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Advanced nuclear technologies for deep space exploration

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Advanced nuclear technologies for deep space exploration

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Content 1
Advanced Nuclear Technologies for Deep Space Exploration

By

Adarsh Rajguru, BEng (Hons) in Aerospace Systems
Engineering

Master’s Thesis
Presented to the Faculty of Viterbi School of Engineering of the University of Southern
California in Partial Fulfillment of the Requirements for the Degree of

Masters of Science in Astronautical Engineering

The University of Southern California
December 2015

2

Copyright ©
by
Adarsh Rajguru
December 2015

3
The Thesis Committee for Adarsh Rajguru
Certifies that this is the approved version of the following thesis:

Advanced Technologies for Deep Space Exploration

APPROVED BY
SUPERVISING COMMITTEE:

(Name typed under line, omitting Ph.D. or Dr.)
(Name typed under line, omitting Ph.D. or Dr.)

Supervisor:
4
Dedication

To the two most important women in my life,
Mrs. Sanjukta Mohapatra
Mrs. Arda Houri Rajguru
TABLE OF CONTENTS
CHAPTER 1 – BI-MODAL RADIOISOTOPE POWER &
PROPULSION SYSTEM
I.
CONCEPT

12

II.
CORE
DESIGN

13

A. RADIOISOTOPE
FUEL
SELECTION

18

B. THERMAL
SUBSYSTEM
MODELING

24

C.
FINAL
CORE
DESIGN
SUMMARY
&
CONCLUSION

25

III.

OPERATIONAL
MODES

26

A.
POWER
CONVERSION
MODE

28

B.
PROPULSION
MODE

33

C.
HEAT
REJECTION
SYSTEM
33

IV.
MISSION
ARCHITECTURE
37

A.
DESIGN
APPROACH
37

B.
SCIENCE
OBJECTIVES
39

V.
CONCLUSION
44

VI.
REFERENCES
45

VII.
BIBLOGRAPHY
46

CHAPTER 2 – COMMERCIALIZING C-CLASS ASTEROIDS IN
THE MAIN ASTEROID BELT
I.
ASTEROID
COMMODITIES
AND
POTENTIAL
MARKETS
47

A.
WATER
47

B.
ALUMINUM,
IRON,
NICKEL,
SILICON
AND
TITANIUM
47

C.
PLATINUM
GROUP
METALS
(PGMS)
47

D.
REGOLITH
49

II.
WATER
DETECTION
50

A.
CONCEPT
OF
USING
A
NEUTRON
SOURCE
FOR
WATER
DETECTION
50

B.
NEUTRON
GENERATOR
CONCEPTS
53

C.
RADIOISOTOPE
CORE
OF
A
RADIOISOTOPE
THERMAL
ROCKET
AS
A
NEUTRON
SOURCE
DESIGN
56

D.
NEUTRON
GUN
AS
A
NEUTRON
SOURCE
DESIGN
60

E. NEUTRON
DETECTOR
64

F.
GROUND
PENETRATING
RADAR
67

G.
CONCLUSION
68

6
III.
WATER
EXTRACTION
69

A.
WATER
EXTRACTION
AND
COLLECTION
CONCEPTS
70

B.
DRILLING
EQUIPMENT
72

C.
PROPELLANT
PRODUCTION
78

IV.
CHALLENGES
OF
AN
ASTEROID
HOPPING
MISSION
83

A.
LAMBERT’S
PROBLEM
84

V.
ACQUIRING
OWNERSHIP
OF
AN
ASTEROID
89

A.
PINGER
DESIGN
89

B.
PINGER
DEVICE’S
POWER
SOURCE
93

C.
PINGER
DEVICE’S
FINAL
ASSEMBLY
AND
MASS
BUDGET
96

VI.
REFERENCES
98

CHAPTER 3 – APPLICATIONS OF RADIOISOTOPE
THERMOPHOTOVOLTAIC POWER SOURCE
I.
ONE
WAY
DEEP-‐SPACE
LASER
COMMUNICATION
FROM
A
CUBESAT
(EUROPA
ORBITER

MISSION)
103

A.
ABSTRACT
103

B.
DATA
REQUIREMENT
103

C.
RE-‐CREATION
OF
AN
EXISTING
EUROPA
MISSION
TRAJECTORY
DESIGN
FOR
RADIATION
ANALYSIS
108

D.
SENSOR
REQUIREMENTS
112

II.
ANALYSIS
OF
ENTRY,
DESCENT
AND
LANDING
(EDL)
OF
AN
AMPHIBIOUS
QUADCOPTER

SWARM
FOR
THE
EXPLORATION
OF
TITAN
119

A.
ABSTRACT
119

B.
INTRODUCTION
119

C.
SCIENCE
OBJECTIVES
119

D.
MISSION
ARCHITECTURE
120

E.
ENTRY,
DESCENT
AND
LANDING
122

F.
POWER
SOURCE
131

G.
EDL
TELECOMMUNICATION
132

H.
PAYLOAD & THERMAL CONTROL
133

III.
REFERENCES
135

IV.
BIBLOGRAPHY
137

V.
ACKNOWLEDGEMENTS
137

APPENDICES
138

7
TABLE OF FIGURES

Figure 1: Early NTP fuel concept [7] .......................................................................................... 14
Figure 2: Heat capacity vs temperature of several materials [8 – 13] ......................................... 15
Figure 3: Sensible v/s latent heat storage ..................................................................................... 17
Figure 4: From left to right:
238
PuO
2
,
241
AmO
2

