abstract i

Resume ii

Acknowledgements iii

Table of Contents iv

List of Figures viii

List of Tables x

Nomenclature xi

CHAPTER 1: INTRODUCTION AND LITERATURE REVIEW 1

1.1 Scientific Investigations on Mars 1

1.2 Case for Aerial Missions on Mars 2

1.3 Scientific Objectives Achievable Using Aerial Platforms 3

1.4 Literature Review 5

1.5 Obj ectives of Thesis 17

CHAPTER 2: MISSION CONCEPTS FOR MARS EXPLORATION 19

2.1 Determining if Life Ever Arose on Mars 21

2.2 Understanding the Processes and History of Climate on Mars 26

2.3 Determining the Evolution of the Surface and Interior of Mars 31

iv CHAPTER 3: AERIAL PLATFORMS FOR MARS EXPLORATION 36

3.1 Lighter-Than-Air Vehicles 36

3.2 Heavier-Than-Air Vehicles 39

3.3 Aerial Platforms Considered in this Study 40

3.3.1 Fixed-Wing Aircraft Configuration 41

3.3.2 Airship Configuration 46

3.3.3 Rotary-Wing Aircraft Configuration 49

CHAPTER 4: AERIAL PLATFORM PERFORMANCE ON MARS 52

4.1 Performance Metrics 53

4.2 Fixed-Wing Aircraft 55

4.2.1 Aircraft Characteristics 55

4.2.2 Method of Evaluation of Performance Metrics for Fixed-

Wing Aircraft 57

4.2.2.1 Gross Takeoff Mass 57

4.2.2.2 Power 61

4.2.2.3 Manoeuvrability 62

4.2.2.4 Complexity 65

4.2.3 Scaling of the Fixed-Wing Aircraft 66

4.3 Airship 67

4.3.1 Method of Evaluation of Performance Metrics for the Airship 68

4.3.1.1 Gross Takeoff Mass 68

4.3.1.2 Power 72

v 4.3.1.3 Manoeuvrability 73

4.3.1.4 Complexity 75

4.3.2 Scaling of the Airship 75

4.4 Rotary-Wing Aircraft 76

4.4.1 Method of Evaluation of Performance Metrics for the Rotary-

Wing Aircraft 76

4.4.1.1 Gross Takeoff Mass 76

4.4.1.2 Power 80

4.4.1.3 Manoeuvrability 85

4.4.1.4 Complexity 87

4.4.2 Scaling of the Rotary-Wing Aircraft 88

4.5 Method for Evaluating a Single Performance Rating for each Aerial

Platform 89

4.5.1 Performance Rating with respect to Mass {Qx) 91

4.5.2 Performance Rating with respect to Power (Q2) 91

4.5.3 Performance Rating with respect to Manoeuvrability (Q3) 92

4.5.3.1 Performance Rating with respect to Turn Radius (q{) 92

4.5.3.2 Performance Rating with respect to Range of Speed (q2) 92

4.5.3.3 Performance Rating with respect to Ability to Hover (q3) 93

4.5.4 Performance Rating with respect to Complexity (Q4) 93

4.5.5 Weighting of Performance Metrics 94

4.5.5.1 Mission 1 94

4.5.5.2 Mission 2 95

vi 4.5.5.3 Mission 3 96

CHAPTER 5: DISCUSSION OF RESULTS 97

5.1 Fixed-Wing Aircraft 97

5.2 Airship 101

5.3 Rotary-Wing Aircraft 104

5.4 Summary of Results 107

5.5 Performance Rating of Aerial Platforms 108

5.6 Factors Affecting the Overall Performance of the Aerial Platforms 110

CHAPTER 6: CONCLUSIONS AND FUTURE WORK 114

6.1 Conclusions 114

6.2 Future Work 118

REFERENCES 119

APPENDIX A: Previous Mars Missions 131

APPENDIX B: Unconventional Aerial Platforms 13 8

APPENDIX C: Fixed-Wing Aircraft Characteristics 141

APPENDIX D: Airship Characteristics 144

vn List of Figures

Figure 1.1 Marsplane Configuration 7 Figure 1.2 Zephyr Configuration 8 Figure 1.3 ARES Scout Airplane 9 Figure 1.4 Hybrid Airship 12 Figure 1.5 Photovoltaic Rotorcraft 14 Figure 1.6 Harris/DeLaurier Model of Engine-Powered Piloted Ornithopter 16 Figure 2.1 Viking Derivative Aeroshell 20 Figure 2.2 Chryse Planitia - Site for Conducting Mission 1 25 Figure 2.3 Chasma Boreale - Site for Conducting Mission 2 30 Figure 2.4 Valles Marineris - Site for Conducting Mission 3 35 Figure 3.1 Balloon Deployment Sequence 38 Figure 3.2 Gossamer Albatross 42 Figure 3.3 Daedalus 88 43 Figure 3.4 Daedalus 88 Configuration 44 Figure 3.5 Robotic Martian Airship 47 Figure 3.6 Martian Autonomous Rotary-Wing Vehicle Configuration 50 Figure 3.7 MARV Body Configuration 50 Figure 4.1 Flow Chart for Iterative Calculation of Fixed-Wing Gross Takeoff Mass 60 Figure 4.2: Flow Chart for Iterative Calculation of Airship Gross Takeoff Mass 69 Figure 4.3: Airship in Steady Turn 74 Figure 4.4: Flow Chart for Iterative Calculation of Rotary-Wing Gross Takeoff Mass 77

Figure 4.5 Power vs. Speed Profile for Rotary-Wing Aircraft 84 Figure 5.1 Originally Scaled Airships (Unstable Configuration) 102 Figure 5.2 Airships with Enlarged Fin and Rudder (Stable Configuration) 102

via Figure 5.3: Turn Rates at Maximum Speed and Maximum Rudder Deflection (£ = 30°) 103 Figure 5.4: Power vs. Speed Profile (Mission 1) 104 Figure A. 1: Mariner 4 Spacecraft 132 Figure A.2: Mars Climate Orbiter 133 Figure A.3: Mars Reconnaissance Orbiter (MRO) 135 Figure A.4: Mars Pathfinder 136 List of Tables

Table 2.1: Requirements for Mission 1 26 Table 2.2: Requirements for Mission 2 31 Table 2.3: Requirements for Mission 3 35 Table 3.1: Characteristics of Daedalus 88 in the Earth Environment 45 Table 3.2: Specifications of the Robotic Martian Airship 48 Table 3.3: RMA Mass Summary 48 Table 3.4: MARV Rotor Specifications 51 Table 3.5: MARV Mass Breakdown 51 Table 4.1: Atmospheric Properties of Mars and Earth at Datum 53 Table 4.2: Weighting of Performance Metrics for each Mission 94 Table 5.1: Common Fixed-Wing Aircraft Characteristics 98 Table 5.2: Specific Fixed-Wing Aircraft Characteristics 99 Table 5.3: Manoeuvrability Factors for Fixed-Wing Aircraft 100 Table 5.4: Complexity Factors for Fixed-Wing Aircraft 100 Table 5.5: Airship Characteristics 101 Table 5.6: Manoeuvrability Factors for Airship 103 Table 5.7: Rotary-Wing Aircraft Characteristics 105 Table 5.8: Manoeuvrability Factors for Rotary-Wing Aircraft 106 Table 5.9: Complexity Factors for Rotary-Wing Aircraft 106 Table B.l: Characteristics of DeLaurier's Proof of Concept Ornithopter 139 Table B.2: Characteristics of DeLaurier's Full-Scale Piloted Ornithopter 140 Table D. 1: Density of Lifting Gas and Atmosphere for each Mission 145

x NOMENCLATURE a Angle of Attack

A Area

A Rotor Disk Area

AR Aspect Ratio b Span

B Buoyancy c Maximum Chord Length c Mean Chord Length

rotor Rotor Chord Length

CD Drag Coefficient

Cn Minimum Drag Coefficient

CL Lift Coefficient

C, Minimum Drag Lift Coefficient LMD D

C, Maximum Lift Coefficient

CD Parasite Drag Coefficient

CD Induced Drag Coefficient

Cf Skin Friction Coefficient

CT Thrust Coefficient

S Rudder Deflection

XI d Maximum Airship Diameter diat Internal Aeroshell Diameter

d0 Fuselage Diameter

dmtor Rotor Diameter

D Drag

£ Eccentricity e Oswald's Efficiency Factor f Fuel Consumption

F Flat Plate Area g Martian Gravity k Constant of Proportionality

K Drag-due-to-Lift Factor

I Characteristic Length

L Lift

LHVH2 Hydrogen Lower Heating Value m Mass rjfc Fuel Cell Thermodynamic Efficiency rlprop Propeller Efficiency

chord Chord-wise Folds

niength Length-wise Folds n Span-wise Folds span

Xll N Load Factor

0 Bank Angle a Rotor Solidity p Density

pco Density of Carbon dioxide

pH Density of Hydrogen

P Pressure

PAmil Available Power

Plnd Induced Power

PPam Parasite Power

PVro Profile Power

PKeq Required Power

Pspec Specific Power

Pt Power Required for Turn q Performance Rating (with respect to performance metric)

Q Platform Performance Rating r Turn Radius

R Range

R Gas Constant

Re Reynolds Number

S Planform Area

xin Swel Wetted Area

t Mission Duration

T Thrust u Airship Velocity in the X Direction

// Advance Ratio

v Airship Volume

vH Hydrogen Volume

v Kinematic Fluid Viscosity v Airship Velocity in the Y Direction

V Forward Platform Velocity

Vyp Minimum Drag Velocity

vRange Range of Speed

Vs Stall Speed

Vt Speed in the Turn

VT Rotor Tip Speed

w Downwash Velocity w Weight Assigned to Influencing Factor

W Weight Assigned to Performance Metric

x X-Coordinate of Airship's Center of Gravity

xiv Chapter 1

INTRODUCTION AND LITERATURE REVIEW

1.1 Scientific Investigations on Mars

For centuries, exploration of Mars has been of significant interest to mankind. In recent decades, this interest has heightened with new evidence indicating that Mars may once have had flowing water as well as lakes and oceans. Some of that water is now frozen at the poles but data from recent Mars missions indicates that a substantial portion of water may be buried beneath Mars' surface. The presence of water on Mars also raises questions about whether life could have arisen, but more importantly, whether life could still persist in these regions of Mars believed to contain a large portion of subsurface water and ice. The discovery of water on Mars would strengthen support for human missions to Mars and pave the way towards establishing human colonies on Mars.

So far, the exploration of Mars has occurred in three stages: flyby missions, orbiter missions and landers and rovers [1]. Appendix A provides a brief account of these previous missions. Although orbital and landed packages have provided many of the high priority scientific measurements, alternative platforms are required to access areas of Mars not currently accessible by orbiters and landers. Future exploration of Mars using aerial platforms (such as airplanes, rotorcraft, and balloons), subsurface explorers and sample return missions, is the next step towards studying Mars from a perspective never achieved before.

The use of aerial platforms would allow us to gather higher resolution data and surface images of Mars than orbital platforms have yielded. Aerial platforms also have the ability

1 to cover more territory than current rover technologies allow. Furthermore, aerial platforms have been used for Earth-based applications for over a century; therefore, the technology to develop aerial platforms already exits and can be easily adapted for Mars applications. Some aerial platforms have already been tested on other planets such as the balloons used on the Venus Vega mission [2].

Subsurface explorers would allow us to determine whether reservoirs of water, in either liquid or frozen form, exist in the Martian subsurface. However, the technology required to perform subsurface exploration is still under development. There is also a strong consensus on the need for sample return missions. Sample return missions would allow us to conduct exhaustive studies of rock, soil, and atmospheric particles that can only be conducted in laboratories here on Earth. However, bringing samples back is challenging and requires the ability to determine which samples are the most scientifically interesting.

Therefore, aerial missions provide the most logical means of accessing areas of Mars that are currently inaccessible by orbiters and rovers and will play an important role in setting the stage for future subsurface and sample return missions.

1.2 Case for Aerial Missions on Mars

Scientific surveys of Mars have been conducted on the macro-scale by satellites and on the micro-scale by rovers. However, there remains a huge gap between the spatial resolution and coverage that can be achieved from orbital spacecraft and surface vehicles. Detailed exploration of Mars requires the capability to survey the planet with a spatial resolution that cannot be achieved with orbiting spacecraft. Surface rovers have limited range because of the rough terrain that must be traversed. Aerial platforms, therefore, provide a means of bridging the gap between orbital and ground based investigations of Mars.

Whereas orbiting platforms provide a resolution of tens of meters, the spatial resolution from an aerial platform is several orders of magnitude higher due to its ability to fly at

2 low altitudes of a few hundred meters to a few kilometers above the surface. Aerial platforms enable ultra-high resolution imaging (20 cm resolution or less) over extensive areas of the planet enabling recognition of individual rocks and identification of areas suitable for sample acquisition [15]. High-resolution imaging is a key factor in interpreting the geologic history of Mars as well as the processes (aeolian, hydrothermal, aqueous, volcanic, cratering, and tectonic) leading to its evolution.

Furthermore, both the spatial resolution and the signal strength for magnetic observations improve dramatically from the height of an aerial platform. Airborne radar sounders have also proven to be powerful tools in detecting subsurface water up to a depth of 4 km. In situ measurements from aerial platforms can, therefore, be used to validate global remote sensing data from orbiting platforms.

Aerial platforms will play an important role in early robotic missions dedicated to exploring and characterizing potential landing sites for sample return and human missions. In addition, aerial platforms will provide useful information on atmospheric parameters and variations that affect atmospheric flight. This in turn will assist in the development of planetary mobility systems such as tele-operated aerial platforms and aerial transportation during the phase of active human exploration.

1.3 Scientific Objectives Achievable Using Aerial Platforms

In order to obtain science input for planning and prioritizing Mars exploration activities for the next several decades, NASA has formed a community-based forum entitled Mars Exploration Program Analysis Group (MEPAG). MEPAG has set out various scientific goals and objectives to address the types of scientific investigations that are of interest to scientists conducting research on Mars.

Through a series of workshops, meetings and online surveys, MEPAG has solicited inputs from the wider scientific community and has documented the detailed scientific goals, objectives and investigations for Mars exploration in the "MEPAG Goals,

3 Objectives, Investigations and Priorities" report [3]. The MEPAG report has become the standard basis for evaluating prospective missions to Mars.

The four major goals for Mars exploration as documented in the MEPAG report include: I. Determining if life ever arose on Mars II. Understanding the processes and history of climate on Mars III. Determining the evolution of the surface and interior of Mars IV. Preparing for human exploration

These four goals are interrelated and encompass the key areas of interest with respect to Mars exploration. Each goal has been further broken down into a set of scientific objectives aimed at gathering information that would provide insight into the factors influencing the high level goal. The objectives are in turn broken down into scientific investigations aimed at taking specific measurements in order to obtain the data required to understand the entire complex Mars system and how it operated through time.

For the purpose of this study, the MEPAG report has been used as a guideline for assessing the types of scientific missions that could be conducted using aerial platforms. In addition, an assessment was made of each scientific investigation listed in the report in order to determine the mission requirements for performing the investigation. The assessment included determining whether the measurements can be made remotely or whether they require direct contact with samples; the type of instrumentation that would be required to make the measurements and the associated constraints that would be imposed on the mission and platform as a result of the instrumentation; and the type of platform (orbital, lander or aerial) that would be best suited to perform each investigation. The detailed assessment is provided in Chapter 2.

As noted in the MEPAG report, a cursory reading of the goals, objectives and investigations show that several crucial technical capabilities need to be developed. "The most important of these are: (1) access to all of Mars - high and low latitudes, rough and smooth surfaces, low and high elevations, plus precision landing; (2) access to the

4 subsurface, from a meter to hundreds of meters, through a combination of drilling and geophysical sounding; (3) access to time varying phenomena; hence the need to be able to make climate studies; (4) access to microscopic scales by instruments capable of measuring chemical and isotopic compositions and determining mineralogy and the nature of mineral intergrowths" [3].

1.4 Literature Review

Various types of aerial platforms have been proposed for Mars exploration including fixed-wing aircraft, balloons and airships, and rotary-wing aircraft. In addition, hybrid aircraft such as winged and rotor airships as well as unconventional flapping-wing aircraft have also been proposed. Reference 4 discusses the role that such aerial platforms will play in the future exploration of Mars. The paper discusses a number of observational techniques that can be employed to great advantage from aerial platforms, the role that aerial platforms can play in sample selection through acquisition of high resolution morphological, mineralogical and topographic data, and the role of aerial platforms in searching for subsurface life through magnetic and radar sounding.

Specific designs for each type of aerial platform have also been developed. A review of the literature associated with the proposed designs for each of these types of aerial platforms is provided below.

Fixed-Wing Aircraft

A conceptual study of an airplane specifically designed for Mars flight was proposed by the Jet Propulsion Laboratory (JPL) along with its contractor, Development Sciences, Inc., in 1979 [5]. Per the proposed design, the airplane would fly at an altitude of up to 15 km above the Martian datum at a cruise speed of 60 — 100 m/s, covering a maximum range of 10,000 km and having a payload capacity of 40 - 100 kg.

5 Five experiment areas were chosen to evaluate the airplane's science capability with respect to aerial survey: visual imaging, gamma ray and infrared reflectance spectroscopy, gravity field, magnetic field and electromagnetic sounding, and atmosphere composition and dynamics. In designing the Mars airplane, the following issues were considered: the air density of Mars; the hazard that uncharted terrain elevation on a local scale presents for low flying missions; limited knowledge of winds and turbulence; the hazard that dust storms present; and the difficulty in landing due to rugged terrain.

Consideration of these issues led to a mission concept that employs 12 airplanes, delivered in three squadrons of four airplanes, rather than the usual two vehicles, like Viking. This concept was deemed plausible due to the low unit cost of each airplane, as well as a high crash rate in the face of the above-mentioned environmental issues. Each airplane would have a wingspan of 21 m, a mean chord of 0.655 m, a wing area of 20 m2, and a two bladed, 4.5 m diameter propeller. The wings would require six folds in order to fit within the Viking-like aeroshell having a diameter of 3.8 meters.

In 1984, a comparative study between the performance of an airplane, a balloon and a dirigible airship was conducted by MIT [6]. The study maintained that if airplanes were used on Mars, they would likely resemble the recent and highly efficient manpowered aircraft, such as the Gossamer Condor or Gossamer Albatross, due to the requirement for low wing loading and lightweight structures. Furthermore, manpowered aircraft on Earth operate at low Reynolds numbers due to the low speeds at which they fly. Fixed-wing aircraft on Mars are expected to operate under a similar Reynolds number regime as manpowered aircraft on Earth due to the low Martian atmospheric density. A comparison between three human-powered aircraft, Gossamer Albatross, Light Eagle and Daedalus 88, indicated that the Daedalus 88 had the best performance with respect to range covered on Earth [7]. With a wingspan of 34 m, a wing area of 31 m and a gross weight of 109.6 kg, the Daedalus 88 covered a range of 116.58 km flying at a speed of 6.87 m/s.

In 1988, a project was undertaken by the University of Illinois to develop designs for a manned Mars airplane [8]. Figure 1.1 illustrates the Mars airplane configuration. The

6 performance specifications for the Mars airplane included an endurance of eight hours while carrying two suited astronauts equipped with life support systems. The design drivers included (1) the low atmospheric density, (2) the need for lightweight materials (including fuel) and highly efficient propulsion systems, and (3) the take-off and landing requirements. The initial design parameters of the aircraft included a maximum gross weight of 7740 N, a maximum installed power of 155 kW, a wingspan of 85 m, an aspect ratio of 15 and a cruise speed of 75 m/s.

1 [ 111 V 1 1 1 1 ^ TOP I io.e» 1 1 1 Tw« •!«• Figure 1.1 Marsplane Configuration [8]

The design results showed that (1) propulsion systems using electric motor driven propellers powered by fuel cells or advanced technology batteries have values of specific power high enough for low-speed cruise and modest climb performance, (2) rocket based VTOL systems, based on existing technology, are light enough to be used to provide versatile take-off and landing performance, and (3) efficient structures using composite materials help to increase payload fractions.

The Zephyr, another manned Martian airplane, was proposed under a study conducted at the University of Toronto in 1996 [9]. Figure 1.2 illustrates the configuration of the

7 Zephyr. The design consists of a 2388 kg aircraft containing two suited astronauts. The wing aerofoil is a Selig 1223 high-lift, low Reynolds number design. The wing area is about 185.8 m2. The optimum propeller consists of 10 blades with a diameter of 1.8 m. The aircraft is powered by a hydrogen-oxygen fuel cell. It is envisioned that the fuel for the Zephyr would be produced in-situ by electrolyzing ground ice found on Mars.

•w-r •MB—

Figure 1.2 Zephyr Configuration [9]

Sverdrup Technology Inc. proposed a preliminary design for a long-endurance Mars aircraft, in 1990 [10]. The design considered both radioisotope/heat engine and PV solar array power production systems. Various cases for each power system were analyzed in order to determine the necessary size, weight and power requirements of the aircraft. The analysis was set up to design an aircraft that had minimum wingspan and maximum endurance. The results showed that a long endurance aircraft is feasible within the

8 Martian atmosphere. Furthermore, the size and weight of the most efficient solar aircraft were comparable to the one powered by the radioisotope/heat engine.

The Canyon Flyer, an airplane for Mars exploration, was proposed by the NASA Ames Research Center in 2000 [11]. The Canyon Flyer is a propeller-driven subsonic airplane having a wingspan of 1.61 m and a chord of 0.35 m. The aircraft has a gross mass of 14.6 kg, providing a payload capacity of 5.2 kg. The configuration of the Canyon Flyer requires the wings to be folded in order to fit within the atmospheric-entry aeroshell. Both battery-powered electric motors and hydrazine-fueled reciprocating engines were considered for power generation. Although the battery-powered electric motor provided sufficient endurance for the purpose of this mission, the continued development of hydrazine-powered motors was recommended for longer duration missions.

ARES, a Mars exploration airplane, was one of four Mars Scout proposals selected in December 2002 for continued study and refinement [12]. Figure 1.3 illustrates the ARES airplane. The airplane has a 6.25 m wingspan with a total mass of 149 kg and uses a bipropellant liquid rocket system for propulsion. The scientific goals include making atmospheric measurements to determine the boundary layer composition, chemistry and dynamics from an altitude of 1 to 2 km above ground level, an aeromagnetic survey of the Mars surface, and the search for near surface water.

.^W

Figure 1.3 ARES Scout Airplane [103]

9 Balloons and Airships

The first interplanetary application of lighter than air vehicles was the 1985 Soviet/French Venus Vega Mission [2]. Two balloons, fabricated from a woven Teflon cloth coated with Teflon film, were deployed on Venus as part of this mission [13]. Each balloon operated for about 46.5 hours and covered more than a third of the planet. The sensors on board the gondolas measured pressure, temperature, vertical wind speed, and cloud particle physical properties.

Due to the relatively low operating temperatures, the balloons were not capable of exploring the hotter, lower atmosphere of the planet. Instead, they maintained a fairly constant altitude of 54 km. The missions were terminated when the batteries used to operate the transmission equipment lost power.

Success of the VEGA Venus balloons led to a serious study of balloon applications on Mars. In 1987, a NASA-sponsored design course at Utah State University led to the design of a single hydrogen super-pressure balloon to study the surface and lower atmosphere of Mars [14]. The balloon, weighing 400 kg and having a diameter of 38 m, would be inflated during descent to avoid fabric damage through contact with the Martian surface. The balloon was designed to carry a payload of 116 kg to an altitude of 2.5 km above the Martian datum and remain aloft both night and day.

The Mars Aerial Platform Concept, conceived in 1995, proposed the use of super- pressure balloons deployed at a constant atmospheric density (nominally set at 6 km above the Mars datum) [15]. Three balloons having a nominal lifetime of 100 sols, each covering a range of more than 500,000 km, would be deployed in the northern, southern, and equatorial zones of Mars. The 18 m diameter super-pressure balloons, fabricated from 11-um biaxial nylon, would carry 9.3 kg gondolas.