and
244
Cm
2
O
3
..................................................... 18
Figure 5:
238
PuO
2
core configurations with Silicon as the thermal capacitor. ............................. 20
Figure 6:
241
AmO
2
core configurations with Silicon as the thermal capacitor. ........................... 21
Figure 7:
244
Cm
2
O
3
core configurations with Beryllium as the thermal capacitor. ..................... 22
Figure 8: 1-D Heat Transfer Analysis through the flow channel of the core .............................. 23
Figure 9: Model of core geometry ............................................................................................... 25
Figure 10: The flow schematics for the two operation modes where (a) is the thermal operation
mode and (b) is the electrical conversion operation mode. .................................................. 27
Figure 11: Schematic of the Brayton cycle for power generation. .............................................. 28
Figure 12: Illustration of Brayton cycle parameters for both working fluids during operation. . 30
Figure 13: Artistic rendering of the system engine and the dual Brayton engines ...................... 31
Figure 14: Rankine Cycle with water as the working fluid ......................................................... 32
Figure 15: Thermal Propulsion Mode .......................................................................................... 33
Figure 16: Effect of number and diameter of flow channels on absorber thermal conductance. 35
Figure 17: Variation of fluid temperature through the absorber. ................................................. 35
Figure 18: One of the six lithium cylinders comprising the absorber. ........................................ 36
Figure 19: Scrutinized hierarchy for the mission design ............................................................. 37
Figure 20: Mission Architecture Trade Tree [19] ........................................................................ 39
Figure 21: Saturn’s smaller moons – (a) Epimetheus [20], (b) Pandora [21] and (c) Telesto [21].
............................................................................................................................................... 40
Figure 22: Saturn’s larger moons – (a) Rhea [22], (b) Mimas [23], (c) Iapetus and (d) Hyperion
[24]. ....................................................................................................................................... 40
Figure 23: (a) Saturn’s ICY moon – Enceladus & (b) Enceladus’ interaction with Saturn’s “e-
ring” [25] ............................................................................................................................... 41
Figure 24: Cassini’s CIRS instrument data of Enceladus South Pole [19]. ................................. 42
Figure 25: Artistic rendering of the 6U toroidal cage. ................................................................. 43

8
Figure 26: Platinum (top left), Palladium (top right) and Rhodium (bottom). Platinum has high
melting point (1772
0
C) and is very stable at high temperatures. Palladium and Rhodium are
both a silvery white metal. Rhodium is used as electrodes for aircraft spark plugs [2]. ...... 48
Figure 27: Artist’s impression of ISRU based robotic construction technologies using asteroid
regolith. This picture is just an illustration of a futuristic space based 3D printer using
asteroid regolith material to fabricate habitat structures [5]. .............................................. 49
Figure 28: The probability of interaction of a neutron with another atom’s nucleus based on
neutron energy & cross section [9]. ...................................................................................... 51
Figure 29: Neutron spectra returned as a function of water content [10]. ................................... 53
Figure 30: Mapper spacecraft after the deployment of lithium acceleration and deceleration
electrostatic columns [10]. .................................................................................................... 53
Figure 31: Spontaneous fission of a large unstable molecule [12]. ............................................. 54
Figure 32: The actual energy spectrum of neutrons from spontaneous fission of
244
Cm [13]. ... 55
Figure 33: Energy spectrum of neutrons from the spontaneous fission of the generated
244
Cm
source in MCNP. ................................................................................................................... 55
Figure 34: Shows the interaction probability of a neutron with a beryllium nucleus based on
neutron energy dominated by elastic scattering [9]. ............................................................. 56
Figure 35: MCNP plot of a radioisotope thermal rocket using tungsten encapsulated
244
CmO
2

fuel rods embedded inside Beryllium acting as a thermal capacitor and neutron reflector. . 57
Figure 36:
244
Cm with Beryllium encapsulated core configurations 1 – 3 designed in MCNP. . 58
Figure 37:
244
Cm with Beryllium encapsulated core configurations 4 – 6 designed in MCNP. . 59
Figure 38: Neutron Gun Design. .................................................................................................. 61
Figure 39: The detected reflected neutron spectrum from the neutron gun with varying distance
from the surface of the asteroid. ........................................................................................... 61
Figure 40: Increasing neutron count with increasing water content in the asteroid composition 62
Figure 41: (Top plot) 5 meters distance away from asteroid. ...................................................... 63
Figure 42: The primary neutron interactions with helium, lithium and boron [9]. ...................... 65
Figure 43: Scintillation detector. [14] .......................................................................................... 65
Figure 44: A typical neutron scintillation detector [10]. ............................................................. 66
Figure 45: System Architecture ................................................................................................... 68