In 1999, a novel hot-air balloon system known as the solar "Montgolfiere" was proposed by JPL [16]. Using entirely solar heat, JPL demonstrated that the hot-air balloon is

10 capable of landing and re-ascending using a novel, lightweight, radio-controlled top air vent. The study indicated that a solar Montgolfiere balloon carrying a 2 kg gondola would weigh just 4.4 kg and have a diameter of 12 .5 m, whereas a balloon carrying a 10 kg gondola would weigh 9.12 kg and have a diameter of 18 m.

At the same time, a low-cost Mars balloon mission was being studied at JPL for the 2001 Mars balloon mission opportunity [17]. The design parameters included the use of a constant density altitude super-pressure balloon system without landing capability; a 10 kg gondola with up to 3-4 kg of science instruments; and a northern hemisphere entry and flight.

Unlike balloons that drift freely in the atmosphere, airships incorporate some means of propulsion in order to control their trajectory. Furthermore, the design of the airship hull also contributes to the aerodynamic component of its lift. The use of control surfaces allows the flight path of airships to be controlled unlike balloons. The presence of propulsion systems, combined with the use of control surfaces, enable scientists to conduct more directed scientific missions with airships than with balloons.

In 1997, the design of a robotic Martian airship was presented at the AIAA Aerodynamic Decelerator Systems Technology Conference in San Francisco, California [18]. The 200 kg super-pressure airship would provide a payload capacity of 10 kg. 50 m in length and 25 m in diameter, it was designed to have an inflated volume of 16500 m3. The airship would travel at a maximum speed of 10 m/s covering a range of 1440 km.

In addition to conventional balloons and airships, various designs for hybrid airships have been proposed. Hybrid airships provide several advantages including reduced size and volume, since part of the lift would be generated aerodynamically, and the ability to control the direction of flight.

In 1984, a comparative study between the performance of an airplane, a balloon and a dirigible airship for Mars exploration, conducted at MIT [6], indicated that the dirigible

11 airship would provide the greatest payload capacity for a 1000 kg overall lift requirement. The study indicated that hybrid buoyant/aerodynamic craft seem to show optimum performance when the vehicle has ten to fifteen percent of its lift derived from its buoyancy and the remainder from aerodynamics.

The Federal University of Rio de Janeiro performed a feasibility study of a hybrid airship operating in ground effect for Earth-based applications in 1977 [19]. The design utilized a low aspect ratio wing to generate aerodynamic lift. The low aspect ratio wing also helped to reduce the structural weight of the wing and allowed operation in ground effect. The airship hull was reshaped from a conventional circular cross section into a rectangular cross section to reduce the projected side area (so as to decrease the gust loading problem) and to increase the hull lift/drag ratio when blended into the wing. Figure 1.4 illustrates the hybrid airship. Further study of the aerodynamic performance of delta- winged hybrid airships indicated that a rigid hybrid airship of fineness ratio 4, incorporating a slender delta wing of aspect ratio 1.5 had superior performance to modern conventional airship designs and existing jet airplanes [20].

Figure 1.4 Hybrid Airship [19]

These hybrid airship designs for Earth-based applications provided inspiration for a similar design for Mars exploration. In 1998, Aereon Corporation proposed a design for an unmanned solar-powered hybrid airship for Mars exploration [21]. The design

12 consisted of an inflatable hull with a low aspect ratio wing and VTOL capabilities allowing a ground launch. In addition, smart materials were incorporated in the propellers allowing them to change the propeller twist distribution in order to provide vertical take off capability. Finally, the shape of the hull provided a large area for solar collection thus making solar power a viable propulsion alternative.

Rotorcraft

Rotorcraft offer several advantages over the other types of aerial platforms including vertical lift capabilities, the ability to hover, low speed flight, and take off and landing at unprepared sites. Several studies have been conducted to determine the feasibility of using rotorcraft for Mars exploration.

In 1993, the Institute of Fluid Mechanics and Flight Dynamics in Romania proposed a design for an autonomous flying robot for Mars exploration [22]. The design consisted of two coaxial counter-rotating and foldable rotors for vertical lift. Solar cells, mounted on the upper side of the rotor blades, would be used to supply electricity to the electro-motor of a gas compressor [23]. Figure 1.5 illustrates the photovoltaic rotorcraft.

In 1995, the NASA Ames Research Center began testing a prototype rotor in an environmental chamber that simulates the atmospheric conditions on Mars [24]. A computational analysis was performed to determine the hover performance of this Mars rotor prototype using an overset-grid, Navier-Stokes CFD flow solver. Through the use of experimental flow conditions, solutions for hovering rotor performance were produced for a series of collective pitch angles. The rotor that was designed for the hover test was simply a proof-of-concept design. Future computational efforts will test possible design changes in order to improve performance.

13 Figure 1.5 Photovoltaic Rotorcraft [24]

In 2000, the NASA Ames Research Center conducted a study specifically addressing the technical challenges facing vertical lift vehicles for Mars, Titan and Venus exploration [25]. The study indicated that autonomous Martian rotorcraft will have large lifting surfaces and will be required to have ultra-lightweight construction. This in turn will pose a challenge in making them sufficiently robust to operate in the Martian environment.

Earlier conceptual design studies at Ames Research Center focused on a tilt rotor configuration [26]. Assuming the use of an Akkerman hydrazine reciprocating engine, a small (10 kg) Mars tilt rotor was shown to potentially have a range capability of 150-250 km. However, it was also clear that deployment of even a small Mars tilt rotor requires human assistance in vehicle assembly. Recent work at Ames has focused on a coaxial helicopter configuration for early Mars exploration missions.

14 In 2000, the University of Maryland proposed the design of a Martian autonomous rotary-wing vehicle [27]. The rotorcraft has a gross mass of 50 kg, allowing a payload capacity of 10.8 kg. It incorporates a coaxial rotor configuration with two blades per rotor. The rotor diameter is approximately 4.3 m and the blades have a chord length of 0.53 m. The rotorcraft is powered by a hydrogen/oxygen fuel cell and can cover a maximum range of 25 km.

In 2004, the Virginia Polytechnic Institute proposed a rotor design for an unmanned helicopter for use on Mars [28]. The design study focused on an analytical determination of the optimum configuration to lift a 100 kg payload effectively in the Martian environment. Single rotor and co-axial rotor configurations were compared as part of the study. The study indicated that a co-axial rotor configuration has significant advantages over a single rotor. For the same overall rotor footprint, a co-axial rotor can lift twice the required mass in hover for just over twice the required power. Alternatively, a co-axial rotor can lift the design mass for half the required power per rotor.

Flapping-Wing Aircraft

Due to the low atmospheric density on Mars, aircraft are required to fly within a very low Reynolds number regime. Insects have succeeded in efficiently exploiting the low Reynolds number flight regime through the flapping motion of their wings. Therefore, flapping wing aircraft may have tremendous potential for Mars applications. However, flapping wing technology is very much in its infancy.

Although the development and testing of an Earth-based, full-scale piloted Ornithopter has been ongoing at the University of Toronto since the 1980's [29], it was only able to achieve sustained flight for 14 seconds when flight-tested in 2006. Its predecessor, a smaller, unpiloted proof of concept model, achieved sustained flight for just over two minutes but was launched off a cliff.

15 Figure 1.6 Harris/DeLaurier Model of Engine-Powered Piloted Ornithopter

Studies into the flight dynamics of insects and birds have shown that the wing motion that allows optimal flapping wing flight is complicated and not easily replicated by mechanical structures [30]. Furthermore, insects and birds have the ability to store energy in their muscles allowing them to fly efficiently. Since energy storage devices are still under development, their application into flapping wing devices is currently not possible resulting in energy intensive flapping wing systems.

In 2003, a study performed at Georgia Tech Research Institute resulted in a preliminary design of an entomopter-based Mars surveyor [31]. The baseline design specifications for the Mars Entomopter included a wingspan of 1 m, an aspect ratio of 5.87, and a wing area of 0.546 m2. The study indicated that majority of the mass budget of the Entomopter would be taken up by fuel since the energy requirements for flapping wing vehicles are high. Therefore, Entomopter flight durations would be brief (5 to 10 minutes).

The Canadian Space Agency is currently experimenting on various flapping wing prototypes [32]. The prototypes have not been optimized for actual flight and are simply first iterations in order to measure the lift forces produced. A four-wing design, developed by the University of Toronto Institute of Aerospace Studies, is one of the prototypes being tested at the Canadian Space Agency [33]. However, to date, no quantitative data has been collected. A RotaFlap prototype that simplifies the usual 'back and forth' motion traditionally required to flap a wing is also being tested at the Canadian Space Agency [33]. Testing of this prototype indicates that its lift coefficient lies in the range of insects and hummingbirds.

16 1.5 Objectives of Thesis

Comparative studies of different aerial platforms for Mars applications performed by the Jet Propulsion Laboratory (JPL) [4] and the Massachusetts Institute of Technology [6] have addressed the advantages offered by each of the different platforms, the types of scientific objectives best accomplished by the use of specific platforms, and very high level estimations of the mass and power requirements for select aerial platforms. However, these studies have been general in nature and have not included a thorough, in- depth comparative analysis of the performance of the different types of aerial platforms.

The primary objective of this thesis is to perform a detailed comparative study of the performance of three types of aerial platforms, namely, the fixed-wing aircraft, the airship, and the rotary-wing aircraft for Mars applications. In order to achieve this objective, the performance of these three aerial platforms will be compared in the context of specific Mars mission scenarios and will only address the performance of the platforms during cruising flight. Performance of the aerial platforms will be measured based on common parameters related to mission requirements.

The results of the comparative study will be used to determine which of the three aerial platforms is best suited to perform the specific Mars missions. The scope of the thesis only includes assessing the relative performance of the three types of aerial platforms. It does not include designing the platforms. Therefore, existing aerial platform designs have been identified for the purpose of performing this comparative study.

The next chapter presents the three Mars mission concepts that provide the context for comparing the three types of aerial platforms. For each mission, the scientific investigations (derived from the MEPAG Report) to be performed by each platform are identified. In addition, the associated mission requirements including the range, altitude and flight path, as well as the associated instrumentation are also presented. The performance of the three aerial platforms will be measured based on the mission requirements defined in this chapter for each mission case.

17 Chapter 3 presents the aerial platform designs that have been selected as baseline designs for performing this comparative study. Depending on the mission requirements, the existing designs will be scaled as required in order to accommodate the different payloads identified in each mission as well as the fuel and propulsion system required for each mission.

Chapter 4 presents the metrics that will be used to measure the performance of each of the three aerial platforms. For each mission, the metrics will be used to determine a performance rating for each platform. This performance rating will provide an indication of which platform is best suited for each mission. The method by which these metrics will be consolidated into a performance rating is also presented in this chapter.

Chapter 5 presents the evaluated performance metrics for each platform and for each mission, as well as a discussion of the results of the analysis. In addition, this chapter presents the performance rating for each platform indicating the type of aerial platform that is best suited to conduct each mission.

Chapter 6 provides the conclusions of this study as well as some recommendations for future research.

18 Chapter 2

MISSION CONCEPTS FOR MARS EXPLORATION

This chapter discusses the types of scientific investigations that can be performed using aerial platforms; the grouping of these investigations into scientific missions; and the associated mission requirements that will be used to compare the performance of the three types of aerial platforms.

The MEPAG report was used as a guideline for determining the types of scientific missions that could be performed using aerial platforms. The MEPAG report outlines four major goals that would contribute to increasing our knowledge of the Martian environment and the history of its evolution including [3]: I. Determining if life ever arose on Mars II. Understanding the processes and history of climate on Mars III. Determining the evolution of the surface and interior of Mars IV. Preparing for human exploration

Associated with each goal are scientific investigations aimed at making specific measurements for the purpose of gathering data to achieve the goal. The investigations associated with each goal were reviewed in order to identify those that could be conducted via aerial platforms. These investigations were then grouped to form scientific missions that would address each goal.

The investigations relating to the fourth goal, preparing for human exploration, pertain to collecting data for the purpose of understanding the hazards in the Martian environment so as to design systems to mitigate them, as well as demonstrating technologies that are

19 currently being developed but that are still in the infant stages or conceptual phases of development. Many of these investigations pertain to ground-based collection of data requiring the use of landers and subsurface penetrators. Since, aerial platforms are less relevant for addressing this goal, it will not be pursued further in the present work.

Consultations with scientists from the Canadian Space Agency were useful in identifying the investigations from the MEPAG report that could be performed via aerial platforms, the locations on Mars that would be suited to perform such investigations and the types of instruments required.

An important consideration in selecting instruments that will be carried on aerial platforms is the mass of the instruments; the size of the aerial platform is a function of the overall takeoff mass. It is envisioned that the aerial platforms will be delivered to Mars in a Viking derivative aeroshell having an outer diameter of 2.65 m and a maximum internal diameter of 2.48 m [34]. This type of aeroshell, illustrated in Figure 2.1, has been proposed for similar aerial missions such as the ARES mission [34].

Backshetl

Figure 2.1 Viking Derivative Aeroshell [34]

20 The following sections describe in more detail each of the scientific missions that have been developed to address the goals of the MEPAG report.

2.1 Determining If Life Ever Arose on Mars

The MEPAG report identifies three scientific objectives related to the goal of determining whether life ever arose on Mars [3]: A. Assess the past and present habitability of Mars B. Characterize carbon cycling in its geochemical context C. Assess whether life is or was present on Mars

With respect to the first objective, in order to determine whether Mars had a high potential for habitability, three conditions need to be satisfied [3]: i) The presence of liquid water ii) The presence of the key elements such as Nitrogen, Phosphorous and Sulfur that are critical elements for life iii) The presence of a source of energy to support life (e.g. chemical redox, pH gradients, geothermal heat, radioactivity, incident radiation)

The MEPAG report identifies specific scientific investigations aimed at gathering data to validate whether these conditions are present on Mars. The investigations pertaining to the first objective of searching for the presence of water on Mars can be accomplished by taking remote measurements via aerial platforms. However, the investigations pertaining to validating the presence of the other two conditions require either sample collection via landers or global investigations via orbiters.

The other two scientific objectives focus on characterizing organic and inorganic material and require sample collection via landers. Therefore, the only investigations that can be accomplished via aerial platforms to determine whether life ever arose on Mars are those pertaining to searching for water on Mars since the associated data can be obtained by

21 taking remote measurements. Hence, the first mission concept, developed for the purpose of evaluating the performance of aerial platforms, is based on searching for water on Mars.

Mission Concept 1: Searching for Water on Mars

The following scientific investigations from the MEPAG report pertain to searching for water on Mars [3]. 1. Establishing the current distribution of water in all its forms on Mars (liquid water, massive ground ice and polar layered deposits) 2. Understanding at a global scale the abundance, form and distribution of water in Mars' geologic past 3. Finding geomorphic evidence of the former presence of an ocean or seas or overflow channel activity 4. Determining sedimentary stratigraphy and the distribution of aqueous weathering products 5. Characterizing the stratigraphic record of climate change preserved in polar layered deposits, residual ice caps, and other climate-modulated deposits of H2O 6. Determining regionally the present state, 3-dimensional distribution, and cycling of water on Mars 7. Performing morphological mapping of Mars

These investigations can be performed remotely through low altitude surface reconnaissance using aerial platforms and will form the basis for addressing the objectives of the first mission of Searching for Water on Mars.

Various instruments can be used to perform these investigations such as cameras, sounding radars and imaging spectrometers. Cameras provide high-resolution visible imagery of surface features that could provide evidence of the presence of water on Mars. For instance, images taken in 2004 and 2005 from the Mars Orbiter Camera on NASA's

22 Mars Global Surveyor provided new evidence of deposits that suggest water carried sediment through the deposits sometime during the past seven years [35].

Sounding radars provide the ability to probe the subsurface of Mars to search for shallow water and ice deposits. Ground penetrating radar operates by transmitting pulses of ultra high frequency radio waves into the ground through a transducer or antenna. The transmitted energy is reflected back from various buried objects. The antenna receives the reflected waves and stores them in a digital control unit. The Mars Reconnaissance Orbiter that has been in operation on Mars since March 2006 carries a Shallow Subsurface Radar (SHARAD) designed to seek liquid or frozen water in the first few hundreds of feet (up to 1 kilometer) of Mars' crust using this principle [36].

Spectrometers can be used to identify minerals, especially those likely formed in the presence of water, by detecting the emission or absorption of certain energies or wavelengths of the corresponding atoms. The Mars Reconnaissance Orbiter carries a Compact Reconnaissance Imaging Spectrometer for Mars (CRISM) that performs exactly this function [37]. However, sounding radars offer advantages over spectrometers since they have the ability to produce continuous cross-sectional profiles that can be used to investigate the presence and continuity of natural subsurface conditions and features.

Therefore, for the purpose of this mission, the Shallow Subsurface Radar (SHARAD), identical to the one used on the Mars Reconnaissance Orbiter, will be used to detect liquid water and profile ice in the Martian subsurface. In addition, the instrument will perform surface characterization, three-dimensional stratigraphy and morphological mapping. SHARAD has a mass of 15 kg and a peak power consumption of 60 W [38].

From its orbital platform, SHARAD has a vertical resolution of approximately 10 to 20 m [39]. Since SHARAD is currently operated from an orbital platform it is required to operate at higher frequencies of 10-25 MHz in order to achieve the desired high depth resolution [40]. However, low frequency (mHz - kHz) soundings are ideally suited to groundwater detection due to their great depths of penetration [41], but low frequency

23 soundings are only feasible from aerial platforms. Therefore, operating SHARAD from an aerial platform using a lower frequency band will produce even greater vertical resolution than is currently achieved.

Visible imagery will be gathered of the surface in order to determine the distribution of ground ice, as well as to obtain high-resolution imagery of surface features, such as overflow channels, providing evidence of the presence and activity of liquid water on the surface of Mars. The Mars Imaging Camera, a visual wavelength camera identical to the one used on the failed 1998 Nozomi mission, will be used to perform the optical imaging of Mars. The camera has a mass of 2.7 kg and a power consumption of 14 W [42].

In principle, the investigations pertaining to this mission should be performed across all regions of Mars in order to obtain a view of the global distribution of water on Mars. However, it is not feasible for a single aerial mission to cover such a large distance since mass is a limiting factor, but multiple platforms carrying the same set of instruments can be deployed in different regions on Mars in order to achieve global coverage. For the purpose of this study, a region of Mars thought to have a high potential for preserving subsurface water in the form of ice was selected.

Based on data gathered from the MARSIS antenna on the Mars Express spacecraft, it is believed that a 250-km-wide circular structure that lies between 1.5 and 2.5 km below the surface of Chryse Planitia is an impact crater that was buried with volcanic ash or soil several billion years ago [43]. No radar boundaries were detected in the material that fills the bowl of the crater and the radar signals lost little strength when passing though it. This suggests that the infill must contain a large proportion of ice, which is nearly transparent to radar. Therefore, Chryse Planitia has been selected as the location on Mars at which to conduct this mission.

The aerial platforms will be expected to cover a range of 500 km between the following coordinates: (28.4N, 41.6W) -> (30.9N, 40W) -> (28.4N, 38.4W) -> (28.4N, 41.6W) as illustrated in Figure 2.2.

24 This triangular flight path was selected so as to cover as much of the 250-km-wide circular region over Chryse Planitia thought to have water buried beneath it. Each corner of the triangular flight path touches the outer perimeter of the 250-km-wide circular region. The flight path requires the platform to make two 120-degree right turns. The aerial platforms will fly at the Martian datum where the atmospheric density is approximately 0.01245 kg/m3. The Martian datum is the elevation at which the atmospheric pressure is 6.1 millibars, or 610 Pascals (atmospheric pressure has to be used because Mars has no ocean; therefore, there is no "sea level" equivalent to that on Earth) [44]. The scientific payload to be carried by the aerial platforms has a total mass of 17.7 kg and a power requirement of 74 W. Table 2.1 provides a summary of the mission requirements.

Figure 2.2 Chryse Planitia - Site for Conducting Mission 1

25 Location Chryse Planitia Range 500 km Instruments SHARAD-15kg, 60 W Mars Imaging Camera - 2.7 kg, 14 W Flight Path Triangular flight path through the following coordinates (28.4N, 41.6W) -> (30.9N, 40W) -> (28.4N, 38.4W) -> (28.4N, 41.6W). Two 120-degree turns required.

Table 2.1 Requirements for Mission 1

2.2 Understanding the Processes and History of Climate on Mars

The fundamental scientific questions that underlie the second MEPAG goal are how the climate of Mars has evolved over time to reach its current state, and what processes have operated to produce this evolution. The MEPAG report identifies three scientific objectives related to this goal [3]: A. Characterize Mars' atmosphere, present climate, and climate processes B. Characterize Mars' ancient climate and climate processes through study of the geologic and volatile record of climate change C. Characterize the state and processes of the Martian atmosphere of critical importance for the safe operation of spacecraft

The first objective is crucial to understanding the present state of the entire atmospheric system (from the surface-atmosphere boundary to the exosphere) and forms the baseline for interpreting the climate on Mars. The scientific investigations pertaining to the first objective entail taking measurements of [3]: • Atmospheric parameters such as water vapour, wind speed, temperature, humidity and dust in the lower atmosphere (<80 km) • The upper atmosphere including volatile escape rates, interaction of the upper atmosphere with solar wind, and variability of the neutral and plasma environment above -80 km

26 Aerial platforms provide the best means of performing investigations pertaining to the lower atmosphere of Mars (<80 km) since higher resolution data can be obtained from aerial platforms than from orbiters. However, orbiters are required to perform investigations pertaining to the upper atmosphere of Mars since the atmospheric density in this region is too low to permit the operation of aerial platforms.

The second objective focuses on specific investigations that will measure key indicators of the past climate of Mars. Such investigations include the characterization of the stratigraphic record of climate change preserved in polar layered deposits, residual ice caps, other climate-modulated deposits of H2O, CO2 and dust found elsewhere on the planet [3]. The scientific investigations performed in Mission 1 tie into this objective and the data collected from that mission could also serve to answer questions about the ancient climatic conditions on Mars.

The third objective highlights mission critical atmospheric measurements that will reduce mission risk and enhance overall science return, benefiting all future missions to Mars. The scientific investigations pertaining to this objective include determining the thermal and dynamic behaviour of the planetary boundary layer as well as the atmospheric mass density in the upper and lower portions of the Martian atmosphere. As with the first objective, investigations pertaining to determining the atmospheric density of the lower atmosphere of Mars can be accomplished via aerial platforms; however, orbiters are required to perform investigations pertaining to the upper atmosphere of Mars.

The second mission concept, developed for the purpose of evaluating the performance of aerial platforms, is based on studying the present climate of Mars. The mission focuses on performing the scientific investigations related to the lower atmosphere on Mars.

27 Mission Concept 2: Studying the Present Climate of Mars

The following scientific investigations from the MEPAG report pertain to studying the present climate of Mars [3]: 1. Investigating the lower atmospheric climate and processes 2. Searching for micro-climates 3. Determining the present-day composition of the atmosphere 4. Determining the thermal and dynamical behaviour of the lower portion of the atmosphere 5. Determining the atmospheric density of the lower portion of the atmosphere 6. Understanding the mechanism of regional and global dust storm development

The above investigations entail taking measurements of atmospheric parameters such as temperature, pressure, humidity, wind speeds, and atmospheric gases in the lower atmosphere. Determining the atmospheric composition of Mars, through measurements of atmospheric gases, will enable scientists to determine the atmospheric mass density.

Meteorological instrument packages that include temperature, pressure, relative humidity and wind speed sensors have been used on various missions including the Mars-96, Mars Pathfinder and Mars Polar Lander missions. For the purpose of this mission, simultaneous temperature, pressure and humidity profiles, as well as wind speed and direction will be measured using the Meteorology Instrument Package identical to the one used on 1996 Mars Pathfinder. The Meteorological Instrument Package has a mass of 2.04 kg and a power consumption of 3.2 W [45].

Measurement of atmospheric gases can be best accomplished via spectrometers. For the purpose of this mission, a Tunable Laser Spectrometer, similar to the one proposed for the 2009 Mars Science Laboratory mission [46], will be used to obtain atmospheric gas measurements. The instrument has the capability of making accurate measurements of the mixing ratios of several target gases such as H20, CH4, CO, C02, OCS, and H2O2, and their isotope ratios [46]. Sensitivities for methane and hydrogen peroxide will be down to

28 a minimum detectable amount of 20 parts per trillion [47]. A laser absorption technique enables the instrument to be mounted on the outside of an airframe with air passing between two mirrors. The "open path" configuration eliminates the need for a sampling system, thus simplifying the requirements and improving speed of response. Furthermore, the instrument design is conducive to use on aerial platforms [48]. The Tunable Laser Spectrometer has a mass of 0.2 kg and low power consumption [49].

Finally, a visual wavelength camera, such as that used in Mission Concept 1, will be used to monitor cloud coverage as well as dust storm development.

In principle, the investigations pertaining to this mission should be performed across all regions of Mars in order to obtain a view of the global climate of Mars. However, it is not feasible for a single aerial mission to cover such a large distance since mass is a limiting factor, but multiple platforms carrying the same set of instruments can be deployed in different regions on Mars in order to achieve global coverage.

For the purpose of this study, Chasma Boreale, a valley that cuts into Mars' north polar ice cap, was selected as the site for conducting this mission. Images of this region recently obtained by NASA's Mars Reconnaissance Orbiter (MRO) show a dust layer sandwiched between layers of ice near the north pole suggesting that the planet's climate has shifted dramatically in the past 100,000 years or so [50].

"Three layers are visible in the cap. The one on top corresponds to the most recent period in Mars' climate history, during which ice has built up around the north pole. The bottom layer is also made of ice and appears to represent a period with a similar climate. In between is a layer of dust. Its presence suggests an era with a very different climate, when ice was no longer being deposited in the north polar region" [50]. By studying the present climatic conditions in this region, scientists will be able to gain deeper insight into the dynamic change of Mars' climate over time.

29 The aerial platforms will be expected to cover a range of 1000 km between the following coordinates: (89N, 47.1W) -> (82.9N, 47.1W) -> (89N, 5E) -> (84.7N, 5E). This "Z- shaped" flight path, illustrated in Figure 2.3, requires the platform to make two 110- degree turns (one to the left and one to the right).

Figure 2.3 Chasma Boreale - Site for Conducting Mission 2

The aerial platforms will fly at an altitude of 3 km above the Martian datum where the atmospheric density is approximately 0.00944 kg/m . The scientific payload to be carried by the aerial platforms has a total mass of 4.94 kg and a power requirement of approximately 20 W. Table 2.2 provides a summary of the mission requirements.

30 Location Chasma Boreale Range 1000 km Instruments Meteorological Instrument System - 2.04 kg, 3.2 W Tunable Laser Spectrometer - 0.2 kg, ~ 2 W Mars Imaging Camera - 2.7 kg, 14 W Flight Path "Z-shaped" flight path through the following coordinates (89N, 47.1W) -> (82.9N, 47.1W) -> (89N, 5E) -> (84.7N, 5E). Two 110-degree turns required.

Table 2.2 Requirements for Mission 2

2.3 Determining the Evolution of the Surface and Interior of Mars

Insight into the composition and structure of Mars is fundamental in providing insight into the history and processes of our own planet. Geology is an important aspect in studying the conditions that are conducive to the origin and persistence of life on Mars. In addition, the study of the interior provides important clues about a wide range of topics including the early history of Mars, sources of volatiles, and geothermal energy. The MEPAG report identifies two scientific objectives related to this goal [3]: A. Determine the nature and evolution of the geologic processes that have created and modified the Martian crust and surface B. Characterize the structure, composition, dynamics and evolution of Mars' interior

Scientific investigations pertaining to both these objectives include [3]: • Studying the chemical and mineralogical composition of the crust, near- surface rocks and large-scale vertical structures • Investigating igneous processes such as volcanic outgassing and volatile evolution • Determining the distribution of water on Mars since water is an important geologic material that influences most geological processes such as the

31 formation of sedimentary, igneous and metamorphic rocks, the weathering of geological materials, and deformation of the lithosphere (this investigation is covered in Mission 1) • Determining the tectonic history of the Martian crust through magnetic and gravity data as well as seismic monitoring • Geologic mapping using global topographic data.

Aerial platforms provide an ideal means of obtaining localized high-resolution mineralogical data that can be correlated against data obtained from orbiters. The third mission concept, developed for the purpose of evaluating the performance of aerial platforms, is based on performing a mineralogical, thermophysical and magnetic study of Mars.

Mission Concept 3: Mineralogical, Thermophysical and Magnetic Study of Mars

The following scientific investigations from the MEPAG report pertain to studying the mineralogical, thermophysical and magnetic makeup of Mars [3]: 1. Determining the chemical and mineralogical composition of the crust, near- surface rocks and large-scale vertical structures 2. Investigating geologic deposits that have been affected by hydrological processes 3. Determining the array of potential energy sources available on Mars 4. Determining the thermal properties of Martian surface materials 5. Establishing the nature of the magnetic field of Mars and gravitational anomalies associated with crustal density variations

Various spectral imaging techniques such as Mossbauer spectroscopy, Raman spectroscopy, and Infrared spectroscopy can be used to determine the chemical and mineralogical composition of Mars. Mossbauer spectroscopy can only be applied to a relatively small group of atoms including: 57Fe, 129I, U9Sn, and 121Sb [51]. The main difficulty of Raman spectroscopy is separating the weak inelastically scattered light from

32 the intense Rayleigh scattered laser light [52]. Infrared (IR) spectroscopy, on the other hand, is one of the most common spectroscopic techniques used by organic and inorganic chemists.

The main goal of IR spectroscopic analysis is to determine the chemincal functional groups contained in a sample; different functional groups absorb characteristic frequencies of IR radiation. IR spectrometers can accept a wide range of sample types such as gases, liquids and solids. Thus IR spectroscopy is an important tool for compound identification [53].

For the purpose of this mission, the Miniature Thermal Emission Spectrometer (Mini- TES) - an infrared spectrometer, identical to the one developed for the 2003 Mars Exploration Rover Mission, will be used to determine the mineralogy of rocks and soils from a distance by detecting their patterns of thermal radiation. Mini-TES will record the spectra of various rocks and soils. These spectra will be studied to determine the type of minerals and their abundances at the selected location. Mini-TES weighs 2.4 kg and has an average power requirement of 5.6 W [54].

High-resolution visible imagery of the large-scale vertical structures on Mars can provide evidence of processes such as weathering and erosion, as well as evidence of volcanism and the activity of liquid water at the surface. A visual wavelength camera similar to the one used in Mission Concept 1 will be used to obtain imagery of the large-scale vertical structures on Mars.

Mars has, at best, a weak surface magnetic field although it may once have had a fluid core and a dynamo. If such a fluid core and dynamo did exist, those geologic units which were formed early in the history of Mars may well carry a memory of this effect. Magnetic mapping over a variety of well-mapped geological features of varying age could reveal a great deal about the history of any Martian magnetic field. The strength of the magnetic field on the surface of Mars will be determined by the iron content of Martian rocks together with the strength of the external magnetic field to which these

33 rocks were exposed. Measurements of magnetic field are performed using magnetometers.

Various types of magnetometers exist including fluxgate magnetometers, quantum magnetometers, proton precession magnetometers, Overhauser magnetometers and potassium magnetometers. The fluxgate magnetometer has most commonly been used for space-based applications. For instance, the Russian MARS-96, Pioneer Venus, Mars Global Surveyor, Voyager, Magsat and Giotto missions all included fluxgate magnetometers to measure the magnetic field of Mars [55, 56, 57]. For the purpose of this mission, a fluxgate magnetometer identical to the one used on the Mars-96 mission will be used to measure the magnetic field on Mars. The fluxgate magnetometer has a mass of 0.25 kg and a power requirement of 1 W [58].

The site on Mars selected to conduct this mission is Valles Marineris, a system of canyons located just south of the Martian equator. "The geologic history of the central canyon system is complex: first the surface collapsed into a few deep depressions that later became filled with layered material, perhaps as lake deposits. Then graben-forming faults cut across some of the older troughs thus widening existing troughs, breaching barriers between troughs, and forming additional ones. At that time the interior deposits were locally bent and tilted. Water, if still present, would have spilled out and flowed toward the outflow channels" [59]. In order to determine the history of the Valles Marineris, the current-day features must be noted and carefully examined, with attention paid to how they may relate to possible previous processes. Aerial platforms provide the ideal means of exploring this region since orbiters cannot obtain high-resolution imagery of the canyon walls and the rough terrain poses challenges for rovers.

The aerial platforms will be expected to cover a range of 150 km between the following coordinates: (1 IS, 70W) -> (1 IS, 67.4W). As illustrated in Figure 2.4, the platforms will be required to fly as close to the walls of the canyon as possible making it necessary for the platforms to be highly maneuverable since the walls of the canyon are uneven.

34 Figure 2.4 Valles Marineris - Site for Conducting Mission 3

The aerial platforms will fly at the Martian datum where the atmospheric density is approximately 0.01245 kg/m3. The scientific payload to be carried by the aerial platforms has a total mass of 5.35 kg and a power requirement of 20.6 W. Table 2.3 provides a summary of the mission requirements.

Location Valles Marineris Range 150 km Instruments Mini-TES (IR Spectrometer) - 2.4 kg, 5.6 W Fluxgate Magnetometer - 0.25 kg, 1 W Mars Imaging Camera - 2.7 kg, 14 W Flight Path Flight path through the following coordinates (US, 70W) - > (1 IS, 67.4W). Multiple sharp turns required.

Table 2.3 Requirements for Mission 3

35 Chapter 3

AERIAL PLATFORMS FOR MARS EXPLORATION

Aerial platforms can be grouped into two broad categories: Lighter Than Air (LTA) vehicles and Heavier Than Air (HTA) vehicles. Within each category there are many levels of capability with varying degrees of technological maturity. The following sections describe the capabilities of various LTA and HTA concepts.

3.1 Lighter-Than-Air Vehicles

LTA vehicles are the simplest kind of aerial vehicle and in principle can operate for extended periods in the atmosphere. The primary challenge for LTA vehicles on Mars is the very thin atmosphere resulting in the need for large gasbag volume. The atmospheric density at Mars' datum corresponds to the Earth's atmospheric density at 33,500 m above sea level. The average temperature on Mars is -63 °C, with a maximum temperature of 20 °C and a minimum temperature of -140 °C [74]. Stratospheric balloons have flown as high as 51,800 m on Earth; therefore, there are no fundamental obstacles to LTA flight on Mars.

Balloons are the simplest of all LTA vehicles, consisting of nothing more than a flexible "envelope" filled with a "lifting gas" that is lighter than the surrounding atmosphere. Although a balloon has no propulsion system, a degree of directional control can be attained by rising or sinking in altitude to find favorable winds.

There are two types of light-gas balloons: "zero-pressure" and super-pressure". Zero- pressure balloons are the traditional form of light-gas balloon. When ground-launched on

36 Earth, they are partially inflated with the light gas prior to launch, and the gas pressure is maintained the same on the inside and the outside of the balloon (i.e., the pressure difference is zero). As the zero-pressure balloon rises, its gas expands to maintain the zero pressure difference, and the balloon's envelope swells. At night, the gas in a zero- pressure balloon cools and contracts, causing the balloon to sink. A zero-pressure balloon can only maintain altitude by releasing the expanding gas, which may rupture the envelope if the balloon rises too high; or by releasing ballast when it sinks too low. Loss of gas and ballast limits the endurance of zero-pressure balloons to a few days.

A super-pressure balloon, in contrast, is sealed so as to prevent the release of gas as it expands when the balloon rises. Therefore, instead of maintaining a zero pressure difference, the internal gas pressure of the balloon is allowed to exceed that of the external atmosphere while maintaining a constant volume. In this manner, the super- pressure balloon maintains an altitude of constant density in the atmosphere, and can maintain flight until gas leakage or a command from the ground (the more likely situation) gradually brings it down. In recent months, super-pressure balloons have achieved a flight endurance of weeks, rather than days.

Balloon designs for possible planetary missions have involved a few unusual concepts. One is the solar, or infrared (IR) Montgolfiere [16]. "This is a hot-air balloon where the envelope is made from a material that traps heat from sunlight, or from heat radiated from a planetary surface. The Solar Montgolfiere Balloon is perhaps the simplest LTA concept since it involves no inflation system. A balloon is filled with atmospheric gas by ram pressure as the vehicle descends on a parachute and solar heating of the balloon envelope rapidly raises the temperature of the contained gas allowing it to stay aloft. However, it functions as long as there is sunlight: several hours, except near the poles where it could be much longer" [16].

Another concept is the "reversible fluid" balloon that consists of an envelope connected to a reservoir containing a fluid that is easily vaporized. Vaporizing the fluid into gas causes the balloon to rise and condensing the gas back into fluid causes it to descend.

37 Balloons have a number of advantages for planetary exploration including their low cost, their ability to cover considerable range, and their ability to examine wide swaths of terrain with greater detail than would be available from an orbiting satellite due to the low altitude at which they can fly. In addition, the balloon envelope can be compactly folded into the aeroshell and easily inflated upon deployment. Figure 3.1 illustrates the deployment sequence of the balloon [60].

Figure 3.1 Balloon Deployment Sequence [60]

Airships also achieve lift through buoyancy of the lifting gas within the airship envelope, however, aerodynamic surfaces such as wings, ailerons and vertical and horizontal stabilizers provide directional control. Airships are more sophisticated than balloons in that a higher degree of directional control can be achieved through the use of thrusters and aerodynamic control surfaces.

Three main categories of airships exist. Rigid airships maintain an external shape independent of the internal pressure within their gas cells. Semi-rigid airships rely upon gas pressure for envelope shape, but use rigid structural members for load distribution

38 and appendage mounting points. Non-rigid airships use gas pressure for both shape maintenance and structure. Non-rigid airships are preferred for planetary exploration since they are lighter and easier to pack into an aeroshell.

3.2 Heavier-Than-Air Vehicles

HTA vehicles depend on relative motion through the air for lift. Therefore, these vehicles will typically fly on Mars at velocities much greater than those of LTA vehicles due to the thin Martian atmosphere, and will require propulsion systems. As a result, HTA flights will often be of shorter duration than those of LTA flights covering equivalent surface area.

Various types of conventional propulsion systems can be considered for planetary flight including electrical propulsion systems (such as fuel cells, batteries and solar panels), combustion engines, bipropellant and monopropellant rocket systems, and un-powered systems such as gliders. These systems require consumable fuel or energy sources, thus having an impact on the gross takeoff mass and operational duration. Another constraint is that the oxygen content in the Martian atmosphere is insufficient to sustain internal combustion or jet engines. As a result, these types of propulsion systems will need to be modified to operate on monopropellant fuel, or carry along an oxidizer. Monopropellant propulsion systems are preferred for planetary applications.

HTA concepts that have been considered and are clearly feasible include gliders and powered fixed-wing aircraft. Rotary wing aircraft (such as gyrocopters and helicopters) are aerodynamically less efficient than fixed-wing aircraft but may also be feasible in instances where hovering and vertical landings and take-offs outweigh the performance penalty. The main challenge for HTA vehicles on Mars is to generate sufficient lift from the very thin atmosphere. This is further complicated by the fact that Mars HTA vehicles will need to operate at low Reynolds numbers and high subsonic Mach numbers due to the lower speed of sound in CO2 (Mars' most abundant atmospheric constituent). The

39 detailed properties of the Martian atmosphere will be presented and discussed in more detail in Chapter 4.

Another significant challenge in the use of HTA vehicles is their deployment. Since there are no ground takeoff sites on the surface of Mars, fixed-wing aircraft will have to be deployed during descent. Folding of the aircraft wings will likely be required in order to accommodate the platform in the aeroshell. This will necessitate mechanical unfolding of the structure during deployment, further increasing the complexity and risk associated with the use of HTA vehicles. Although rotary-wing aircraft may allow ground deployment, the likely need to mechanically unfold the rotor blades again results in increased complexity and risk.

As indicated in the literature review, multiple designs for both Lighter-Than-Air and Heavier-Than-Air vehicles have been proposed for Mars exploration. Since the objective of this study is to analyze the relative performance of the different types of aerial platforms, existing aerial platform designs have been selected to serve as baseline models for analysis. These will then be scaled to suit the requirements of each of the candidate missions. The following sections provide a description of the three types of aerial platform designs that have been selected for comparison in this study.

3.3 Aerial Platforms Considered in this Study

For the purpose of this study, three types of aerial platforms were considered: fixed-wing aircraft, airships, and rotary-wing aircraft. The use of flapping-wing aircraft was initially considered; however, given that flapping-wing technology is still in its infancy (we have still not been able to mechanically replicate efficient flapping wing motion such as that of birds and insects), this mode of flight was not considered in this study. Appendix B provides more detailed information on flapping-wing flight and the performance of existing flapping-wing aircraft.

40 3.3.1 Fixed-Wing Aircraft Configuration

Airfoil characteristics are strongly affected by the Reynolds number, which represents the ratio between the dynamic and the viscous forces in a fluid and is given by:

VI Re = — (3.1) v where V represents the mean fluid velocity, / represents the characteristic length and v represents the kinematic fluid viscosity.

The thin atmosphere of Mars results in fixed-wing aircraft operating at Reynolds numbers in the range of a few hundred to a few hundred thousand in the Martian atmosphere. By comparison, a typical conventional aircraft wing on Earth operates at a Reynolds number of about ten million [61]. However, Earth-based human-powered aircraft operate in the same range of Reynolds numbers as fixed-wing aircraft on Mars due to the low speeds at which they fly in comparison to conventional Earth-based aircraft. This results in Earth- based human-powered aircraft having physical configurations that could be similar to Martian fixed-wing aircraft.

In 1984, a comparative study between the performance of an airplane, a balloon and a dirigible airship in the Martian environment was conducted by MIT [6]. The study postulated that fixed-wing aircraft on Mars will likely resemble efficient Earth-based human-powered aircraft (such as the Gossamer Albatross, Light Eagle and Daedalus 88) since their light structures and low wing loading are essential attributes required for fixed-wing aircraft to remain aloft in the Martian atmosphere. As such, the design of an Earth-based human-powered aircraft will be used as a baseline model for analyzing the fixed-wing case in this comparative study of Martian aerial platforms.

The Gossamer Albatross was the first generation of human-powered aircraft to be built and flown successfully on Earth. On June 12, 1979 the Gossamer Albatross, powered by

41 an amateur cyclist, Bryan Allen, using pedals to drive a large two-bladed propeller, successfully crossed the English Channel covering a range of 35.8 km in 2 hours and 49 minutes, achieving a top speed of 29 km/h and an average altitude of 1.5 m [62]. Figure 3.2 illustrates the Gossamer Albatross crossing the English Channel.

Figure 3.2 Gossamer Albatross [62]

The Daedalus Project, inspired by the famous Greek myth about a craftsman named Daedalus who built his own wings to fly to freedom from imprisonment, was established by a group of students, professors, and alumni from the Massachusetts Institute of Technology in the mid 1980's. The goal of the project was to design, build and test a human-powered aircraft that could fly farther than 115 km [63].

Three human-powered aircraft were constructed as part of the project: The Light Eagle, a prototype aircraft that set a closed course distance record of 59 km on January 22, 1987; the Daedalus 87; and the Daedalus 88, that flew 119 km on April 23, 1988, from the Iraklion Air Force Base on Crete, Greece to the island of Santorini in 3 hours, 54 minutes

42 and set a new record in distance and endurance for a human powered aircraft [63]. The prototype Light Eagle was designed to be capable of a cliff launch and sustained flight at altitudes of about 300 m. The design of the Daedalus, on the other hand, was optimized for low-level flight by decreasing the load factor.

Due to its superior performance over its predecessors, the Daedalus 88 human-powered aircraft has been selected as the baseline model for analyzing the performance of fixed- wing aircraft in this comparative study. Figure 3.3 depicts the Daedalus 88 flight model. The Daedalus 88 has an empty mass of 31.1 kg and a wingspan of 34.1 m. An 8.8-meter fuselage accommodates the pilot and consumables. Leg-powered bicycle pedals are used to rotate a 3.35-meter propeller and hand controls are used to manipulate the rudder and elevator and to adjust the variable pitch propeller. The lack of ailerons in the current design necessitates a large turn radius.

NASA Dryden Flight Research Center Photo Collection http:/Awww,dfrc.nasa.gov/gallery/photo/index.html NASA Photo: EC88-0059-002 Date: March 7,1988 Photo by: Beastey

Daedalus - Last Dryden flight

Figure 3.3 Daedalus 88

43 Figure 3.4 depicts orthogonal projections of the Daedalus 88 [64].

XL jj^manrmi"'

WBBIpfnBJ

Oatdalus The Daedahu Projtet V«y LMk| Kl*p HiMH PMrtrtfd AUvnft 9 * M

Figure 3.4 Daedalus 88 Configuration [64]

Table 3.1 presents the characteristics of the Daedalus 88 in the Earth environment [7]. Although the Daedalus 88 is equipped with control surfaces, the current design requires a human to steer the controls and to produce the power required to fly the aircraft. On Mars the aircraft will be required to fly autonomously; as such, a flight control system as well as a propulsion system will be required. These items are not reflected in the mass budget presented in Table 3.1. The characteristics of the Daedalus 88 in the Martian environment will be presented in Chapter 5 along with the results of the analysis.

44 Wing Span (m) 34.1 Wing Area (mz) 31 Aspect Ratio 37.51 Weights (kg) Empty 31.1 Pilot 72.5 Consumables 6 Gross 109.6 Cruise Conditions Specific Power (W/kg) 2.80 Power (W) 203 V(m/s) 6.87 q (N/m2) 28.87 Aero Coefficients

CL 1.20 Cdo 0.011 Cas(parasite) 0.18 Drag (N) Induced 12.15 Profile 9.4 Parasite 5.06 Total 26.61

Table 3.1 Characteristics of Daedalus 88 in the Earth Environment

Since the aerial platforms will have to operate autonomously on Mars, an on-orbit sensor and perception system providing global navigation is required. This system would consist of an on-board computer with state estimation algorithms and sensors such as rate gyros, accelerometers, barometers, thermocouples, and anemometers. The state estimation algorithms combine pressure and vertical airspeed measurements from the accelerometer and anemometer with a relatively simple on-board model of the platform up and down motion to produce a continuous estimate of how high the platform is with respect to a selected reference altitude. Accelerometer and angular rate gyro measurements are combined with an on-board model of the platform attitude dynamics to produce continuous estimates of the platform attitude and its angular rates [17]. Temperature, pressure and wind speed measurements can be incorporated into flight software, such as the Application Navigation System designed by the NASA Advanced Supercomputing

45 Division [65], to manipulate the aerial platform control surfaces in order to maintain the pre-determined flight path. Flight control systems and navigation sensor packages for Earth-based unmanned aerial vehicles, such as the GuideStar 211HG-7 developed by Athena Technologies, are readily available [66]. Such a package weighs approximately 3 kg and has a volume of less than (0.15m x 0.15m x 0.15m).

As mentioned above, monopropellant propulsion systems are preferred for planetary applications. As such, a hydrazine-based Akkerman engine, similar to the one proposed for the Mini-Sniffer Mars aircraft concept [67], has been selected as the propulsion system. This monopropellant engine has been successfully used on high altitude, long endurance terrestrial experimental aircraft. The Akkerman engine has a specific fuel consumption, /, of 2.7 kg/kW-h and a specific power, ps , of 1.62 kW/kg [68].

On Earth, a pilot having a mass of 72.5 kg provided the power required to fly the Daedalus 88. In addition, 6 kg of consumables were provided to the pilot for the duration of the flight. For Mars-based applications, the mass of the pilot and consumables will be replaced by the mass of the Akkerman engine, hydrazine fuel, fuel tank and scientific payload. The Daedalus 88 fuselage has been retained in order to accommodate these components. Scaling of the aircraft may be required depending on the mass of the payload to be carried and the range to be covered in each mission. Details of how the aircraft will be scaled are presented in Chapter 4.

3.3.2 Airship Configuration

Since airships provide the advantage of directional control that balloons do not offer, the design of a Robotic Martian Airship (RMA) [18], shown in Figure 3.5, has been selected as the baseline LTA vehicle for analyzing the performance of lighter-than-air vehicles in this comparative study. The RMA was jointly designed by Anser and JPL for the purpose of assessing the feasibility of operating an airship on Mars. It is a non-rigid airship with an envelope made of composite material ("2 layers of 0.14 mil mylar glued to 0.25-mil polyethylene with loose weaved skrim of 55 denier Kevlar fibre" [18]).

46 Figure 3.5 Robotic Martian Airship [18]

The airship has a generic elliptical shape and a fineness ratio of 2. Cruising flight of the airship is modeled at zero angle of attack. The literature does not indicate the type of propulsion system proposed for the RMA. For the purpose of consistency, a hydrazine- based Akkerman engine, similar to the one selected for the fixed-wing aircraft, has been selected as the propulsion system as it provides the same advantages (the use of a high energy density monopropellant) as the fixed-wing case. The Akkerman engine has a specific fuel consumption, /, of 2.7 kg/kW-h and a specific power, pSpec, of 1.62 kW/kg

[68].

Table 3.2 provides the specifications of the current RMA design [18]. Of course these characteristics will differ for each of the mission concepts conceived in this study due to the different sized payloads carried in each mission. The airship specifications for the three mission concepts considered in this study are provided in Chapter 5 along with the analysis.

47 Gas Bag Volume (inflated) 16500mJ Factor of Safety for Gas Bag 1.2 Total Mass of Airship 199.66 kg Payload Mass 10 kg Maximum Speed lOm/s Duration 40hrs Max Powered Range at Max Speed 1440 km Length 50.12 m Width 25.06 m

Table 3.2 Specifications of the Robotic Martian Airship

Table 3.3 provides a breakdown of the component masses of the current design [18]. This includes the masses of the operational airship vehicle as well as ancillary equipment, the entry vehicle, and spacecraft masses.

Envelope Fabric 126.27 kg Inflatable Tail 8.1kg Rigging & Miscellaneous 24.6 kg Fuel 13.02 kg Fuel System 2.60 kg Engine Mass 1.07 kg Payload 10 kg Hydrogen (lifting gas) 14.02 kg Hydrogen Tank, Temperature Control 13.77 kg System, and Inflation System (Jettisoned) Parachute Mass 30 kg Total Entry Mass 351kg Total Launch Mass 540 kg

Table 3.3 RMA Mass Summary

48 3.3.3 Rotary-Wing Aircraft Configuration

The Martian Autonomous Rotary-wing Vehicle (MARV), designed in response to the American Helicopter Society's Seventeenth Annual Student Design Competition, has been selected as the baseline model for a rotary-wing vehicle in this study [27]. This vehicle was selected due to its mature design and conformance with NASA's typical design requirements for Scout missions. MARV's design requirements include the ability to maintain sustained controlled flight for a minimum of 30 minutes; a range of at least 25 km; and a maximum mass of 50 kg with a payload capacity of 10.8 kg [27]. Figure 3.6 depicts the configuration the overall MARV [27]. Figure 3.7 depicts the MARV body configuration [27].

A low Reynolds number airfoil, AGRC 1506, was designed for use on MARV. This airfoil has a maximum thickness of 15% chord and a maximum camber of 6% chord [27]. After analyzing numerous energy options, a Proton Exchange Membrane (PEM) fuel cell, that uses pure hydrogen and pure oxygen as the fuel, was selected as the power plant. The fuel cell supplies power to an electric motor that rotates the rotor. The fuel cell system consists of the fuel cell stack, fuel, fuel tanks, thermal and water management units and ancillary control equipment.

49 Forward flight direct

Figure 3.6 Martian Autonomous Rotary-Wing Vehicle Configuration [27]

Bsr

Helicol directional ontenna

Figure 3.7 MARV Body Configuration [27]

50 Table 3.4 provides the rotor specifications for MARV [27].

Number of blades 2 per rotor Radius 2.13 m Maximum chord 0.670 m Tip chord 0.366 m Tip speed 143.75 m/s Q, 4049 RPM Tip Mach number 0.5 Solidity 0.25 Effective chord 0.530 m Thrust coefficient 0.0232 Mean lift coefficient 0.85 Tip Reynolds number 64800 Ratio of forward speed to rotor tip speed (u) 0.08 Climb rate 2.5142 m/s

Constant drag coefficient (Cdo) 0.038 Total flat plate area 62.9 m

Table 3.4 MARV Rotor Specifications

Table 3.5 provides a breakdown of the component masses of the current design [27].

Fuselage 3.4 kg Motor 2.0 kg Fuel Cell 5.6 kg Fuel 2.06 kg Fuel Tanks 2.94 kg Blades 6.0 kg Hub 3.3 kg Control group 4.1kg Drive system 2.2 kg Landing gear 3.6 kg Avionics 4.1kg Payload 10.8 kg Gross weight 50 kg

Table 3.5 MARV Mass Breakdown

51 Chapter 4

AERIAL PLATFORM PERFORMANCE ON MARS

A common basis for comparing the performance of the three types of aerial platforms (fixed-wing aircraft, airship, and rotary-wing aircraft) is required in order to assess their suitability for conducting each of the missions described in Chapter 2. As such, a set of performance metrics based on common mission parameters was identified. The following sections identify these performance metrics and explain how each performance metric will be evaluated for each aerial platform. In addition, the methodology by which the performance metrics will be combined into a single performance rating, that will determine which platform is best suited for each mission, is presented.

To begin, a brief overview of the Martian atmosphere is essential in order to analyze the possibility of flight on Mars since atmospheric conditions play a critical role in the design and performance of aircraft. The Martian atmosphere is composed primarily of carbon dioxide. The variability of Mars' atmosphere in terms of temperatures, dust content, clouds, water vapor, and even total atmospheric pressure, is extreme on all time scales. The average recorded temperature on Mars, as recorded by the Viking 2 Lander over a period of 1281 Mars days, is -63 °C, with a maximum recorded temperature of 20 °C and a minimum recorded temperature of -140 °C [70]. Mars' surface pressure ranges between 4.1 to 6.7 mbar. Wind speeds on Mars can vary between 5 m/s and 40 m/s. Table 4.1 provides a comparison between the properties of Mars and those of Earth [71, 72, 73, 74]. Note that the atmospheric properties presented in Table 4.1 are those at the Martian datum and at sea level on Earth. As mentioned previously, the Martian datum is defined as the elevation at which the atmospheric pressure is 6.1 millibars, or 610 Pascals. Elevation of the Martian surface varies between 9 km below the datum and 21 km above the datum.

52 Mars Earth Mass [kg] 6.42 x 102" 5.98 x 10"4 Radius [m] 3.39 xlOb 6.38 x 10b Gravity [m/s2] 3.73 9.81 Density [kg/m ] 0.01245 1.23 Pressure [Pa] 610 101350 Temperature [K] 207.5 288.2 Dynamic Viscosity [Ns/m2] 1.06 xlO"3 1.8 xlO"5 Kinematic Viscosity [m2/s] 8.51 x 10"4 1.5x10"* Speed of Sound [m/s] 239 340 Atmospheric Composition [%] C02 [95] N2 [79] N2 [2.7] 02 [20] Ar [1.6] Other [1] Other [0.7]

Table 4.1 Atmospheric Properties of Mars and Earth at Datum

One benefit of flying on Mars is that its gravity is lower than that of Earth; as a result, less force is required to lift an aircraft on Mars than on Earth. However, the atmospheric density of Mars is two orders of magnitude lower than that of Earth, thus requiring a Martian aircraft to have larger aerodynamic surfaces and higher speeds to stay aloft. The thin Martian atmosphere, therefore, poses significant design challenges for aerial platforms.

4.1 Performance Metrics

In order to determine which of the three aerial platforms is best suited to conduct each mission described in Chapter 2, a common basis for comparing their performance is required. Performance measures that are common to each platform and that form an important basis for mission design include the gross takeoff mass, power, manoeuvrability and complexity.

Mass is an important parameter due to launch considerations. The cost of launching payloads to Mars is proportional to the mass of the payload to be carried; thus,

53 minimizing the gross takeoff mass is essential in order to minimize launch costs. Additional power, over the cruising power requirements, may be needed to provide flexibility in terms of varying the speed of the platform, performing turns, climbing, hovering or landing. However, the need to build in this flexibility needs to be weighed against the mass penalty since power has an impact on the amount of fuel required.

The ability to manoeuvre the aerial platform is essential for meeting the requirements of a directed mission where the flight path is specified and requires the aerial platform to vary its heading. In addition, obstacle avoidance becomes a concern when flying at low altitudes; thus, the ability to steer the platform away from obstacles becomes necessary. Furthermore, the ability to land or to hover greatly adds to the value of the mission since specific areas of interest can be investigated at close range.

Complexity of the aerial platform has an impact on the risk associated with using a particular platform. Past history of failed missions, due to the degree of complexity involved in their execution, has shown that such disasters result in years of delay in acquiring the scientific data and creates a negative image for space programs. As mentioned in Chapter 2, a Viking derivative aeroshell having an internal diameter of 2.48 m will be used to deliver the aerial platforms to Mars. Therefore, the physical dimensions of each platform when stowed within the aeroshell cannot exceed 2.48 m. An aerial platform having larger dimensions will necessitate multiple folding of the structure, which leads to a high probability of failure during deployment.

Although other factors such as the size and volume of the aerial platforms could be compared, the gross takeoff mass, power, manoeuvrability and complexity were deemed to be the most important factors from a performance perspective and will form the basis for comparing the performance of the three types of aerial platforms addressed in this study. The following sections describe how each of these performance metrics is evaluated for each aerial platform.

54 4.2 Fixed-Wing Aircraft

To begin, a description of some aircraft characteristics is presented as these characteristics play an important role in evaluating the performance metrics.

4.2.1 Aircraft Characteristics

The drag force, D, which retards the motion of the aircraft through the air, and the lift force, L, which is a measure of the usefulness of the aircraft, are the primary forces that influence the performance of the aircraft and are defined as:

2 D = \pV SCD (4.1)

2 L = $pV SCL (4.2) where \pV2 is the dynamic pressure (p is the density of the surrounding fluid and V is the forward aircraft velocity), S is the wing planform area, CD is the drag coefficient and CL is the lift coefficient.

For an aircraft, the drag coefficient can be written as [61]:

2 CD = CDo + K CL (4.3)

2 where CD is the parasite (zero-lift) drag coefficient and the second term, KC L , is the induced drag coefficient. The method by which CD and K are determined is presented in Appendix C.

In this study, an attempt was made to determine the optimum flight condition for each platform so as to derive the best performance. For the fixed-wing aircraft, maximum

55 range is achieved by flying at the minimum drag condition [76]. Therefore, as a starting point, the performance of the fixed-wing aircraft will be evaluated at the minimum drag condition.

It can be shown [76] that at the minimum drag condition, CL and CD can be expressed as:

"f? (4.4)

= 2C (4.5) UMD D

Furthermore, it can be shown [76] that the velocity of the fixed-wing aircraft at the minimum drag condition is: v--m where m is the mass of the aircraft and g is the local gravity.

For flight conditions other than the minimum drag condition, the lift equation, (4.2), allows us to find a relationship between CL and V as follows:

Q=-^" (4-7) pV2S

From this, it is apparent that CL is inversely proportional to the square of the velocity. For instance, for the same lift force, if the velocity is increased by 30%, the associated

CL is decreased by 69%. Equation (4.3) can then be used to determine the associated value for CD.

56 A design decision was made to set the cruising speed as the maximum speed at which the fixed-wing aircraft can fly.Consequently , the maximum available power was limited to that required for cruising flight. Note that this design decision is subjective and may be changed depending on mission requirements or for the purpose of performance optimization. For safety reasons, a sufficient margin between the cruising speed and the stall speed should be maintained. FAA guidelines recommend a margin of 30% [84].

Therefore, the cruising speed of the fixed-wing aircraft will be set as l.Ws. Again, this design decision may be changed depending on mission requirements or for the purpose of performance optimization.

4.2.2 Method of Evaluation of Performance Metrics for Fixed-Wing Aircraft

4.2.2.1 Gross Takeoff Mass

The gross take-off mass, MTakeoff, of the fixed-wing aircraft as it begins the mission for which it was designed, can be broken down into payload mass, fuel mass, mass of the propulsion system, mass of the fuel system (including fuel tanks and piping), and the mass of the structure (including the fuselage, aerodynamic surfaces, fixed equipment, flight control system, and anything else not considered as part of payload or fuel). The equation below summarizes the takeoff mass buildup:

^Takeoff ~ "^Structure "^Payload "^ Fuel "*Prop System + *&Fuel_Syslem V*-°)

A given mission can be broken down into various mission segments or "legs", e.g., 1) warm-up and takeoff, 2) climb, 3) cruise, 4) loiter, 5) descent and 6) landing. For each of the mission concepts outlined in Chapter 2, it is assumed that the mission will comprise of a single cruise leg. For simplification, the descent will be ignored, assuming that the cruise ends with a descent and that the distance traveled during descent is part of the

57 cruise range. It is assumed that at the end of the mission, all the fuel will have been consumed. Therefore, for the simple cruise mission, the mass of the aircraft at the end of the mission will be equal to:

MFi„ai — ^Structure "* Payload "^Prop_ System ^ Fuel _ System V*-*)

A relationship between MTakeoff and MFjnal can be derived from the Breguet range equation [76]:

"Takeoff 'Hvmv C, dm R = f 'l^^^L (4.10) ML f CD mg where R is the range, TJ is the propeller efficiency, / is the specific fuel consumption, CL and CD are the lift and drag coefficients respectively, m is the mass, and g is the local gravity.

Integrating Equationii (4.10)y*.iv), , undeunucir thuiec assumptioaoouiiipuun thauiati rji/pmp, , f,j , g,g , anaiidu -^- are all CD constant, results in:

I prop ^l (M Takeoff R = In (4.11) fs [cD V MFinal J

Rearranging Equation (4.11) provides us with a relationship between MTakeoff and MFinal as follows:

RfgCD IpropC-L ivl Takeoff ivl Final K Y+-1*)

58 The range, R, to be covered by the aircraft in each mission is specified in Chapter 2. The specific fuel consumption, /, is dependent upon the power plant selected. As mentioned in Chapter 3, a hydrazine-based Akkerman engine has been selected as the propulsion system for the fixed-wing aircraft. The specific fuel consumption, /, for the hydrazine engine is 2.7 kg/kW-h and the specific power, ps , is 1.62 kW/kg [68]. The Daedalus

88 has a design-point propeller efficiency, rj , of 90% [7]. However, a conservative propeller efficiency of 85% will be used in this analysis.

When evaluating the mass of the aircraft at the minimum drag condition, the associated lift and drag coefficients, as evaluated in Equations (4.4) and (4.5), should be substituted into Equation (4.12). If the mass is being evaluated at a flight condition other than the minimum drag condition, then the lift and drag coefficients associated with the specific flight condition should be used in Equation (4.12).

Due to the interdependency between the gross takeoff mass and the power required to operate the fixed-wing aircraft, the gross takeoff mass must be determined through an iterative process. This iterative process is described in the flow chart in Figure 4.1.

The process begins by determining MFinal from Equation (4.9). The components of

MFinai are determined in the following manner:

• ^structure ' Specified for the Daedalus 88 in Table 3.1 plus the mass of the flight control system, 3 kg, as mentioned in Chapter 3. If scaling of the aircraft is

required, MStructure is determined using Equation (4.28).

• MPayload - Specified in Chapter 2 for each mission

• M?Top System - An initial guess is made for this component

# MFue, System - Assumed to be 10% of MStmcture [77]

59 GueSS Mprop System Equation (4.9) i k w

Mpinal Mprop System

Equation (4.12)

Mrakeoff

Equation (4.16)

•* Avail

Equation (4.13)

N<3

Mpr0p_System

MpropSystem (guess) - Mpr„p System (4.13) ~ 0? 1 Yes 1 Mxakeoff (4.12)

Figure 4.1 Flow Chart for Iterative Calculation of Fixed-Wing Gross Takeoff Mass

The result of this equation is then substituted into Equation (4.12) to obtain MTakeoJf.

Once MTakeoff has been obtained, the available power, PAvail, can be calculated from

Equation (4.16) presented in Section 4.2.2.2. Based on this available power, the mass of the propulsion system, MPr op _ System ' can be determined as follows:

60 ^ trap System \*-*-3) PsP,

where pSpec, is the specific power for the hydrazine engine (1.62 kW/kg).

If the value of M?rop System calculated in Equation (4.13) and the initial guess that was substituted into Equation (4.9) are close enough, then the value of MTakeoff determined in

Equation (4.12) is the appropriate gross takeoff mass for the aircraft. If not, the result of Equation (4.13) is substituted back into Equation (4.9) and the process is repeated until there is a convergence between the value of MVlop System substituted into Equation (4.9) and that obtained from Equation (4.13).

The fuel mass, MFuel, is determined as:

MFuel = MTakeoff - MFinal (4.14)

4.2.2.2 Power

The power required for cruising flight for a fixed-wing aircraft is determined by:

D-V P^ = (4-15) /prop

where D is the drag calculated from Equation (4.1), V is the cruising velocity and rjprop is the propeller efficiency (specified as 85% as mentioned previously).

Note that if the required power is being calculated at the minimum drag condition, then the velocity, V, in Equation (4.15) is the minimum drag velocity, V^, as calculated from Equation (4.6) with m = MTakeoff and the minimum drag coefficient as calculated

61 from Equation (4.5). The drag, D, in Equation (4.15) is the minimum drag as calculated from Equation (4.1) using the minimum drag coefficient and the minimum drag velocity calculated from Equations (4.5) and Equation (4.6) respectively.

For conditions other than the minimum drag condition, the drag will be calculated from Equation (4.1) using the drag coefficient as calculated from Equation (4.3) where the lift coefficient in this equation is determined from Equation (4.7).

The available power, PAvajl, is the maximum amount of power available to the aircraft for its operations and is defined as:

"Avail = ":Req "*"' *Payload "*" •* Avionics (4.16)

where, PPayload is the power required by the payload as specified in Chapter 2 for each mission, and PAvionics, assumed to be 200 W, is the power required to operate the avionics (the power consumption of an average computer is 175 W [78]). Since the cruising velocity is the maximum velocity at which the fixed-wing aircraft can fly, the available power is limited to the amount of power required to fly at the cruising condition.

4.2.2.3 Manoeuvrability

The manoeuvrability of each aerial platform is quantified by evaluating the following three parameters: • Turn radius • Range of speed • Ability to hover

The turn radius for the fixed-wing aircraft is determined as [76]:

62 V2

2 (4.17) g VAT -]

where N is the 'load factor', defined as — (L is the lift and J^is the weight), Vt is the speed in the turn, and g is the local gravity. The load factor can be shown to be a function of the bank angle, ^, as follows [76]:

JV^l + tan2^ (4.18)

In the case of the Daedalus 88, the design load factor is 1.75 in the Earth environment [7]. However, given that the gravity of Mars is lower than that of Earth, a higher load factor can be sustained on Mars when using the same structure as that used on Earth. For the purpose of this study, a conservative value of 1.75 is used; however, this design decision may be changed for the purpose of performance optimization.

The shortest radius turn is limited by the total available power since, for a given velocity, the power required to perform a turn is higher than the power required in cruising flight at the same velocity. The power required to perform a turn can be determined as follows:

2 ° 1 * (4.19)

The speed in the turn, V,, and the bank angle, ^, must be selected so that Pt does not exceed the required power, PKeq, as defined in Equation (4.15). Thus, the turn radius is calculated via an iterative process. From Equation (4.17), it is apparent that decreasing the speed in the turn and increasing the bank angle will result in the turn radius being minimized. For safety reasons, a sufficient margin between the speed in the turn and the

63 stall speed should be maintained. Therefore, the speed in the turn, Vt, is set as 15% above the stall speed. Again, this is a design decision that may be changed for the purpose of performance optimization.

Given the speed in the turn, an estimate is made of the bank angle, ^, that will result in the shortest turn radius. The associated load factor, N, is calculated from Equation

(4.18). N and Vt are then substituted into Equation (4.19) to determine the power required to perform the turn, Pt. The value of Pt should be as close as possible to the value of PRe?. If it is not, the bank angle, ^, is adjusted until Pt is as close to PReq as possible while maintaining a load factor below 1.75.

The range of speed, VRange, is defined as the difference between the absolute minimum and maximum allowable speeds at which the aerial platform can fly. As mentioned in Section 4.2.1, the absolute maximum allowable speed is set as the cruising speed

(V = 13VS) while the absolute minimum speed is set as the stall speed, Vs , which is defined as: r-l£t

Therefore, the range of speed of the fixed-wing aircraft is defined as:

VRange=\-Ws-Vs=Q3Vs (4.21)

A more conservative approach of maintaining a buffer above the stall speed as the minimum allowable speed may be taken. However, for the purpose of this study, the stall speed is set as the minimum allowable speed.

64 The ability to hover is measured based on the platform's ability to remain static over the ground in the presence of Martian winds blowing at 15 m/s. The fixed-wing aircraft has absolutely no hovering capability because it must have a velocity relative to its environment in order to generate lift.

4.2.2.4 Complexity

The complexity of the aerial platform has an impact on the risk associated with performing the mission since failures during deployment could result in failure of the entire mission. As mentioned in Chapter 2, a Viking derivative aeroshell having an outer diameter of 2.65 m and a maximum internal diameter of 2.48 m will be used to stow the platform [34]. As a result, the structure of the platform may have to be folded such its maximum span and length do not exceed 2.48 m when stowed.

The need to fold the platform structure significantly increases its complexity as well as the risks associated with deployment. Regardless of whether the platform is being deployed during descent, in which case the deployment sequence is time critical, or from the ground, failure to successfully unfold the structure will result in mission failure. Therefore, the complexity of the aerial platforms is measured based on the number of structural folds required to stow the platform in the aeroshell. The number of span-wise folds for the fixed-wing aircraft is determined as:

"van =-7- (4-22) "int where b is the span andd^ is the internal diameter of the aeroshell (2.48 m).

The number of length-wise folds for the fixed-wing aircraft is determined as:

aircraft "length j (4.23) *int

65 where laircraft is the length of the fixed-wing aircraft. The length of the Daedalus 88 is

determined from Figure 3.4. However, if scaling of the aircraft is required then laircrafl can

be determined from Equation (4.25). The total number of folds for the fixed-wing aircraft is determined as:

n total = nspcm + n length (4.24)

4.2.3 Scaling of the Fixed-Wing Aircraft

Scaling of the aircraft may be required depending on the mass of the payload to be carried and the range to be covered in each mission. For geometrically similar objects, the

linear length scale, /, is proportional to the one-third-power of the weight, WTakeoff, that is [79]:

l = KWTakeoffy (4.25)

where A: is a constant of proportionality and WTakeoff = MTakeoffg.

Using this relation, dimensions such as the wingspan, b, mean chord, c, and aircraft

length, lajrcraft, can be determined given the gross takeoff weight. The constant of proportionality for the Daedalus 88 in the Martian environment is determined as:

k = 7- (4.26) V™Takeoff Daedalus

66 where / is any dimension (wingspan, mean chord, or aircraft length) of the Daedalus in its current Earth-based design, MTakeoff Daedalus is the gross takeoff mass of the

Daedalus 88 as defined in Table 3.1, and g is the Martian gravity.

As an example, in order to scale the wingspan of the Daedalus 88 in the Martian environment, the following relation is derived:

b = 4.594(WTakeo/ry (4.27)

The structural mass is scaled as follows:

M.Takeoff M( M Structure Daedalus (4.28) \M Takeoff Daedalus J

where M Structure Daedalus is the structural mass of the Daedalus 88, defined as the empty weight in Table 3.1.

4.3 Airship

The primary lift component of lighter-than-air vehicles is derived from the buoyancy of the lifting gas. In the case of airships, control surfaces incorporated into the design contribute a small fraction of the lift and provide directional controllability. The manner in which the buoyancy force as well as the density of the lifting gas (hydrogen) and the density of the Martian atmosphere (comprised primarily of carbon dioxide) are calculated is presented in Appendix D.

67 4.3.1 Method of Evaluation of Performance Metrics for the Airship

4.3.1.1 Gross Takeoff Mass

The gross takeoff mass of the airship can be broken down into payload mass, mass of the lifting gas (H2), mass of the gasbag envelope, mass of the propulsion system, fuel mass, mass of the fuel system (including fuel tanks and piping), and the mass of the structure (including inflatable tail, rigging, rudder, fins and miscellaneous structural components). The gross takeoff mass usually also includes the mass of the gasbag inflation system. However, since this equipment is jettisoned shortly after the airship has been inflated, it is not accounted for in this study. The equation below summarizes the takeoff mass buildup:

M M M M M M M M 4 2 Takeoff = Structure + Payload + H2 + Gasbag + VroP_Sys,em + Fuel + Fuel _Sys,em ( " 9)

The gross takeoff mass is determined through an iterative process that is described in the flow chart in Figure 4.2. Based on an initial guess ofMTakeoff, the values of MSlruclure,

M M M M H^ Gasbag Pr op _ System, Fuel > **& MFml System, determined from the equations below, and the value of MPayload determined from Chapter 2 for each mission, are substituted into Equation (4.29) and the equation is solved for MTakeoff. If the difference between the initial guess and the calculated value of MTakeoff does not equal zero, then a new value of MTakeoff is selected and the process is repeated until there is convergence.

The mass of the hydrogen, MHi, required to lift the vehicle is calculated as:

MHl=pHvHi (4.30)

where pH is the hydrogen density as specified in Table D. 1 for each mission, and vH is the volume of gas required to lift the airship and is determined as:

68 Guess MTakeoff 1

" ir w Equation (4.31) Equation (4.37) Equation (4.38)

MFueLSystem ^Structure vHl

^ Equation (4.40) Equation (4.32) ' f Moasbag Equation (4.30) PReq

ir Mpayioad Equation (4.16) MH2 lvlPayload

* Avail Equation (4.13)

i ' Equation (4.35)

Mpmp System MFuei

' Equation (4.29)

Mrakeoff '

Mjakeoff (gUeSS) - M akeoff (4.29) ~ 0? No T

Yes 1 Mxakeoff_Guess = Mxakeoff

Figure 4.2 Flow Chart for Iterative Calculation of Airship Gross Takeoff Mass

69 Mrake°ff (4.31) Pco7 ~ PH7

The mass of the gasbag, MGasBag, is calculated as:

MGasBag=Aa (4.32) where A is the surface area of the airship and a is the density of the composite gasbag material taken to be 32 g/m2 [18]. Since the shape of the Robotic Martian Airship is that of a prolate spheroid, the surface area, A, is determined as:

A = 2nc2 + In—sin"1 e (4.33) £

a,rs ip mrs ip where a - and c - . The values of lairship and dairship are calculated from

Equation (4.42) and Equation (4.43) respectively. The eccentricity, s, is defined as:

/ 2 2 s = (4.34) a

The mass of the propulsion system, M?ropulsion Syslem, is calculated in the same manner as the fixed-wing aircraft using Equation (4.13), where PAvail is determined from Equation

(4.16) using PReq from Equation (4.40) presented in Section 4.3.1.2. pSpec in Equation

(4.13) is the specific power of the propulsion system. As mentioned in Chapter 3, a hydrazine-based Akkerman engine will be used to power the airship. Therefore, ps for the case of the airship is also 1.62 kW/kg [68].

The mass of the fuel, MFuel, is calculated as:

70 MFuel=ftPAvail (4.35) where, /, the specific fuel consumption for the hydrazine engine, is 2.7 kg/kW-h, and t is the mission duration determined as:

t = y (4.36)

where R is the range, specified in Chapter 2 for each mission, and V is the velocity of the airship. From the Viking mission, normal wind speed peaked at 10 m/s, with a daily average of 5 m/s [18]. During a dust storm, winds can reach 14 - 32 m/s. Therefore, a cruising speed, V, of 15 m/s has been selected in order to ensure that the airship can maintain its heading under moderately windy conditions while minimizing power requirements. This design decision with respect to setting the cruising speed at 15 m/s may be changed for the purpose of performance optimization.

The mass of the fuel system, MFuel System, is determined by scaling the mass of the RMA fuel system, MFuel Syslem RMA, relative to the gross takeoff mass of the airship, MTakeoff, as follows:

™*Fuel ^System ,, "*Fuel _System RMA (4-J') M Takeoff RMA

M Takeoff RMA ' me gross takeoff mass of the RMA, is approximately 200 kg and

^Fuei_system_RMA is 2.60 kg as provided in Table 3.2 and Table 3.3 respectively

The structural mass, MStruclure, is determined by scaling the structural mass of the RMA,

^structure RMA > relative to the gross takeoff mass of the airship, MTakeoff, as follows:

71 ^Takeoff "*• Structure ~ ~7Z "^ Structure _ RMA (4.J0) M Takeoff RMA

where MStructure_RMA is 32.7 kg as provided in Table 3.3.

4.3.1.2 Power

The power required for cruising flight for the airship is determined in the same manner as the fixed-wing aircraft using Equation (4.15), where the drag, D, is determined as:

2 D = \pV ACD (4.39) where p is the density of the atmosphere, V is the cruising velocity, A is the reference area, and CD is the drag coefficient. The reference area for a buoyant craft is typically

273 chosen to be (Buoyant Volume) , or vH * [80]. Therefore, the power required for cruising flight for the airship becomes:

\pv\H\cD P», = 2 2 (4-40) /prop

The drag coefficient, CD, is determined by extrapolating empirical data gathered on other airships on the basis of thickness ratio, —, where d is the maximum diameter and / is the length of the airship [80]. Based on the RMA design, a thickness ratio of 0.5 is used. The drag of the airship can be broken down into two components; the drag associated with the bare hull (representing 80% of the total drag), and the drag associated with the fins and rudder (representing 20% of the total drag) [80]. The drag coefficient associated with the bare hull, which is extrapolated from the data provided in Reference 80 based on

72 the thickness ratio, is approximately 0.026. Therefore, the total drag coefficient, CD, is estimated here to be 0.033.

The propeller efficiency, tj^ , is assumed to be 85% as with the fixed-wing aircraft.

A design decision was made to limit the available power to the amount of power required to fly at the cruising speed; however, this design decision may be changed based on mission requirements.

4.3.1.3 Manoeuvrability

The turn radius for the airship can be determined using the 6-DOF nonlinear equations of motion that incorporate the flight mechanics, aerostatics and complete aerodynamic effects of airships [81]. The 6-DOF nonlinear equations of motion can be simplified for the case of a steady turn (shown in Figure 4.3), to the following two nonlinear algebraic equations [81]:

- mur + FAD (u, v, r) + Fc (u, v, r,S)-0 (4.41 a)

- mxgur + MAD (u, v, r) + Mc (u, v, r,S) = 0 (4.41 b)

The first terms in Equations (4.41 a) and (4.41 b) are inertial terms where m, u, r and

xg represent the mass of the airship, the forward speed, the yaw rate and the x-coordinate of the airship's center of gravity in the body frame respectively. The second terms are the aerodynamic force and moment terms respectively, including added-mass terms, viscous effects, forces on the fins, and forces on the hull due to fins. The third terms are the control force and moment terms respectively, u, v, and 8 represent the velocity in the x direction, velocity in the y direction and the rudder deflection (positive trailing edge left) respectively.

73 Figure 4.3 Airship in Steady Turn [81]

As mentioned in Section 4.2.2.3, the range of speed, VRange, is the difference between the absolute minimum and maximum allowable speed at which the aerial platform can fly. For the airship, the maximum allowable speed is set as the cruising speed (15 m/s). Unlike the fixed-wing aircraft that must have a velocity relative to the environment in order to generate lift, airships have the ability to remain aloft without forward motion. Therefore, the airship can maintain a minimum ground speed of 0 m/s for wind speeds up to 15 m/s resulting in VRange =\5mls .

As mentioned in Section 4.2.2.3, the ability to hover is measured based on the platform's ability to remain static over the ground in the presence of Martian winds blowing at 15 m/s. The airship, therefore, has the ability to hover.

74 4.3.1.4 Complexity

As mentioned in Section 4.2.2.4, the complexity of the aerial platform has an impact on the risk associated with performing the mission. The need to fold the platform structure significantly increases its complexity as well as the risks associated with deployment. The complexity of the aerial platforms is measured based on the number of structural folds required to stow the platform in the aeroshell.

The airship being considered in this study is a non-rigid airship that uses gas pressure for both shape maintenance and structure thereby eliminating the need for structural hardware for the purpose of deployment. Even the tail of the airship is inflatable. The mass of the structural components of the RMA represent only 16% of its gross mass, as indicated by the mass breakdown in Table 3.3, and do not require folding for launch purposes. The deployment scenario envisioned here is much like that shown in Figure 3.1.

4.3.2 Scaling of the Airship

Depending on the mass of the payload to be carried and the range to be covered in each mission, the size of the airship will vary. The length of the airship, lairshi , is scaled in relation to the volume per the following relation:

vn, 'airship ~ \\ '•RMA (4.42)

where vH is the volume of the airship determined per Equation (4.31), vRMA and lRMA are the volume and length of the RMA respectively, defined in Table 3.2. Given a fineness ratio of 2, the diameter of the airship is determined as:

75 j airship airship « (4.43)

As mentioned earlier, Equations (4.42) and (4.43) are used to calculate the surface area, A, of the airship, which in turn is used to calculate the mass of the gasbag envelope.

4.4 Rotary-Wing Aircraft

AAA Method of Evaluation of Performance Metrics for the Rotary- Wing Aircraft

4.4.1.1 Gross Takeoff Mass

As with the fixed-wing aircraft, the overall take-off mass, MTakeoff, of the rotary-wing aircraft can be broken down into payload mass, fuel mass, mass of the propulsion system, mass of the fuel system (including fuel tanks), and the mass of the structure (including the body, aerodynamic surfaces, fixed equipment, piping, avionics, and anything else not considered as part of payload or fuel). The expression summarizing the mass buildup is provided in Equation (4.8).

Again, due to the interdependency between the gross takeoff mass and the power required to operate the rotary-wing aircraft, the gross takeoff mass must be determined through an iterative process. This iterative process is described in the flow chart in Figure 4.4.

76 GueSS Miakeoff Equation (4.61) "1 Equation (4.44) Req

Equation (4.16) MStructure

Equation (4.45) * Avail

M_ Equation (4.46)

Mpr0p Systei

M„ Equation (4.47)

Equation (4.48) Equation (4.49) M,'Fuel

MFud MFuel Syste,

Equation (4.8) M Payload

MTakeoff7 + Mxakeoff (gUeSS) - M keoff (4.8) ~ 0? No Ta 1 Yes

M Takeoff Guess — Mxakeoff

Figure 4.4 Flow Chart for Iterative Calculation of Rotary-Wing Gross Takeoff Mass

77 The process begins by making an initial guess of the gross takeoff mass, MTakeoff. This initial guess of MTakeoff is then used to determine the structural mass, MStruaure as well as the power required to operate the rotorcraft, PR . The structural mass, including the fuselage, blades, hub, control group, drive system, landing gear and avionics, is determined by scaling the structural mass of the MARV relative to the gross takeoff mass of the rotorcraft, MTakeoff, as follows:

M = T-^l M C4 44) lv* Structure •.. lyl Structure_MARV y+.i^) M Takeoff MARV

where MTakeoff MARV is the gross takeoff mass of the MARV baseline rotorcraft model, and MStructure_mRV is the structural mass of MARV. MTakeoffMARV and MStnclm_lURr are

50 kg and 26.7 kg respectively as provided in Table 3.5.

^Takeoff is also used in Equation (4.61), presented in Section 4.4.1.2, to determine the power required, PKeq. The result from Equation (4.61) is substituted into Equation (4.16) in order to determine the available power, PAvail, which is then used to determine the mass of the propulsion system, MProp_Syslem.

MARV's power plant consists of a fuel cell system that supplies power to an electric motor that rotates the rotor. The mass of the propulsion system, therefore, includes the mass of the fuel cell as well as the motor. The mass of the propulsion system to be used on the rotary-wing aircraft is determined by scaling the mass of MARV's fuel cell,

MFC an MARV > d MARV's motor, MMgtor MARV based on the available power as follows:

= + M-pvop_System "^ V** FC_MARV ^Motor MARV ) (4-45) "Avail MARV

78 where PAvail is the available power as determined from Equation (4.16), and PA] the available power for the MARV mission specified as 4.63 kW [27]. MFC MiARV is 5.6 kg as specified in Table 3.5 and MMotor mRV is 2 kg [27].

The fuel cell uses pure hydrogen and pure oxygen as fuel. The amount of hydrogen required for the mission can be determined from the following relation [82]:

tP M = Avail (446) 2 >lfcLHVH2

where, rjfc is the thermodynamic efficiency of the fuel cell, and LHVH is the lower heating value of hydrogen. The baseline thermodynamic efficiency of the PEM fuel cell is 45% [82] and the lower heating value of hydrogen is 120.1 MJ/kg [83].

Given that the overall chemical reaction in a hydrogen/oxygen fuel cell is

H2 +\02 - H20,a single oxygen atom is required for every two hydrogen atoms. Since the molecular mass of oxygen is 16 g/mol, and only half the number of oxygen atoms is required in comparison to hydrogen atoms, the total mass of oxygen, M0 , required for the mission is eight times the mass of hydrogen required for the mission. Therefore:

M02=8(MH2) (4.47)

The mass of the fuel system, MFuel, therefore, is determined as:

MFM=MHi+M02 (4.48)

The hydrogen and oxygen are stored in their respective storage tanks at cryogenic temperatures. The mass of the fuel system, MFuel System, is determined by scaling the mass of the MARV fuel system based on the amount of fuel required as follows:

79 MF«el M Fuel _ System M Fuel _ System MARV (4.49) \MFuel_MARV J

where M Fuel MARV is the mass of the fuel required for the MARV mission, specified as

2.06 kg [27] and M^^ is the mass of MARV's fuel system, specified as 2.94 kg [27].

Based on the initial guess of MTakeoff, the values of MFuel, M?ropulsionSystem, MFuelSystem, and MStructure determined from the equations above and MPayload provided in Chapter 2 for each mission are then substituted back into Equation (4.8). If the result of MTakeoff from Equation (4.8) does not match the initial guess, then the resulting value from

Equation (4.8) is used as the new estimate of MTakeoff. The process is then repeated until there is convergence between the estimated value of MTakeoff and that obtained from Equation (4.8).

4.4.1.2 Power

The power required to operate the rotorcraft in forward flight is composed of three parts: the parasite power, PPara, required to move the rotary-wing aircraft through the air; the induced power, Plnd, required to produce the rotor thrust; and the profile power, PPro, required to turn the rotor blades through the air.

Parasite Power

The parasite power, required to move the rotary-wing aircraft through the air, is calculated as [84]:

3 PPara-DV = ^2pV F (4.50)

80 where D is the drag force and F is the flat plate area. The flat plate area of MARV attributable to the fuselage body is 0.2 m2 [27].

Induced Power

The induced power, required to produce the rotor thrust, is calculated as [84]:

PInd=Tw (4.51)

where T is the thrust or weight of the rotorcraft, and w is the downwash velocity. PInd obtained from Glauert's hypothesis, Equation (4.51), represents an unattainable minimum value, which, based on experience, should be increased by approximately 15% [84].

Generally, the downwash velocity, w, is obtained iteratively by assuming initially that w is equal to the value for hovering, w0, expressed as follows [84]:

w--te (452)

where Ar is the rotor disk area. The first iteration is substituted into the following equation:

V = ^{w-Vsma)2+{Vcosa)2 (4.53) where a is the angle-of-attack of the rotor plane and V is the forward velocity. In steady level flight the sum of the forces acting on the rotorcraft in the direction of flight leads to:

D+ T since = 0 (4.54)

81 Therefore, the angle-of-attack is defined as:

a = sin-1f^| (4.55)

where D = —1 pV^2 , F and T is the thrust or weight of the rotorcraft.

V is then substituted into the following equation to determine the next iteration of w.

w = T - (4.56) 2pArV

This value is then substituted back into Equation (4.53) and the process is repeated until the value of w converges.

Profile Power

The profile power, required to turn the rotor blades through the air is calculated as [84]:

3 2 PPvo=pArVT ^(l + 3M ) (4.57)

where Ar is the rotor disc area, VT is the tip speed defined as a>R where co is the rotor rotational speed and R is the rotor radius, a is the rotor solidity (the ratio of the lifting area of the blade to the area of the rotor), Cd is the drag coefficient, and // is the ratio between the forward speed and the rotor tip speed rv\ For the purpose of this study, a KVTJ design decision was made to maintain the rotor solidity, a, and the tip speed, VT, at the same value as that of MARV, 0.25 and 143.75 m/s respectively [27]. Again, these values may be changed for the purpose of performance optimization.

82 To determine the drag coefficient, the characteristics of the FX 63-137 airfoil having a maximum thickness of close to 15% of the chord and a maximum camber of 6% of the chord [69], which is similar to the AGRC 1506 airfoil used on MARV, were used. Based on these characteristics, the drag coefficient can be determined from the following equation:

2 Cd =0.009 + 0.004C; (4.58)

Ct is determined from the following equation [84]:

6CT c>=v^n <4'59) 1 + V 2 ,

where CT, the thrust coefficient, is defined as:

CT= —-i (4-60) pArVT

Power Required

Thus, the final expression for the total power required, PReg, for forward flight of the rotorcraft is:

3 3 2 PReq =-pV F + Ll5Tw+pArVT -^(l + 3M ) (4.61) 2'

Figure 4.5 illustrates the variation with speed of the power required by the rotary-wing aircraft in level flight [85]. From this diagram it can be seen that the induced power,

83 PInd, is the largest component in hover and that it decreases quickly with speed. The profile power, P?TO, exhibits a slight increase with speed. The parasite power, PPara, is negligible at low speeds but increases proportionally to V3 to dominate at high speed. Thus the total power required is high at hover, has a minimum value in the middle of the rotorcraft speed range, and then increases again at high speed because of the parasite power.

Power Required Parasite

Speed

Figure 4.5 Power vs. Speed Profile for Rotary-Wing Aircraft [85]

As with the fixed-wing aircraft, an optimum flight condition was sought in order to derive the best performance. This optimum flight condition was selected as the point at which the power consumption of the rotorcraft in level flight is at a minimum. Again, an alternate design choice for the cruising condition may be selected based on mission requirements. Therefore, for each mission, Equation (4.61) was evaluated for a range of

84 speeds in order to determine the cruising velocity, Vp , corresponding to the minimum power consumption of the rotorcraft in level flight.

It can also be seen from Figure 4.5 that the power required to hover is significantly higher than the power required to fly at the cruising speed, Vp . Since hovering is not a requirement in any of the missions outlined in Chapter 2, to minimize the gross takeoff mass of the rotary-wing aircraft, the available power is limited to the amount required to enable the rotorcraft to fly at 30% above or below the cruising velocity, Vp . Again an alternate design choice may be selected depending on mission requirements. The maximum available power, PAvajl, is calculated from Equation (4.16) using PK from

Equation (4.61).

4.4.1.3 Manoeuvrability

The turn radius for the rotary-wing aircraft is determined as follows [86]:

V2 r = —'-— (4.62) gtan^ where V, is the speed in the turn, g is the local gravity and ^ is the bank angle.

For the rotary-wing aircraft, the turn radius is limited by the total amount of available power, PAvail, and is determined through an iterative process. An initial estimate is made of the bank angle. Given the bank angle, the load factor is determined from the following relation:

N = Vl + tan2^ (4.63)

85 Given the load factor, the lift force during the turn is determined from the following relation:

T = Nmg (4.64)

Given the lift force during the turn, the associated power is calculated using Equation (4.61). The resulting power should not exceed the maximum available power for propulsion, PReq. Once the maximum bank angle has been determined, Equation (4.62) is used to calculate the turn radius. In order to maximize the bank angle, Vt is selected as the velocity associated with the minimum power consumption during straight and level

flight, VP . * mm

The range of speed is evaluated as the difference between the absolute minimum and maximum allowable speed at which each aerial platform can fly. As mentioned earlier, in order to minimize the gross takeoff mass of the rotary-wing aircraft, the available power is limited to the amount required to enable the rotorcraft to fly at 30% above or below the

cruising velocity, Vp . Therefore, the range of speed, VRange, of the rotary-wing aircraft

is defined as:

V [ Rame=' -Wp ~0.7Vp = 0.6F„ (4.65) Range fM„ Fm„ fmin V /

The ability to hover is measured based on the platform's ability to remain static over the ground in the presence of Martian winds blowing at 15 m/s. As mentioned earlier, for the purpose of minimizing the gross takeoff mass, the available power is not sufficient to allow the rotary-wing aircraft to hover.

86 4.4.1.4 Complexity

As mentioned in Section 4.2.2.4, the complexity of the aerial platform is measured based on the number of structural folds required to stow the platform in the aeroshell. The number of span-wise folds required for the rotor is determined as:

n««=^T- (4-66)

where drotor is the diameter of the rotor andofint is the internal diameter of the aeroshell (2.48 m).

The number of chord-wise folds required for the rotor is determined as:

n***=^f- (467) a-rat.

where crotor is the maximum chord length of the rotor blade.

Since there are two rotors, the total number of folds for the rotary-wing aircraft is determined as:

nto,ai = 2(nspan + "chord ) (4-68)

87 4.4.2 Scaling of the Rotary-Wing Aircraft

Depending on the mass of the payload to be carried and the range to be covered in each mission, the size of the rotary-wing aircraft will vary. Regression analysis indicates that the rotor diameter, drotor, scales as [87]:

dmlor=k(MTakeoffr (4.69)

Using the rotor diameter and the gross takeoff mass of MARV, we find that k = 0.891. and so

0.4 drolor=o.m(MTakeoff) (4.70)

The chord of the rotary-wing aircraft, crolor, is determined by scaling the maximum chord of MARV based on the rotor diameter as follows:

'rotor MARV (4.71) \" rotor _MARV J

where drotor MARV is the rotor diameter of the MARV baseline rotorcraft model, and

crotor_MARv & ™e maximum chord length of the MARV rotor. drotorMARV and cmk>r_iaRr are 4.266 m and 0.67 m respectively [27].

The flat plate area, F, will also vary as the size of the aircraft changes. The flat plate is determined by scaling the MARV flat plate area relative to the square of the rotor diameter as follows:

F = Ft, (4.72) \^" rotor MARV J 2 where FMARV is 0.2 m [27].

4.5 Method for Evaluating a Single Performance Rating for each Aerial Platform

The four performance metrics discussed above serve as the common basis for comparing the performance of each aerial platform. For each scientific mission, the performance metrics will be used to calculate an overall performance rating, QPlatform , for each aerial platform and this rating will be used to determine which platform is best suited for each mission.

The performance of each platform relative to each performance metric is denoted Qk where k : 1 -> 4 and represents each performance metric:

1 = Mass 2 = Power 3 = Manoeuvrability 4 = Complexity

Qk can be a number between 0 and 10 whereby a lower value indicates poor performance and a higher value indicates best performance. The method by which the value of Qk is determined for each metric will be discussed later.

The weight assigned to the importance of each performance metric in achieving the mission requirements is denoted Wk. Wk can be a number between 0 and 10 whereby

^_Wk - 10. A lower value of Wk indicates that the metric is less important in achieving the mission requirements and a higher value indicates that the metric is more important in

89 achieving mission requirements. The method by which the value of Wk is determined for each metric will also be discussed later.

The overall performance rating, QPlatform , is then determined as:

QPlatform ~ V"1 nr (4-73) iwk which will yield a value between 0 and 10. The platform having the highest value of Qpiai/orm is deemed best suited for the mission.

A given performance metric, Qk, may be dependent upon various influencing factors. For instance, the factors influencing the manoeuvrability of an aerial platform include the turn radius, the range of speed and the ability to hover. In this case, Qk is determined through the same methodology as described above for QPlalform resulting in the following relation:

Qk=^ (4.74)

where qm (a number between 0 and 10) denotes the performance of the platform with respect to the influencing factor m, and wm denotes the weight assigned to the importance of each influencing factor in achieving the mission requirements. wk can also be a number between 0 and 10 whereby ^wk = 10. n denotes the number of influencing factors associated with the performance metric, k. The method by which the value of qk and wk are determined for each influencing factor will be discussed later.

90 In the case where there is only a single influencing factor, as is the case with the gross takeoff mass, power and complexity, then wl = 10 and Qk =qx.

4.5.1 Performance Rating with respect to Mass (Qx)

The performance rating relating to mass is calculated as follows: for a given mission, the platform having the lowest gross takeoff mass is assigned a value of 0, = 10 ; the value assigned to the other platforms is determined as follows:

( M ^

a =io ly± Takeoff _Best (4.75)

V^ Takeoff Platform J where, Qx is the performance rating of the platform being evaluated with respect to mass,

M is the oss Takeoff _BeSt & takeoffmass of the best case, and MTakeoff _Platform is the gross takeoff mass of the platform being evaluated.

4.5.2 Performance Rating with respect to Power (Q2)

The performance rating relating to power is calculated in a manner similar to the mass performance rating: for a given mission, the platform having the lowest available power,

s PAvau' i assigned a value of Q2 -10; the value assigned to the other platforms is determined as follows:

( P \ e =io Avail _ Best 2 P (4.76) \ Avail _ Platform J

where, Q2 is the performance rating of the platform being evaluated with respect to power, PAvail Best is the available power of the best case, and PAmil Plalform is the available power of the platform being evaluated.

91 4.5.3 Performance Rating with respect to Manoeuvrability (£>,)

The performance of an aerial platform with respect to manoeuvrability, Q3, is affected by three influencing factors:

1 = Turn Radius 2 = Range of Speed 3 = Ability to Hover

4.5.3.1 Performance Rating with respect to Turn Radius (q{)

For a given mission, the platform having the smallest turn radius, r, is assigned a value

of qx = 10; the value assigned to the other platforms is determined as follows:

( rBesl \ qx =10 \^ Platform J (4.77)

where, qx is the performance rating of the platform being evaluated with respect to the turn radius, rBest is the turn radius of the best case , and rplalform is the turn radius of the platform being evaluated.

4.5.3.2 Performance Rating with respect to Range of Speed (q2)

For a given mission, the platform having the greatest range of speed, Vmnge, is assigned a value of q2 = 10; the value assigned to the other platforms is determined as follows:

(y >\ Range _ Platform 02=10 v \ Range _ Best J (4.78)

92 where, q2 is the performance rating of the platform being evaluated with respect to the range of speed, VRange plalform is the range of speed of the platform being evaluated, and

VRange Best *s me range of speed of the best case.

It is important to note that in this study, the range of speed is defined as the absolute minimum and maximum speeds at which the platforms can fly. However, for some missions the requirements may be such that a given range of speed is preferred. In that case, the performance of the platforms with respect to the range of speed will have to be evaluated based on different criterion.

4.5.3.3 Performance Rating with respect to Ability to Hover (q3)

For a given mission, the platform having the ability to hover up to wind speeds of 15 m/s is assigned a value of q3 =10 and the platform that does not have the ability to hover is assigned a value of

4.5.4 Performance Rating with respect to Complexity (<94)

The performance rating relating to complexity was evaluated using a single influencing factor: the number of structural folds. For a given mission, the performance rating with respect to complexity, Q4, is determined as:

( n ^ a =10 . ,l Platform (4.79) y " Platform _ Most J

where, nplalform is the total number of structural folds required for a given platform, and

nPlatform Most l& tne total number of structural folds of the platform requiring the most number of folds.

93 4.5.5 Weighting of Performance Metrics

The weight assigned to each performance metric, Wk, and each influencing factor, wm, is dependent upon its relative importance in achieving the mission requirements and is, therefore, mission dependent. For each mission, the selection of the weights was performed empirically and the justification for each selection is provided in the following sections. Table 4.2 provides the weights assigned to each performance metric and each influencing factor for each of the three missions.

Mission 1 Mission 2 Mission 3

Mass (Wx) 3 4 1

Power (w2) 2 1 1

Manoeuvrability (W3) 1 1 4 Turn Radius (w,) 5 5 4

Range of Speed (w2) 5 5 3

Ability to Hover (w3) 0 0 3

Complexity (W4) 4 4 4 4 10 10 10 Total Y,W, i=\

Table 4.2 Weighting of Performance Metrics for each Mission

4.5.5.1 Mission 1

Mission 1 is a mid-range mission (500 km) whereby the aerial platform is expected to fly at the Martian datum at a relatively constant speed carrying a heavy payload. The area of Mars being observed, Chryse Planitia, is approximately 3 to 5 km below the Martian datum and is relatively flat with the exception of some craters within the 250 km radius.

Since this is a mid-range mission, both the mass and the power of the platform are equally important. However, due to the interdependence between mass and power, the

94 power required ultimately impacts the overall mass of the platform. Therefore, the mass metric is given greater importance.

For Mission 1, a high degree of manoeuvrability for the purpose of obstacle avoidance is not required since the platform will be flying at an altitude of at least 3 km above ground level. For that reason the manoeuvrability metric has the lowest weighting. The platform will be expected to make two 120-degree turns in order to complete the flight path. The factors influencing the platform's ability to complete the flight path are its ability to change its heading efficiently, which is measured by the turn radius, and the platform's range of speed. The range of speed is important since the platform may have to reduce its cruising speed in order to perform the turns due to the limitation in the available power. These two influencing factors are, therefore, weighted equally. Given that for Mission 1 hovering is not a requirement in order to successfully complete the mission, this influencing factor has a weighting of zero.

For each mission, the complexity of the platforms has a significant impact on the risk associated with deployment. Since failure to successfully deploy the platforms would result in mission failure, this metric has been weighted highest in each mission.

4.5.5.2 Mission 2

Mission 2 is a long-range mission (1000 km) whereby the aerial platform is expected to fly at 3 km above the Martian datum at a relatively constant speed carrying a light payload. The area of Mars being observed, Chasma Boreale, ranges in altitude between the Martian datum and 3 to 5 km below the Martian datum.

Since this is a long-range mission, concerns with respect to the amount of propellant required for the mission as well as the mass of the fuel storage tanks are much greater than in Mission 1. Therefore, the mass metric is weighted highest for this mission. Again, due to the interdependence between mass and power, the power required ultimately

95 impacts the overall mass of the platform. Therefore, the mass metric is given greater importance.

As with Mission 1, a high degree of manoeuvrability for the purpose of obstacle avoidance is not required since the platform will be flying at an altitude of at least 3 km above ground level. Therefore, the weighting of the metric as well as its influencing factors is the same as in Mission 1.

4.5.5.3 Mission 3

Mission 3 is a short-range mission (150 km) whereby the aerial platform is expected to fly at the Martian datum at a relatively constant speed carrying a light payload. The area of Mars being observed, Valles Marineris, ranges in altitude between 6 km above the Martian datum and 3 to 5 km below the Martian datum.

Since this is a short-range mission, concerns with respect to the mass and power required are low especially because the payload to be carried is relatively light. Therefore, these two metrics have been weighted lowest for this mission.

On the other hand, since the platform will be flying close to the walls of the canyon, obstacle avoidance becomes a concern. Therefore, a high degree of manoeuvrability is required for this mission resulting in a high weighting for this metric. The need for obstacle avoidance may require the platform to either fly slowly or remain static above the ground in order to assess the safest flight path. In this case, both the range of speed and the ability to hover become important factors and are, therefore, weighted equally. The ability of the platform to change its heading quickly becomes especially important if the platform does not have the ability to hover. Therefore, this influencing factor is weighted higher.

96 Chapter 5

DISCUSSION OF RESULTS

The performance metrics identified in the previous chapter were evaluated for each aerial platform in each mission. The following sections present the results of the evaluation. In addition, this chapter also presents the evaluation of the performance rating indicating which aerial platform is best suited for each mission.

5.1 Fixed-Wing Aircraft

With a view to optimize its performance, as a starting point, the fixed-wing aircraft was evaluated at the minimum drag condition, as described in Section 4.2.1, since maximum range is achieved at this condition. For each mission, assuming an aircraft having the same physical dimensions as the Daedalus 88 (i.e., wingspan of 34.1 m and chord length of 0.91 m), the minimum drag velocity resulted in a Reynolds number in the range of 3 x 104. The Daedalus 88 airfoil (DAE 31) was designed for a Reynolds number of 2.5 x 105 [7] and, therefore, needed to be replaced before proceeding further.

The search for a more suitable airfoil for Mars applications resulted in the E205B-PT (Reed) airfoil, designed for a Reynolds number as low as 4 x 104, being selected for operation on Mars [69]. It was assumed that the CD and K of the entire aircraft are not changed significantly as a result of this change in airfoil. The E205B-PT airfoil has a maximum lift coefficient of 1.0 at Re = 4 x 104. However, the minimum drag lift coefficient for each mission was found to be 1.17. This resulted in the minimum drag speed being lower than the stall speed as detailed in Table 5.2. Furthermore, as specified in Chapter 4, for reasons of safety, the cruising speed should be 30% above the stall speed. Therefore, the minimum drag condition was not a feasible flight condition for this scenario.

97 Re-evaluating the same aircraft (i.e., having the same physical dimensions as the Daedalus 88 and incorporating the E205B-PT airfoil) at a cruising speed of 13VS resulted in a Reynolds number in the range of 4x 104 and an operating lift coefficient of approximately 0.6. Two things are evident from this result: 1) the use of the E205-PT airfoil is suitable for Mars applications, 2) with the cruising speed fixed at l.3Vs, the dimensions of the aircraft cannot be scaled down since doing so would further decrease the Reynolds number and no other airfoils were found that could operate effectively at Reynolds numbers below 4x 104. Table 5.1 presents a summary of the fixed-wing aircraft characteristics that are common to both the minimum drag condition as well as the actual flight condition.

AR 37.51 Wing Area (mz) 31 Span(m) 34.1 Chord (m) 0.91 K 0.0121 CDO 0.0165

Table 5.1 Common Fixed-Wing Aircraft Characteristics

Table 5.2 presents the fixed-wing aircraft characteristics that are specific to the minimum drag condition and the actual flight condition.

98 Minimum Drag Characteristics Actual Flight Characteristics Mission 1 Mission 2 Mission 3 Mission 1 Mission 2 Mission 3 1.17 1.17 0.60 0.62 0.60 cL 1.17 CD 0.033 0.033 0.033 0.021 0.021 0.021 Reynolds Number 33204 25943 28772 46837 36789 40635 Density (kg/m3) 0.01245 0.00944 0.01245 0.01245 0.00944 0.01245 MPavload (kg) 17.7 4.94 5.35 17.7 4.94 5.35 MFuei(kg) 2.65 4.2 0.6 3.3 5.0 0.75 System (kg) 0.35 0.3 0.24 0.42 0.35 0.3 Mpuel System (kg) 3.4 3.4 3.4 3.4 3.4 3.4 Mstructure (kg) 34.1 34.1 34.1 34.1 34.1 34.1

MT„tal(kg) 58.2 46.9 43.7 58.9 47.8 43.9 PRequired (kW) 0.22 0.19 0.15 0.4 0.34 0.26 Ppayload (kW) 0.074 0.020 0.021 0.074 0.02 0.021 ^Avionics 0.20 0.20 0.20 0.2 0.2 0.2 PTotal(kW(kW) ) 0.49 0.41 0.37 0.67 0.56 0.48 V(m/s) 31.1 32.0 26.9 43.8 45.4 38 Vstaii (m/s) 33.5 34.6 29.1 33.7 34.9 29.1

Table 5.2 Specific Fixed-Wing Aircraft Characteristics

In comparison to the minimum drag condition, an increase in the total power of approximately 30% was observed when evaluating the performance of the aircraft at the actual flight condition. However, the associated increase in the gross takeoff mass was less than 1%. This can be attributed to the high specific power of the hydrazine engine resulting in a negligible increase in the mass of the propulsion system.

The gross takeoff mass and power required by the aircraft in Mission 2 is lower than that of Mission 1 despite the range being double in Mission 2. This is mainly because the payload being carried in Mission 2 is approximately 70% lighted than that of Mission 1. In addition, the aircraft in Mission 2 is flying at a higher altitude where the atmospheric density is lower resulting in decreased drag and thus a reduced power requirement.

The shortest radius turn, r, is limited by the total available power since the power required to perform a turn is higher than the power required in cruising flight at the same velocity. Since the aircraft was designed such that maximum available power is required for cruising flight, the speed of the aircraft in

99 the turn, Vt, had to be lower than the cruising speed, V. To obtain the shortest turn radius, the speed in the turn was minimized to \.\5VStall, maintaining a 15% margin above the stall speed for safety reasons. The factors associated with the manoeuvrability of the fixed-wing aircraft for each mission are summarized in Table 5.3. For each mission, the maximum available power was used to perform the turns. It can be seen from Table 5.3 that the associated load factor for each mission scenario is well below the design load factor of 1.75, indicating that the available power is more of a limiting factor on the turn radius than the load factor.

Mission 1 Mission 2 Mission 3 Turn Velocity (m/s) 38.8 40.1 33.6 Bank Angle (deg) 42 42 43 Load Factor (g's) 1.35 1.35 1.37 Total Power in Turn (kW) 0.67 0.56 0.48 Turn Radius (m) 448 479 324 Range of Speed (m/s) 10.1 10.5 8.8 Ability to Hover None None None

Table 5.3 Manoeuvrability Factors for Fixed-Wing Aircraft

As mentioned above, the dimensions of the aircraft could not be scaled down. In order to accommodate each aircraft in the Viking derivative aeroshell having a maximum internal diameter of 2.48 m, the wings, the body and the propeller must be folded. Each aircraft has a span of 34.1 m, a length of 8.8 m, and a 3.4 m propeller. This results in the need for 14 span-wise folds, 4 length-wise folds, and 2 propeller folds giving a total of 20 structural folds. The factors associated with the complexity of the fixed-wing aircraft for each mission are summarized in Table 5.4.

Mission 1 Mission 2 Mission 3 # Span-wise Folds 14 folds 14 folds 14 folds # Length-wise Folds 4 folds 4 folds 4 folds # Propeller Folds 2 folds 2 folds 2 folds # Total Structural Folds 20 folds 20 folds 20 folds

Table 5.4 Complexity Factors for Fixed-Wing Aircraft

100 5.2 Airship

Section 4.3.1 presents the methodology by which the characteristics of the airship for each mission were determined. In keeping with the RMA design, a fineness ratio of 2 was maintained in each case. Table 5.5 summarizes the characteristics of the airship for each mission. Due to the thin Martian atmosphere, a large volume of lifting gas is required to carry the payload resulting in a massive airship for each mission. The large frontal area results in a significantly larger drag force in comparison to the fixed-wing aircraft, which in turn results in a greater fuel and power requirement for the airship. The dimensions of each airship were scaled according to the volume of H2 required to lift the payload. The airships associated with Mission 1 and Mission 3 were scaled down, while the airship associated with Mission 2 was scaled up according to Equations (4.50) and (4.51).

Mission 1 Mission 2 Mission 3 Range (km) 500 1000 150 Length (m) 48 60.6 41.2 Diameter (m) 24 30.3 20.6 Volume (m3) 14743 29998 9435 3 H2 Density (kg/m ) 0.00058 0.00044 0.00058 j C02 Density (kg/m ) 0.01245 0.00944 0.01245 Velocity (m/s) 15 15 15 Reynolds Number 846063 809439 726204

MH2 (kg) 8.5 13.2 5.5 Moasbag (kg) 99 165.2 76.4 Mpayload (kg) 17.7 4.94 5.35

MFuei(kg) 18.4 38.1 4.1

Mprop System (kg) 0.45 0.47 0.34

MFuel System (kg) 2.3 3.5 1.5 Mstructure (kg) 28.6 44.2 18.3

MTotai(kg) 175 270 112 PRequired (kW) 0.46 0.54 0.33 Ppayload (kW) 0.074 0.02 0.021 PAvionics (kW) 0.2 0.2 0.2 PTotal(kW) 0.73 0.76 0.55

Table 5.5 Airship Characteristics

When determining the turn rate of the airship via a MATLAB program [81], it was found that the dimensions of the control surfaces (rudder) were not large enough to provide stability. The rudder was,

101 therefore, extended to the rear of the airship to achieve a stable configuration. Figure 5.1 illustrates the original dimensions of the control surface, which include a fixed fin and a movable rudder, and Figure 5.2 illustrates the modified configuration with both the fixed fin and the rudder enlarged.

20 15 . ~ ^~^^ 10 s^ ^^X 5 / \ 1 0 -5 -10 -15 V_^ _^' -20

30 x(m) x(m) (a) Mission 1 (b) Mission 2 (c) Mission 3

Figure 5.1 Originally Scaled Airships (Unstable Configuration)

(a) Mission 1 (b) Mission 2 (c) Mission 3

Figure 5.2 Airships with Enlarged Fin and Rudder (Stable Configuration)

In order to minimize the turn radius, the turn rate was calculated assuming a turning velocity of 15 m/s and a rudder deflection of 30°. Figure 5.3 illustrates the corresponding turn rate of the airship for each mission.

102 11 , , I .51 , , I .71 , , 1 0 5 10 15 0 5 10 15 0 5 10 15 Air speed (m/s) Air speed (m/s) Air speed (m/s)

(a) Mission 1 (b) Mission 2 (c) Mission 3

Figure 5.3 Turn Rates at Maximum Speed and Maximum Rudder Deflection (8 = 30° )

The factors associated with the manoeuvrability of the airship for each mission are summarized in Table 5.6.

Mission 1 Mission 2 Mission 3 Yaw Rate (deg/s) 5.1 4.1 6 Turn Radius (m) 168.5 210 143 Range of Speed (m/s) 15 15 15 Ability to Hover Yes Yes Yes

Table 5.6 Manoeuvrability Factors for Airship

For the airship, no structural folds are required to stow it in the aeroshell since deployment simply entails inflation of the gasbag. Successful demonstrations of gasbag inflation systems have been performed during the Venus Vega mission, further reducing the risk associated with using this platform.

103 5.3 Rotary-Wing Aircraft

As with the fixed-wing aircraft, an optimum flight condition for the rotary-wing aircraft was sought. As illustrated in Figure 4.5, there is an operating speed at which the power required is at a minimum; this speed was selected as the cruising speed for the rotary-wing aircraft. For each mission, Equation (4.61) was evaluated for a range of speeds in order to develop a power vs. speed profile similar to Figure 4.5 so as to determine the cruising velocity. As an example, Figure 5.4 illustrates the power vs. speed profile for Mission 1.

Figure 5.4 Power vs. Speed Profile (Mission 1)

In order to keep the gross takeoff mass within reasonable bounds, the available power was limited to the amount required to enable the rotorcraft to fly at 30% above or below the cruising velocity. As an example, for Mission 1, the cruising velocity is 75 m/s corresponding to a total power requirement of approximately 4.5 kW; the minimum allowable velocity is 53 m/s and the maximum allowable velocity

104 is 98 m/s corresponding to a total power requirement of approximately 5.0 kW. The characteristics associated with the rotary-wing aircraft for each mission are summarized in Table 5.7. The dimensions of each rotorcraft were scaled according to the gross takeoff mass. The rotorcrafts associated with Mission 1 and Mission 2 were scaled up, while the rotorcraft associated with Mission 3 was scaled down according to Equations (4.70) and (4.71).

Mission 1 Mission 2 Mission 3 Range (km) 500 1000 150 Chord (m) 0.83 0.89 0.44 Rotor Diameter (m) 5.3 5.7 2.8 Flat Plate Area (m2) 0.30 0.36 0.09 Solidity 0.25 0.25 0.25 [i (at optimum velocity) 0.52 0.62 0.44 Minimum Velocity (m/s) 53 62 44 Maximum Velocity (m/s) 98 116 82 Cruising Velocity (m/s) 75 89 63 Rotor Tip Speed (m/s) 143.75 143.75 143.75 Mpayload (kg) 17.7 4.94 5.35 MFuel (kg) 5.6 13.1 0.44

Mprop System (kg) 8.2 11.6 1.83 Mpuel System (kg) 8.0 18.7 0.63 Mstructure (kg) 45.4 56.1 9.6

MTotal(kg) 85 104.5 18 PRea(kW) 4.73 6.76 0.90 Ppayload (kW) 0.074 0.02 0.021 PAyionics (kW) 0.2 0.2 0.2 PTotal(kW) 5.0 6.98 1.12

Table 5.7 Rotary-Wing Aircraft Characteristics

As indicated by the power requirement for each mission, it is evident that the rotary-wing aircraft is less efficient than the other platforms. However, the results indicate that for short-range missions, the rotary-wing aircraft becomes a competitive platform with respect to the gross takeoff mass despite the high power requirement.

As with the fixed-wing aircraft, the turn radius of the rotary-wing aircraft is limited by the total available power. It can be seen from Equation (4.62) that minimizing the turn radius is achieved by minimizing the speed in the turn while increasing the bank angle. To maintain a reasonable bank angle,

105 the speed in the turn is selected as the cruising speed since minimum power is required. The factors associated with the manoeuvrability of the rotary-wing aircraft are summarized in Table 5.8.

Mission 1 Mission 2 Mission 3 Turn Velocity (m/s) 75 89 63 Bank Angle (deg) 22 23 23 Load Factor (g's) 1.08 1.09 1.09 Power Required in the Turn (kW) 5.0 6.97 1.11 Turn Radius (m) 3733 5003 2507 Range of Speed (m/s) 45 54 38 Ability to Hover None None None

Table 5.8 Manoeuvrability Factors for Rotary-Wing Aircraft

Based on the results above, it is evident that the rotary-wing aircraft demonstrates poor performance with respect to manoeuvrability with its large turn radius and inability to hover. The inefficiency of the rotorcraft with respect to power combined with the limitation in the available power result in the large turn radius.

In keeping with the co-axial rotor configuration of MARV, each rotary-wing aircraft is comprised of two rotors; therefore, the blades of both rotors need to be folded for stowage purposes. No chord-wise folds were required for each of the three missions. The parameters associated with the complexity of the rotary-wing aircraft for each mission are summarized in Table 5.9.

Mission 1 Mission 2 Mission 3 # Span-wise Folds 8 folds 8 folds 4 folds

Table 5.9 Complexity Factors for Rotary-Wing Aircraft

106 5.4 Summary of Results

Presented below is a summary of the evaluated performance metrics for each platform and for each mission.

Mission Concept 1 Fixed-Wing Airship Rotary-Wing Mass (kg) 58.9 175 85 Power (kW) 0.67 0.73 5.0 Turn Radius (m) 448 168.5 3733 Range of Speed (m/s) 10.1 15 45 Ability to Hover No Yes No Complexity (# folds) 20 0 8

Mission Concept 2 Fixed-Wing Airship Rotary-Wing Mass (kg) 47.8 270 104.5 Power (kW) 0.56 0.76 6.98 Turn Radius (m) 479 210 5003 Range of Speed (m/s) 10.5 15 54 Ability to Hover No Yes No Complexity (# folds) 20 0 8

Mission Concept 3 Fixed-Wing Airship Rotary-Wing Mass (kg) 43.9 112 18 Power (kW) 0.48 0.55 1.12 Turn Radius (m) 324 143 2507 Range of Speed (m/s) 8.8 15 38 Ability to Hover No Yes No Complexity (# folds) 20 0 4 5.5 Performance Rating of Aerial Platforms

Presented below is a summary of the weighted performance (calculated as Weight x Performance) of each platform with respect to the performance metrics as well as the overall performance rating indicating which platform is best suited for each mission. The platform with the highest value of Qpiat/orm xs deemed best suited for the mission.

Mission 1

Fixed-Wing Airship Rotary-Wing Mission Concept 1 Weight Weighted Weighted Weighted Performance Performance Performance Mass 3 30 10.2 20.7 Power 2 20 18.4 2.6 Manoeuvrability 1 3 6.7 5.3 Turn Radius 5 19 50 2.5 Range of Speed 5 11 16.5 50 Ability to Hover 0 0 0 0 Complexity 4 0 40 24 Qpiatform 5.3 7.5 5.3

Based on the results presented above, the fixed-wing aircraft demonstrated the best performance with respect to the gross takeoff mass and power requirement. However, the need for numerous structural folds significantly limited its overall performance. In comparison to the fixed-wing aircraft, the rotary- wing aircraft demonstrated very poor performance with respect to power and turn radius. Although it required far fewer structural folds, resulting in better performance with respect to complexity, the inefficiency of the rotary-wing aircraft, reflected by its high power requirement, limited its overall performance.

Although the airship demonstrated the worst performance with respect to the gross takeoff mass, its low power requirement, high degree of manoeuvrability and simple design make it the best platform for this mission. The airship does not require any structural folds since deployment of the airship simply requires inflation of the gasbag. Since gasbag inflation technology has been successfully demonstrated

108 previously on the Venus Vega Balloon missions, the risk associated with the deployment of this platform is significantly lower. In addition, the airship can be deployed either during descent or from the ground providing further flexibility. The airship is, therefore, best suited for Mission 1.

Mission 2

Fixed-Wing Airship Rotary-Wing Mission Concept 2 Weight Weighted Weighted Weighted Performance Performance Performance Mass 4 40 7.2 18.4 Power 1 10 7.4 0.8 Manoeuvrability 1 3.2 6.4 5.2 Turn Radius 5 22 50 2 Range of Speed 5 9.5 14 50 Ability to Hover 0 0 0 0 Complexity 4 0 40 24 Qpiatform 5.3 6.1 4.8

The results above indicate that for a long-range mission, the fixed-wing aircraft and the airship demonstrate reasonable performance. As with the previous mission, the fixed-wing aircraft demonstrated the best performance with respect to the gross takeoff mass and power requirement but its overall performance was significantly limited by its complexity. Again, although the rotary-wing aircraft requires fewer structural folds than the fixed-wing aircraft, its overall performance is limited by its inefficiency, which is reflected by its high power requirement. The airship demonstrated the worst performance with respect to the gross takeoff mass but its low power requirement, high degree of manoeuvrability and simple design once again make it the best platform for Mission 2.

109 Mission 3

Fixed-Wing Airship Rotary-Wing Mission Concept 3 Weight Weighted Weighted Weighted Performance Performance Performance Mass 1 4.1 1.6 10 Power 1 10 8.7 4.3 Manoeuvrability 4 10 32.8 12.8 Turn Radius 4 17.6 40 2.4 Range of Speed 3 6.9 11.7 30 Ability to Hover 3 0 30 0 Complexity 4 0 40 32 Qpiatform 2.4 8.3 5.9

The results above clearly indicate that for missions requiring a high degree of manoeuvrability, the airship is the best platform. In addition, its low power requirement, simple design and deployment flexibility make it the best platform for most mission scenarios. The complexity associated with the fixed-wing aircraft and the inefficiency associated with the rotary-wing aircraft make these platforms less suitable. However, for short-range missions, the rotary-wing aircraft outperforms the fixed-wing aircraft since it requires less mass and is less complex.

5.6 Factors Affecting the Overall Performance of the Aerial Platforms

Various subjective design choices were made that ultimately had an impact on the overall performance of the platforms. These design choices included:

Fixed-Wing Aircraft: • Selecting a cruising speed of 30% above the stall speed - this assumption was made primarily to enable a reasonable Reynolds number in cruising flight. The speed has an impact on the power required for flight. A higher speed would result in a greater required power and, therefore, a higher gross takeoff mass. However, based on the results of Section 5.1, it appears that the mass penalty is negligible.

110 • Setting the cruising speed as the maximum speed at which the aircraft can fly - this assumption resulted in the maximum available power being limited to that required for cruising flight. This in turn had an impact on the allowable turn speed and bank angle, thus affecting the aircraft's manoeuvrability. • Limiting the load factor to 1.75 - this assumption was made so as to remain consistent with the design of the Daedalus 88. However, limiting the load factor could have an impact on the maximum allowable bank angle, thus affecting the aircraft's manoeuvrability. • Selecting a turning speed of 15% above the stall speed - based on Equation (4.17), a lower turn speed results in a smaller turn radius, which is desirable. Thus, maintaining a turn speed of 15% above the stall speed limits the degree to which the turn radius can be minimized.

Airship: • Selecting a cruising speed of 15 m/s - this assumption was made so as to allow the airship to maintain directional control under moderately windy conditions on Mars. However, a higher cruising speed results in greater required power, and thus a higher gross takeoff mass. Depending on mission requirements, a trade-off can be made between mass savings and reduced directional controllability. • Limiting the maximum available power to that required for cruising flight - increasing the available power results in an increase in the gross takeoff mass; therefore, this assumption was made in an effort to limit the gross mass.

Rotary-Wing Aircraft: • Maintaining a rotor solidity of 0.25 and a rotor tip velocity of 143.75 m/s - this assumption was made so as to remain consistent with the design of MARV. The rotor solidity and rotor tip velocity are factors that affect the profile power. Understanding the impact of altering these values would help to optimize the performance of the rotorcraft. • Setting the cruising speed as the speed associated with minimum power requirement - this assumption was made so as to cruise at the optimum condition whereby fuel consumption is minimized. A higher speed would result in greater required power, thus increasing the fuel consumption as well as the gross takeoff mass.

Ill • Limiting the minimum and maximum speeds at which the platform can fly to 30% below and above the cruising speed respectively - this assumption resulted in the maximum available power being limited to that required to fly the rotorcraft at 30% above or below the cruising speed. This in turn had an impact on the rotorcraft's manoeuvrability since the maximum available power was not sufficient for hovering and performing sharp turns. • Selecting a turning speed equal to the speed associated with minimum power requirement - selecting a different turn speed would require more power to perform a turn, resulting in a smaller allowable bank angle, thus further limiting the rotorcraft's manoeuvrability.

It is important to note that variations in these design choices could result in quite different results. For instance, increasing the cruising speed of the fixed-wing aircraft would allow its dimensions to be scaled down thus improving the complexity metric, which is currently its major limiting factor. An analysis should be performed to determine how sensitive the overall performance of each platform is to variations in the parameters listed above. This in turn would serve to optimize the design of the platforms for each particular mission. For example, if it is known that for a particular mission, the weight on power is low while that on manoeuvrability is high, then a better design choice for the fixed- wing and rotary-wing aircraft would be to allow a larger installed power than was done in the present work.

It is also important to note that the weights assigned to the performance metrics, indicating their relative importance in achieving the mission requirements, is also subjective. Altering the requirements for a given mission may also result in a different outcome with respect to the most suitable platform, since the weights are mission dependent. For instance, if the requirements pertaining to Mission 3 were altered such that a high degree of manoeuvrability was not required and greater emphasis was put on the gross takeoff mass and complexity, then the rotary-wing aircraft would have been the best platform as indicated in the results below.

112 Fixed-Wing Airship Rotary-Wing Mission Concept 3 Weight Weighted Weighted Weighted Performance Performance Performance Mass 4 16.4 6.4 40 Power 1 10 8.7 4.3 Manoeuvrability 1 2.5 8.2 3.2 Turn Radius 5 22 50 3 Range of Speed 5 11.5 19.5 50 Ability to Hover 0 0 0 0 Complexity 4 0 40 32 Qpiatform 2.9 6.3 8

113 Chapter 6

CONCLUSIONS AND FUTURE WORK

6.1 Conclusions

As part of this thesis, a framework was developed to compare the performance of fixed- wing aircraft, airships, and rotary-wing aircraft in the Martian environment. The framework established the context within which the performance of the platforms would be evaluated, a set of performance metrics providing a common basis for comparing their performance, and a set of assumptions with respect to the design of each aerial platform. The framework was then used to determine which of these platforms would be best suited to perform a series of scientific investigations on Mars.

The performance of each platform was evaluated for the cruising portion of the flight only. Existing platform designs were used for the purpose of this comparative study. However, modifications were made to these existing designs to account for specific mission requirements. In addition, various subjective design choices, summarized in Section 5.6, were made that ultimately had an impact on the overall performance of the platforms. For instance, limiting the maximum available power of the fixed-wing aircraft resulted in limitations in the turn speed and bank angle, thus affecting its manoeuvrability. The decision to limit the minimum and maximum speeds at which the rotary-wing aircraft could fly resulted in its inability to hover due to insufficient available power. The selected cruising speed for the airship had a direct impact on its size and ultimately on its gross takeoff mass.

114 An optimum flight condition was sought for each platform so as to optimize the overall performance. It was found that for the fixed-wing aircraft, operating at the ideal minimum drag condition was not feasible on Mars since the minimum drag speed was below the stall speed for the selected airfoil at this condition. As such, a decision to set the cruising speed of the fixed-wing aircraft to 30% above the stall speed was made based on FAA recommendations for gliding flight.

Although the normal wind speed on Mars ranges between 5-10 m/s, maximum wind speed could reach up to 40 m/s. An optimum cruising speed of 15 m/s was selected for the airship to allow it to maintain directional control under moderately windy conditions while maintaining a reasonable power requirement. Since the speed of the airship has a direct impact on its gross takeoff mass, depending on the mass budget of a particular mission, a trade-off can be made between mass savings and reduced directional controllability.

The cruising speed of the rotary-wing aircraft was set as the speed at which the power requirement is at a minimum in order to minimize the amount of fuel required, thus minimizing its gross takeoff mass. However, it was found that due to the inefficient performance of the rotorcraft with respect to power, even the minimum power requirement was substantially higher than the power requirement of the other two platforms. As a result, the total available power was limited such that the range of speed of the rotorcraft was restricted to 30% above or below the cruising speed. This design decision to limit the maximum available power of the rotorcraft had a significant impact on its manoeuvrability since the available power was not sufficient for hovering and for performing sharp turns.

For all three Mars mission scenarios, the fixed-wing aircraft demonstrated the best performance with respect to the gross takeoff mass and power required. Although it demonstrated moderate performance with respect to the turn radius (which was limited by the available power), its narrow range of speed and inability to hover resulted in poor overall performance with respect to manoeuvrability.

115 Assuming a fixed-wing aircraft having the same physical dimensions as the Daedalus 88, the fixed cruising velocity resulted in a very low operating Reynolds number in the range of 4 x 104 for each mission. A consequence of this low Reynolds number is that the dimensions of the aircraft could not be scaled down since it would further decrease the Reynolds number and no other airfoils were found that could operate effectively at Reynolds numbers below 4 x 104.

The result is that a significant number of structural folds were required to accommodate the fixed-wing aircraft in the aeroshell, which in turn resulted in extremely poor performance with respect to complexity. Since the complexity metric was weighted high in each mission, as it has a big impact on mission success, the poor performance of the fixed-wing aircraft with respect to complexity ultimately limited the aircraft's overall performance, making it an unsuitable platform for all mission scenarios.

For all three Mars mission scenarios, the airship demonstrated the worst performance with respect to the gross takeoff mass since the thin Martian atmosphere resulted in a large volume of lifting gas being required to carry the payload. On the other hand, its power requirement was low, and its short turn radius, moderate range of speed and ability to hover resulted in the airship having the best performance with respect to manoeuvrability. In addition, its simple design, requiring no structural folds, made the airship the least complex platform. As a result, the airship demonstrated the best overall performance for each of the three missions.

For all three Mars mission scenarios, the rotary-wing aircraft demonstrated moderate performance with respect to the gross takeoff mass but the worst performance with respect to the power required. Although the rotary-wing aircraft demonstrated the best performance with respect to the range of speed, its performance with respect to the turn radius was the worst due to the limitations in the available power. This, combined with the inability to hover resulted in the rotary-wing aircraft having poor performance with

116 respect to manoeuvrability. A small number of structural folds resulted in moderate performance with respect to complexity.

The inefficiency of the rotary-wing aircraft, reflected by its high power requirement, limited its overall performance. However, it was found that for short-range missions it demonstrated better overall performance than the fixed-wing aircraft. In fact, for short- range missions that do not require a high degree of manoeuvrability and that have severe mass constraints, the rotary-wing aircraft would be the best platform.

Although the airship appears to be the best platform for each of the Mars mission scenarios, the underlining conclusion drawn from this study is that the overall performance of the aerial platforms considered was greatly influenced by the subjective design choices that were made as well as by the mission requirements. Alternative design choices may result in quite different results. For instance, increasing the cruising velocity of the fixed-wing aircraft may allow its dimensions to be scaled down thereby reducing the complexity and making it a more competitive platform for long-range missions. With respect to mission requirements, it was found that for short-range missions that do not require a high degree of manoeuvrability but where the gross takeoff mass and complexity are more important factors, the rotary-wing aircraft becomes the best option.

117 6.2 Future Work

Various subjective design choices were made in this study in relation to the aerial platforms. These design choices ultimately had an impact on the overall performance of the platforms. A future endeavour in this area of research would be to perform a sensitivity study to determine how sensitive the overall performance of each platform is to variations in these design choices. Examples of typical questions that the sensitivity study would aim to address include, but are not limited to, the following: Until what point can the cruising speed of the fixed-wing aircraft be increased before the associated increase in the power begins to limit its overall performance? At this point, is the improvement with respect to complexity sufficient to make the fixed-wing aircraft a more competitive platform? For the airship, how sensitive is the mass metric to changes in the cruising velocity? For the rotary-wing aircraft, how much of an increase in the total power is required to achieve a significant improvement in the turn radius? Which performance metrics are most sensitive to changes in mission requirements?

Another recommendation for future work would be to investigate the performance of hybrid airships in the Martian environment. Hybrid airships combine the advantages of both the fixed-wing aircraft as well as the airship. The inflatable wing structure would eliminate the need for structural folds, thus minimizing the complexity that is normally associated with the fixed-wing aircraft. In addition, a smaller volume of lifting gas would be required to lift the payload since the lift is generated via both the buoyancy of the lifting gas as well as aerodynamically, thus providing a mass savings over conventional airships.

In addition, continued efforts to model and replicate flapping-wing motion as seen in birds and insects would lead to efficient, highly manoeuvrable flapping-wing vehicles that are able to operate effectively in the Martian environment at low flight Reynolds numbers by taking advantage of unsteady aerodynamic phenomena.

118 REFERENCES

[1] http://mars.ipl.nasa.gov/missions/

[2] http://robotics.ipl.nasa.gov/tasks/aerobot/studies/vega.html

[3] http://mepag.jpl.nasa. gov/goals/MEPAGgoals-approved071604.pdf

[4] J. A. Cutts, J. Bauer, D. L. Blaney, L. G. Lemke, S. Smith Jr., D. S. Stetson, and V. V. Kerzhanovich, "Role of Mars Aerial Platforms in Future Exploration of Mars, Prepared in Support of the Mars Architecture Study", August 31, 1998 (http://telerobotics.jpl.nasa.gov/aerobot/reports/reports.html)

[5] V. C. Clarke, Jr. Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California, A. Kerem, Developmental Sciences, Inc., City of Industry, California, R. Lewis, Lear Seigler, Inc., Astronics Division, Santa Monica, California, "A Mars Airplane.. .Oh Really?"

[6] W. M. Clapp, "Dirigible Airships for Martian Surface Exploration", The Case For Mars II: Proceedings of the Second Case for Mars Conference held July 10-14, 1984, University of Colorado, Boulder, Colorado

[7] J. Langford, Atlantic Aircraft Co., Alexandria, VA, "The Daedalus Project: A Summary of Lessons Learned", AIAA/AHS/ASEE Aircraft Design, Systems and Operations Conference, Seattle, WA, July 31 - August 2 1989

[8] K. R. Sivier, M. F. Lembeck, University of Illinois, "The Marsplane Revisited", AIAA/AHS/ASEE Aircraft Design, Systems, and Operations Meeting, September 7-9, 1988, Atlanta, Georgia

119 [9] B. R. Morrissette, J. D. DeLaurier, "The Zephyr: Manned Martian Aircraft", Master's Thesis, Graduate Department of Aerospace Science and Engineering, University of Toronto

[10] A. Colozza, Sverdrup Technology Inc., Brook Park, OH, "Preliminary Design of a Long-Endurance Mars Aircraft", AIAA/SAE/ASME/ASEE 26th Joint Propulsion Conference, July 16-18, 1990, Orlando, FL

[11] S. C. Smith, A. S. Hahn, W. R. Johnson, D. J. Kinney, J. A. Pollitt, J. R. Reuther, NASA Ames Research Center, Moffett Field, CA, "The Design of the Canyon Flyer, An Airplane for Mars Exploration", 38th Aerospace Sciences Meeting and Exhibit, January 10-13, 2000, Reno, NV

[12] C. A. Kuhl, H. S. Wright, C. A. Hunter, NASA Langley Research Center, Hampton, VA, C. S. Guernsey, Jet Propulsion Laboratory, Pasadena, CA, A. J. Colozza, Analex Corporation, NASA Glenn Research Center, Cleveland, OH "Liquid Rocket Propulsion for Atmospheric Flight in the Proposed ARES Mars Scout Mission"

[13] M. S. Smith, S. R. Schallenkamp, Raven Industries, Inc., Sulphur Springs, TX, C. J. Ekstein, K. Blizard, Foster-Miller, Inc., Waltham, Massachusetts, "Development of Venusian Aerobots"

[14] R. J. Levesque, G. E. Williams, F. J. Redd, Utah State University, "Balloon-Borne Characterization of the Martian Surface and Lower Atmosphere", The Case for Mars III: Strategies for Exploration: Proceedings of the Third Case for Mars Conference held July 18-22, 1987, University of Colorado, Boulder, Colorado

[15] R. Greeley, P. R. Christensen, Arizona State University, B. C. Clark, S. R. Price, R. M. Zubrin, Lockheed Marietta Astronautics, R. M. Haberle, NASA Ames Research Center, J. Cantrell, Space Dynamics Laboratory, J. A. Cutts, R. E.

120 Oberto, Jet Propulsion Laboratory, M. C. Malin, Malin Space Science Systems, "The Mars Aerial Platform (MAP) Concept", 34th Aerospace Sciences Meeting and Exhibit: January 15-18, 1996, Reno, NV

[16] J. A. Jones, J. J. Wu, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California, "Solar Montgolfiere Balloons for Mars", AIAA International Balloon Technology Conference, June 28 - July 1, 1999, Norfolk, VA

[17] K. T. Nock, J. A. Jones, G. Rodriguez, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California, "Planetary Aerobots: A Program for Robotic Balloon Exploration", 34th Aerospace Sciences Meeting and Exhibit: January 15-18, 1996, Reno, NV

[18] A. R. Girerd, Anser, Alexandria, VA, "The Case for a Robotic Martian Airship", 14' AIAA Aerodynamic Decelerator Systems Technology Conference, San Francisco, CA, June 3-5, 1997

[19] D. E. Calkins, Universidade Federal do Rio de Janerio, Rio de Janeiro, Brazil, "Feasibility Study of a Hybrid Airship Operating in Ground Effect", Journal of Aircraft, Vol. 14, NO 8, p. 809-815, 1977

[20] P. A. Mackrodt, DFVLR-AVA, Gottingen, Federal Republic of Germany, "Further Studies in the Concept of Delta-Winged Hybrid Airships", Journal of Aircraft, Vol. 17, NO 10, p. 734-740, 1980

[21] J. F. Gundlach IV, Blacksburg, VA, "Unmanned Solar-Powered Hybrid Airships for Mars Exploration", 37th AIAA Aerospace Sciences Meeting and Exhibit, January 11-14, 1999, Reno, NV [22] G. Savu, C. Oprisiu, O. Trifu, Institute of Fluid Mechanics and Flight Dynamics, Bucharest, Romania, "An Autonomous Flying Robot for Mars Exploration", 44th

121 Congress of the International Astronautical Federation, October 16-22, 1993, Graz, Austria

[23] G. Gavu and 0. Trifu, Institute of Fluid Mechanics and Flight Dynamics, Bucharest, Romania, "Photovoltaic Rotorcraft for Mars Missions", 31st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, July 10-12, 1995, San Diego, CA

[24] K. J. Corfeld, L. N. Long, Pennsylvania State University, R. C. Strawn, Ames Research Center, "Computational Analysis of a Prototype Martian Rotorcraft Experiment", 31st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit: July 10-12, 1995, San Diego, CA

[25] L. A. Young, Ames Research Center, Moffett Field, CA , "Vertical Lift - Not Just for Terrestrial Flight"

[26] L. A. Young, R. T. Chen, E. W. Aiken, G. A. Briggs, NASA Ames Research Center, Moffett Field, CA, "Design Opportunities and Challenges in the Development of Vertical Lift Planetary Aerial Vehicles"

[27] I. Chopra, A. Datta, J. Bao, O. Gamard, D. Griffiths, L. Liu, G. Pugliese, B. Roget, J. Sitamaran, Alfred Gessow Rotorcraft Center, University of Maryland, College Park, Maryland, "The Martian Autonomous Rotary-Wing Vehicle (MARV)"

[28] T. W. Streett, J. F. Marchman, Virginia Polytechnic Institute, "Rotor Design for an Unmanned Helicopter for use on Mars"

[29] J. D. DeLaurier, University of Toronto Institute of Aerospace Studies, "The Development and Testing of a Full-Scale Piloted Ornithopter", Canadian Astronautics and Space Journal, Vol. 45, No. 2, June 1999

122 [30] S. P. Sane, "The Aerodynamics of Insect Flight", The Journal of Experimental Biology Vol. 206, p. 4191-4208, 2003

[31] R. C. Michelson, M. A. Naqvi, Georgia Institute of Technology, Atlanta, Gerogia, "Extraterrestrial Flight (Entomopter-Based Mars Surveyor)"

[32] A. Ania, D. Poirel, Royal Military College of Canada, M. Potvin, Canadian Space Agency, "Martian Flight Prototypes"

[33] G. Desmarais, Concordia University, Montreal, Canada, 2003, "Martian Flight Study: prepared for the Canadian Space Agency"

[34] M. D. Guynn, M. A. Croom, NASA Langley Research Center, S. C. Smith, NASA Ames Research Center, R. W. Parks, Aurora Flight Sciences, G. A. Gelhausen, AVID LLC, "Evolution of a Mars Airplane Concept for the Ares Mars Scout Mission", 2" AIAA "Unmanned, Unlimited" Systems, Technologies and Operations - Aerospace Conference, 15 - 18 September, 2003, San Diego, California

[35] http://www.nasa.gov/mission_pages/mars/news/mgs-20061206.html

[36] http://mars.ipl.nasa.gov/mro/mission/sc instru sharad.html

[37] http://mars.ipl.nasa.gov/mro/mission/sc instru crism.html

[38] http://www.lpi.usra.edu/meetings/geomars2001 /virtual.pdf

[39] http://www.sharad.org/index.php?page=instrument description.php&UserSID=

[40] http://mars.jpl.nasa.gov/mro/mission/sc instru sharad.html

123 [41] R. E. Grimm, Blackhawk Geoservices, "Comparison of Ground-Penetrating Radar and Low Frequency Electromagnetic Sounding for Detection and Characterization of Groundwater on Mars", Sixth International Conference on Mars, 2003, http://www.lpi.usra.edU/meetings/sixthmars2003/pdf/3176.pdf

[42] http://www.terrapub.co.ip/iournals/EPS/pdf/5003/50030183.pdf

[43] http://www.newscientist.com/article.ns?id=dn8397

[44] http://www.lpi.usra.edu/education/K12/gangis/marsdatum.html

[45] http://nssdc.gsfc.nasa. gov/database/MasterCatalog?sc=1996-068A&ex=2

[46] http://marstech.ipl.nasa.gov/content/detail.cfm?Sect=IG=&Cat=base&subCat=MI

[47] http://laserweb.ipl.nasa.gov/planetaryinstruments/tls.html

[48] http://www.nasa.gov/centers/dryden/research/AirSci/ER-2/pi inst.html#ATLAS

[49] http://marstech.ipl.nasa.gov/content/detail.cfm?Sect=IG&Cat=base&subCat=MID

[50] http://www.newscientist.com/article.ns?id=dnl0312&feedld=solar-svstem_rss20

[51] http://en.wikipedia.org/wiki/Mossbauer spectroscopy

[52] http://en.wikipedia.org/wiki/Raman spectroscopy

[53] C. -P. Sherman Hsu, Ph.D., Separation Sciences Research and Product Development, Mallinckrodt, Inc., Mallinckrodt Baker Division, Infrared Spectroscopy, Chapter 15

124 [54] http://minites.asu.edu/Mini-TES Qverview.pdf

[55] http://www.iop.Org/EJ/abstract/0957-0233/9/8/013

[56] http://www-ssc.igpp.ucla.edu/personnel/russell/papers/Venus fluxgate/

[57] http://mgs-mager.gsfc.nasa.gov/instrument.html

[58] http://www.msss.com/mars/mars9x/penetratorpayload.html

[59] http://astrogeology.us gs. gov/Proi ects/VallesMarineri s/

[60] I. S. Smith, J. A. Cutts, "Floating in Space", Scientific American, November 1999, pp.103

[61] D. Raymer, Aircraft Design: A Conceptual Approach, Third Edition, AIAA Education Series

[62] http://en.wikipedia.org/wiki/Gossamer Albatross

[63] R. B. Sullivan, S.H. Zerweckh, MIT, Cambridge, MA, October 1988, Grant NAG-1-836, "Flight Test Results for the Daedalus and Light Eagle Human Powered Aircraft", http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19890001519 1989001519.pd f

[64] T. Shalon, J.S. Langford, and R.W. Parks, "Design and Sizing of the MIT Daedalus Prototype", AIAA 87-2907, August 1987

125 [65] M. Frumkin, NASA Advanced Supercomputing Division, Moffett Field, CA 94035-1000, "Optimizing Mars Airplane Trajectory with the Application Navigation System", NAS Technical Report NAS-04-17

[66] http://www.airforce-technology.com/contractors/flight/athena/

[67] R.D. Reed, "High-Flying Mini-Sniffer RPV: Mars Bound", Astronautics and Aeronautics, vol. 16, no. 6, 1978, pp. 26-39

[68] A. J. Colozza, Analex Corporation, Brook Park, Ohio, "Comparison of Mars Aircraft Propulsion Systems", NASA/CR—2003-212350

[69] M. S. Selig, J. F. Donovan, D. B. Frazer, Airfoils at Low Speeds, Soartech 8, H. A. Stokely, Publisher, 1989

[70] http://facultv.erau.edu/ericksol/projects/mars/group2/mars2a.html

[71] D. Anubhav, et al, "Design of a Martian Autonomous Rotary-Wing Vehicle", Journal of Aircraft, vol. 40, No.3, 2003 pp. 461-472

[72] R. Sullivan et al, "Results of the Imager for Mars Pathfinder Windsock Experiment", Journal of Geophysical Research, vol. 105, No. E10, 2002, pp. 24,547-24,562

[73] L. A. Young, G.A. Briggs, "Design Opportunities and Challenges in the Development of Vertical Lift Planetary Aerial Vehicles", American Helicopter Society International Vertical Lift Aircraft Design Specialist's Meeting, San Francisco, CA, January 19-20, 2000

[74] 1974 Mars Atmosphere Model, JPL

126 [75] http://www.engr.uky.edU/~idiacob/me380-aircraft/programs/glidingdescent.m

[76] J. B. Russell, Performance and Stability of Aircraft, John Wiley & Sons, 1996

[77] http://www.lissys.demon.co.uk/pug/c04.html

[78] http://windows.uwaterloo.ca/Hardware/PC Power Consumption.asp

[79] T. Liu, Western Michigan University, Kalamazoo, Michigan 49008, "Comparative Scaling of Flapping- and Fixed-Wing Flyers"

[80] G. A. Khoury, Imperial College, London, and The Airship Association, J. D. Gillett, Formerly of Brunei University and The Airship Association, Airship Technology, Cambridge University Press

[81] Y. Li, Dynamics Modeling and Simulation of Flexible Airships, PhD Thesis, McGill University, 2007

[82] R. Cownden, M. Nahon, M. A. Rosen, "Modeling and Analysis of a Solid Polymer Fuel Cell System for Transportation Applications", International Journal of Hydrogen Energy 26 (2001) 615-623

[83] Ulf Bossel, European Fuel Cell Forum, Morgenacherstrasse 2F, CH-5452, Oberrohrdorf/Switzerland, "Well-To-Wheel Studies, Heating Values and the Energy Conservation Principle"

[84] Barnes W. McCormick, Aerodynamics Aeronautics and Flight Mechanics, 2nd Edition, Pennsylvania State University

[85] http://www.stlouishelo.org/School%20of%20Aviation%20Safety%20- %20Power%20Available%20vs.%20Power%20Required.pdf

127 [86] Transport Canada Helicopter Flight Training Manual (TP 9982), http://www.transportcanada.com/CivilAviation/general/Flttrain/Planes/Pubs/TP99 82/Exercies

[87] R. D. Lorenz, "Flight Power Scaling of Airplanes, Airships, and Helicopters: Application to Planetary Exploration", University of Arizona, Tucson, Arizona 85721-0092

[88] A. Colozza, R. C. Michelson, M. A. Naqvi, et al, "Planetary Exploration Using Biomimetics - An Entomopter for Flight on Mars", Phase II Final Report, NASA Institute for Advanced Concepts Project NAS5-98051, October 2002

[89] S. P. Sane, Department of Biology, University of Washington, "The Aerodynamics of Insect Flight", The Journal of Experimental Biology, Vol. 206, pp. 4191-4208

[90] C. P. Ellington, Department of Zoology, University of Cambridge, "The Novel Aerodynamics of Insect Flight: Applications to Micro-Air Vehicles", The Journal of Experimental Biology, Vol. 202, pp. 3439-3448

[91] W. Shyy, M. Berg, B. Ljungqvist, Department of Aerospace Engineering, Mechanics & Engineering Science, University of Florida, "Flapping and Flexible Wings for Biological and Micro Air Vehicles", Progress in Aerospace Sciences, Vol. 35, 1999, pp. 455-505

[92] J. M. Birch, M. H. Dickinson, "Spanwise Flow and the Attachment of the Leading-Edge Vortex on Insect Wings", Nature, Vol. 412, 2001, pp. 729-733

[93] C. Van Den Berg, C. P. Ellington, Phil. Trans. R. Soc. Lond. B, "The Three- Dimensional Leading-Edge Vortex of a 'Hovering' Model Hawkmoth", Vol. 352, 1977,pp. 329-340

128 [94] J. M. Birch, W. B. Dickson, M. H. Dickinson, "Force Production and Flow Structure of the Leading Edge Vortex on Flapping Wings at High and Low Reynolds Numbers", Journal of Experimental Biology, Vol. 207, 2004, pp. 1063- 1072

[95] M. H. Dickinson, F. Lehmann, S. P. Sane, "Wing Rotation and the Aerodynamic Basis of Insect Flight", Science, Vol. 284, 1999, pp. 1954-1960

[96] R. B. Srygley, A. L. R. Thomas, "Unconventional Lift-Generating Mechanisms in Free-Flying Butterflies", Nature, Vol. 420, pp. 6600-6604

[97] M. S. Vest, J. Katz, Department of Aerospace Engineering and Engineering Mechanics, "Aerodynamic Study of a Flapping-Wing Micro UAV", 37th AIAA Aerospace Sciences Meeting and Exhibit: January 11-14, 1999, Reno, NV

[98] V. Malolan, M. Dineshkumar, V. Baskar, Madras Institute of Technology, Anna University, Chennai, India, "Design and Development of Flapping Wing Micro Air Vehicle", 42nd AIAA Aerospace Sciences Meeting and Exhibit: January 5-8, 2004, Reno, NV

[99] A. Muniappan, V. Baskar, V. Duriyanandhan, Department of Aerospace Engineering, Madras Institute of Technology, "Lift and Thrust Characteristics of Flapping Wing Micro Air Vehicle (MAV)", 43r AIAA Aerospace Sciences Meeting and Exhibit: January 10-13, 2005, Reno, NV

[100] J. D. DeLaurier, J. M. Harris, "A Study of Mechanical Flapping-Wing Flight", Aeronautical Journal, Vol. 97, 1993, pp. 277

[101] http://www.ornithopter.ca/Publications/OrnithopterReportfor8July2006.pdf

129 [102] J. D. DeLaurier, Institute for Aerospace Studies, University of Toronto, "The Development of an Efficient Ornithopter Wing", AeronauticalJournal, May 1993, pp. 153-162

[103] http://marsairplane.larc.nasa.gov/platform.html

130 APPENDIX A

Previous Mars Missions

Presented below is a brief account of previous flyby, orbiter, and lander and rover missions to Mars. The information contained herewith has been directly extracted from NASA's primary website relating to Mars exploration [1].

Flyby Missions

Between 1962 and 1973, NASA's Jet Propulsion Laboratory designed and built 10 spacecraft named Mariner to explore the inner solar system. Mariner 3 and 4 were identical spacecraft designed to carry out the first flybys of Mars. Mariner 3 was launched on November 5, 1964, but the shroud encasing the spacecraft atop its rocket failed to open properly, and Mariner 3 did not get to Mars. Three weeks later, on November 28, 1964, Mariner 4 was launched successfully on an eight-month voyage to the red planet. The spacecraft flew past Mars on July 14, 1965, collecting the first close- up photographs of another planet.

Mariner 6 and 7 were the second pair of Mars missions in NASA's Mariner series of solar system exploration. In 1969, Mariner 6 and Mariner 7 completed the first dual mission to Mars, flying by over the equator and south Polar regions and analyzing the Martian atmosphere and surface with remote sensors, as well as recording and relaying hundreds of pictures. By chance, both flew over cratered regions and missed both the giant northern volcanoes and the equatorial Grand Canyon that were discovered later.

131 Figure A.1 Mariner 4 Spacecraft [1]

Orbiter Missions

Mariner 8 and 9 were the third and final pair of Mars missions in NASA's Mariner series of the 1960s and early 1970s. Both were designed to be the first Mars orbiters, marking a transition in our exploration of the red planet from flying by the planet to spending time in orbit around it. Unfortunately, Mariner 8 failed during launch on May 8, 1971.

Mariner 9, launched successfully on May 30,1971, observed that a great dust storm was obscuring the whole globe of the planet. The storm persisted for a month, but after the dust cleared, Mariner 9 proceeded to reveal a very different planet than expected, one that boasted gigantic volcanoes and a grand canyon stretching 4,800 kilometers (3,000 miles) across its surface. More surprisingly, the relics of ancient riverbeds were carved in the landscape of this seemingly dry and dusty planet. Mariner 9 also provided the first closeup pictures of the two small, irregular Martian moons: Phobos and Deimos.

132 Mars Observer, launched on September 25, 1992, consisted of a payload of science instruments that was designed to study the geology, geophysics and climate of Mars. The mission ended with disappointment on August 22, 1993, when contact was lost with the spacecraft shortly before it was to enter orbit around Mars. The science instruments from Mars Observer were re-flown on two other orbiters, Mars Global Surveyor and 2001 Mars Odyssey.

Mars Global Surveyor became the first successful mission to the red planet in two decades when it launched November 7,1996, and entered orbit on September 12, 1997. The mission studied the entire Martian surface, atmosphere, and interior. Among key science findings so far, Global Surveyor has taken pictures of gullies and debris flow features that suggest there may be current sources of liquid water, similar to an aquifer, at or near the surface of the planet. Magnetometer readings show that the planet's magnetic field is not globally generated in the planet's core, but is localized in particular areas of the crust. New temperature data and close-up images of the Martian moon Phobos show its surface is composed of powdery material at least 1 meter (3 feet) thick, caused by millions of years of meteoroid impacts. Data from the spacecraft's laser altimeter have given scientists their first 3-D views of Mars' north polar ice cap.

Figure A.2 Mars Climate Orbiter [1]

Mars Climate Orbiter, launched on December 11,1998, was designed to function as an interplanetary weather satellite and a communications relay for Mars Polar Lander. The orbiter carried two science instruments: a copy of an atmospheric sounder on the Mars

133 Observer spacecraft lost in 1993, and a new, lightweight color imager combining wide- and medium-angle cameras. Mars Climate Orbiter was lost on arrival September 23, 1999. Engineers concluded that the spacecraft entered the planet's atmosphere too low and probably burned up.

2001 Mars Odyssey, launched on April 7, 2001, was a spacecraft designed to map chemical elements and minerals on the surface of Mars, look for water in the shallow subsurface, and analyze the radiation environment to determine its potential effects on human health. The surface of Mars has long been thought to consist of a mixture of rock, soil and icy material. However, the exact composition of these materials is largely unknown. Odyssey collected images that were used to identify the minerals present in the soils and rocks on the surface and to study small-scale geologic processes and landing site characteristics.

Mars Express, a European mission launched on June 2, 2003, was sent to explore the atmosphere and surface of Mars. The mission's main objective was to search for sub surface water from orbit and deliver a lander to the Martian surface. The orbiter successfully entered the Martian orbit on December 25, 2003. Unfortunately, the Beagle 2 lander was declared lost after it failed to make contact with the orbiting spacecraft and Earth-based radio telescopes.

Seven scientific instruments onboard the Mars Express orbiting spacecraft were used to study the Martian atmosphere, as well as the planet's structure and geology. MARSIS, an instrument designed to map the sub-surface structure to a depth of a few kilometers, has detected ancient impact craters buried beneath the smooth, younger surface of the Northern Martian hemisphere. This buried terrain signifies that the underlying crust is extremely old, perhaps as ancient as the heavily cratered highland crust in the southern hemisphere. Mars Express will also serve as a communications relay, as part of NASA's Deep Space Network, to support other lander missions.

134 Figure A.3 Mars Reconnaissance Orbiter (MRO) [1]

NASA's Mars Reconnaissance Orbiter (MRO), launched on August 12, 2005, is on a search for evidence that water persisted on the surface of Mars for a long period of time. Mars Reconnaissance Orbiter successfully arrived at Mars on March 10, 2006. The mission has shown unprecedented detail in orbital images of Mars. With its suite of six instruments, MRO will zoom in for extreme close-up photography of the Martian surface, analyze minerals, look for subsurface water, trace how much dust and water are distributed in the atmosphere, monitor daily global weather, and characterize future potential landing sites.

Landers and Rovers

NASA's Viking Project found a place in history when it became the first mission to land a spacecraft safely on the surface of another planet. The Viking 1 lander, launched on August 20, 1975, touched down on the western slope of Chryse Planitia (the Plains of Gold), while the Viking 2 lander, launched on September 9, 1975, settled down at Utopia Planitia.

Besides taking photographs and collecting other science data on the Martian surface, the two landers conducted three biology experiments designed to look for possible signs of

135 life. These experiments discovered unexpected and enigmatic chemical activity in the Martian soil, but provided no clear evidence for the presence of living microorganisms in soil near the landing sites.

Mars Pathfinder, launched on December 4, 1996, was originally designed as a technology demonstration of a way to deliver an instrumented lander and a free-ranging robotic rover to the surface of the red planet. An ancient flood plain in Mars' northern hemisphere known as Ares Vallis, among the rockiest parts of Mars, was chosen as the landing site. Mars Pathfinder returned more than 16,500 images from the lander and 550 images from the rover, as well as more than 15 chemical analyses of rocks and soil and extensive data on winds and other weather factors. Findings from the investigations carried out by scientific instruments on both the lander and the rover suggest that Mars was at one time in its past warm and wet, with water existing in its liquid state and a thicker atmosphere.

Figure A.4 Mars Pathfinder [1]

Mars Polar Lander, launched on January 3, 1999, was an ambitious mission to set a spacecraft down on the frigid terrain near the edge of Mars' south polar cap and dig for water ice with a robotic arm. Piggybacking on the lander were two small probes called

136 Deep Space 2 designed to impact the Martian surface to test new technologies. Mars Polar Lander and Deep Space 2 were lost on arrival on December 3, 1999.

The Mars Exploration Rovers, Spirit and Opportunity, launched toward Mars on June 10 and July 7, 2003, respectively in search of answers about the history of water on Mars. Spirit and Opportunity landed on Mars on January 2, 2004, and January 24, 2004, respectively. Primary among the mission's scientific goals was to search for and characterize a wide range of rocks and soils that hold clues to past water activity on Mars. The rovers were targeted to sites on opposite sides of Mars that appear to have been affected by liquid water in the past. The landing sites selected were the Gusev Crater, a possible former lake in a giant impact crater, and Meridiani Planum, where mineral deposits (hematite) suggest Mars had a wet past. The presence of hematite was confirmed, providing the first mineral evidence that Mars' history may have included water. The rovers are still in operation over three years after landing on Mars.

On June 2, 2003, a lander called Beagle 2 was launched on Mars Express to perform exobiology and geochemistry research. The separation of Beagle 2 from Mars Express occurred on 19 December, 2003. The satellite continued its mission with its successful insertion into a Mars orbit on 25 December, 2003, the day on which Beagle 2 was due to land. The first radio contact with Beagle 2 was expected shortly after the scheduled landing time but no signal was received. Many radio contacts were attempted over the following days and weeks, but without result.

137 APPENDIX B Unconventional Aerial Platforms

In addition to the conventional aerial platforms considered in this study, the use of unconventional platforms such as flapping-wing aircraft have also been considered for Mars exploration. Significant interest in flapping-wing flight for Mars exploration has been generated due to the enhanced aerodynamic performance of flapping wings in the low Reynolds number flight regime.

"With a low flight Reynolds number, a conventional aircraft has a number of aerodynamic issues that severely limit its performance. The main issue is laminar separation of the boundary layer. This separation can cause loss of lift resulting in a catastrophic loss of the aircraft. To avoid this flow separation, the boundary layer must be transitioned from laminar to turbulent. Within low Reynolds number flow it is very difficult, if possible at all, to transition to a turbulent boundary layer. This flow restriction is a major factor that severely limits the flight envelope and capabilities of a conventional aircraft" [88].

Significant research into the flight of birds and insects has been conducted in an effort to quantify the complex wing motions and to measure the forces and flows around the flapping wings [89, 90, 91]. Recent developments in high-speed videography and tools for computational and mechanical modeling have revealed that the fluid dynamic phenomena underlying flapping flight are different from those of non-flapping, 2-D wings. Flapping wings take advantage of several unique unsteady aerodynamic phenomena such as delayed stall with the leading edge vortex [92, 93, 94] and wake capture [95, 96]. "It has been found that even at high angles of attack, a prominent leading edge vortex remains stably attached on the insect wing and does not shed into an unsteady wake as would be expected from non-flapping 2-D wings. Its presence greatly enhances the forces generated by the wing, thus enabling insects to hover and manoeuvre" [89].

138 Such insight has led to the development of Micro Air Vehicles (MAV) concepts [97, 98, 99] and ornithopters [29]. However, difficulties in understanding and modeling the wing motion of birds and insects, has hindered progress on the development of efficient flapping-wing aircraft. In terms of size, MAVs correspond to small birds, which do not glide like large birds, but instead flap with considerable change in wing shape during a single flapping cycle. The efficiency with which birds can achieve flapping-wing flight is attributed to the flexibility of their wings enabling them to take advantage of unsteady aerodynamic phenomena. The change in wing shape during a flapping cycle necessitates flexible airfoils posing yet another challenge in the development of flapping-wing aircraft.

The only flapping-wing device reported to have achieved short sustained flight is DeLaurier's ornithoper [29, 100]. Table B.l provides some characteristics of the proof of concept model developed in the 1980's [100].

Mass (kg) 3.96 Wing Span (m) 3.05 Wing Area (m2) 0.714 Flapping Frequency (Hz) 3.0-3.8 Engine 1 hp O.S. Max 45 FSH

Table B.l Characteristics of DeLaurier's Proof of Concept Ornithopter

On September 4, 1991, the proof of concept ornithopter was flight tested near Newton- Robinson in rural Ontario. With winds from the north at speeds of 2-7 m/s, the aircraft was launched from the top of a ridge. The ornithopter climbed and performed sustained flight for 2 min 46 sec [100].

Subsequently, a full-scale piloted ornithopter designed for ground takeoff was developed. Table B.2 provides some characteristics of the full-scale ornithoper [29].

139 Mass (kg) 322 Wing Span (m) 12.6 Flapping Frequency (Hz) 1.0 Engine 24 hp Konig SC-430

Table B.2 Characteristics of DeLaurier's Full-Scale Piloted Ornithopter

On July 8, 2006, the full-scale piloted ornithopter was flight tested. A takeoff speed of 22.4 m/s was achieved and the aircraft lifted off the runway and performed sustained flight for 14 seconds reaching a height of over one meter above the runway and a range of about a third of a kilometre. After about 10 seconds of straight and level flight, a cross wind caused the aircraft to experience roll divergence, forcing the pilot to throttle back and land the ornithopter [101].

In both cases, the wing of the ornithopter was flapped by a periodic variation of the root dihedral angle [102]. Therefore, the flapping motion of the wing was not representative of that observed in bird flight resulting in inefficient flapping-wing flight that could not be sustained for long periods of time.

It is only in recent years that we have gained some insight into the aerodynamics of flapping-wing flight. However, significant research and technological development, over the span of decades, is required to model and design flexible, flapping-wing motion that will produce efficient flapping-wing flight such as that observed in nature. As a result, serious consideration of flapping-wing aircraft for Mars exploration is premature at this time.

140 APPENDIX C Fixed-Wing Aircraft Characteristics

For an aircraft, the drag coefficient can be written as [61]:

2 CD = CDo + K CL (C.l)

where CD is the parasite (zero-lift) drag coefficient and the second term, KC\, is the induced drag coefficient.

The estimation of the parasite drag coefficient, CD , is based on the fact that a well- designed aircraft in subsonic cruise will have parasite drag that is mostly skin-friction drag plus a small separation pressure drag [61]. The latter is a fairly consistent percentage of the skin friction drag for different classes of aircraft. This leads to the concept of an

"equivalent skin friction coefficient" (Cf ), which includes both skin friction and separation drag. Cf is multiplied by the aircraft's wetted area to obtain an initial estimate of parasite drag. Thus, the parasite drag may be estimated by [61]:

CDt = Cfm ?f- (C.2)

where Cf is the equivalent skin friction coefficient, Swet is the wetted area of aircraft and S is the wing planform area. A typical value of the equivalent skin friction coefficient, Cf , for a light aircraft with a single engine is 0.0055 [61].

The wetted area of the Daedalus 88 is determined as follows:

*~*wet ~ "Wings ^Cockpit ^Fuselage "Fin ^Tail V^"*)

141 where the area of both sides (top and bottom, or left and right) of each component of the aircraft are included since both surfaces of each component are exposed to the fluid flowing past the aircraft as it flies through the atmosphere. The wetted area of each component of the Daedalus 88 is estimated from Figure 3.4 and expressed as a percentage of the wing planform area, S, leading to the following:

SCockpit=(20%S)x2

SFuselage=(lO%S)x2

SFin=(10%S)x2

STai!=(W%S)x2

Swings = ^ X 2

Therefore, CD is estimated as:

r3xS CD = 0.0055| ' = 0.0165 (C.4)

The drag-due-to-lift factor, K, is estimated by starting from classical wing theory. According to this theory, the induced drag coefficient of a 3-D wing with an elliptical lift distribution is [61]: C2 C =—L- (C.5) D' nAR v ' where AR is the Aspect Ratio.

However, few wings actually have an elliptical lift distribution. Also, this does not take into account the wing separation drag. Moreover, the presence of other non-lifting surfaces such as the fuselage, cockpit and fin, tends to decrease the efficiency of lift generation. The extra drag due to these effects can be accounted for using e, the "Oswald

142 efficiency factor", which effectively reduces the aspect ratio, producing the following equation for K [61]:

K = —— (C.6) 7tA Re

The Oswald efficiency factor for an entire aircraft is typically between 0.7 and 0.85 [61]. An Oswald efficiency factor of 0.7 is used in this analysis since this value is typical of gliders having a high aspect ratio [75].

143 APPENDIX D Airship Characteristics

The amount of buoyant lift, B, derived from lighter-than-air vehicles is based on the difference between the density of Mars' atmosphere, largely made up of CO2, and that of the hydrogen used to displace it. The buoyancy force is defined as:

B = v(pC02-pHi)g (D.l)

where v is the volume of the airship, pco represents the density of the Martian atmosphere, which is comprised primarily of carbon dioxide, and pH is the density of the hydrogen, which is used as the lifting gas. pco and pH at the cruising altitude are calculated using the ideal gas law:

(D.2) RJ

where P is the pressure, p is the density, Rg is the gas constant, and T is the absolute temperature. The gas constant for CO2 is 192 J/kg-K, and the gas constant for H2 is

8249.3 J/kg-K based on a H2 molar mass of 0.0010079 kg/mol. The density of the lifting gas and atmosphere associated with the cruising altitude of each mission are presented in Table D.l.

144 Mission 1 Mission 2 Mission 3 3 J pCOi (Atmosphere) 0.01245 kg/W 0.00944 kg/m 0.01245 kg/m

j j PH2 (Lifting Gas) 0.00058 kg/W 0.00044 kg/m 0.00058 kg/m

Table D.l Density of Lifting Gas and Atmosphere for each Mission

145