9
Figure 46: Phase chart of water depending on the temperature and pressure of the surroundings.
............................................................................................................................................... 69
Figure 47: Illustration showing most of the water molecules escaping near the surface and the
rest of the water molecules spreading out in all directions [19]. .......................................... 71
Figure 48: (Left) Auger drill bit [20] and (Right) Internal sample collecting drill bit [21]. ........ 72
Figure 49: The dimensions of the described Auger drill bit (not to scale). ................................. 74
Figure 50: The dimensions of the required housing for the Auger drill bit (not to scale). .......... 74
Figure 51: The dimensions of the hollow drill bit [21]. ............................................................... 75
Figure 52: The dimensions of the electric coil. ........................................................................... 75
Figure 53: Water condensation system ........................................................................................ 78
Figure 54: Possible set up the electrolysis system that could be used on the spacecraft. ............ 80
Figure 55: Hydrogen gas and oxygen gas production by volume v/s electric current. ................ 82
Figure 56: Mass flow rates of hydrogen gas and oxygen gas produced v/s electric current. ...... 82
Figure 57: Asteroids in the Solar System (left) and half of the known asteroids in the main
asteroid belt (right) [36]. ....................................................................................................... 83
Figure 58: Illustration of Lambert’s problem [37, Fig. 5.3]. ....................................................... 85
Figure 59: Trajectory for 100 hops. ............................................................................................. 85
Figure 60: Orbital elements of all known asteroids in the main asteroid belt. ............................ 86
Figure 61: Prediction of 10
7
asteroids not discovered and logged yet. ........................................ 86
Figure 62: 10
7
asteroids reduced to 600,000 to handle computation load. .................................. 87
Figure 63: Main belt asteroid size – frequency distributions [38]. .............................................. 88
Figure 64: ΔVs of 100 best hops from 2.5 x 10
9
asteroids choices. ............................................ 88
Figure 65: Pinger device’s communication system architecture. ................................................ 91
Figure 66: Typical physical parameters of a helical antenna. ...................................................... 92
Figure 67: City Labs’ NanoTritium™ beta-voltaic power source [50]. ...................................... 94
Figure 68: Construction of the radioisotope beta-voltaic cell [51]. ............................................. 94
Figure 69: Energy flux v/s distance between the beta-voltaic material and the receiver plate. ... 95
Figure 70: Length of beta-voltaic surface and the receiver p-n diode plate v/s plate gap. .......... 96
Figure 71: Pinger device’s penetrator probe. ............................................................................... 97
Figure 72: X-123CdTe (X-ray and Gamma-ray detector system [2], Argus Infrared Spectrometer
[3] and NanoCam C1U (High Resolution Camera) [4] ...................................................... 105

10
Figure 73: Laser Anemometer and Martian Dust Analyzer (LAMBDA) [7], Low Voltage Gated
Electrostatic Mass Spectrometer (LVGEMS) [5] and Highly Integrated Micropayload for
Broadband Infrared Spectrometry (HIBRIS) [6] ................................................................ 105
Figure 74: Data rate requirement to portray the quality of information [1 – Fig 1.1] ............... 106
Figure 75: Orbital Altitude (km) v/s Picture Count for Jupiter – Europa Orbiter Mission ....... 108
Figure 76: Parking orbit altitude (km) v/s ground image area (km
2
) for Europa Orbiter. ......... 109
Figure 77: Top view of Earth – Jupiter Cruise Phase (Refer Table 37 in the Appendix) .......... 109
Figure 78: Jovian Moons Tour Phase (Refer Table 37 in the Appendix) .................................. 110
Figure 79: (Below) Combined Electron and Proton Fluence Output for the Europa Orbiter
Mission. ............................................................................................................................... 111
Figure 80: Calculated dosage based on 7 mm Aluminum shielding. ........................................ 111
Figure 81: Fluence (cm
-2
) v/s Energy (MeV). ........................................................................... 112
Figure 82: Block diagram showing jitter isolation and rejection for a laser communication
terminal [1 – Fig 2.10] ........................................................................................................ 113
Figure 83: Comparison of RF and optical beam spreads from Saturn [1 – Fig 1.2]. ................. 114
Figure 84: ATP subsystem in the laser transceiver using Earth as the beacon [15 – Fig 3.11]. 114
Figure 85: Wavelengths of commercially available lasers [16]. ............................................... 116
Figure 86: Top view of the laser transmitter layout [1 – Fig 5.15]. ........................................... 116
Figure 87: (a) Left – Cross-section and bias supply of the InGaAs pin-photodiode [18 – Fig 7.6]
(b) Right – Equivalent circuit of the pin-photodiode. [18 – Fig 7.8] .................................. 117
Figure 88: CAD model of the entry vehicle .............................................................................. 121
Figure 89: EDL Concept of Operations (CONOPS) ................................................................. 122
Figure 90: Monte Carlo Simulation Results showing Altitude (km) v/s Peak Heat Flux (W/cm
2
)
............................................................................................................................................. 123
Figure 91: Monte Carlo Simulation Results showing Altitude (km) v/s Peak Deceleration (g) 123
Figure 92: Trajectory plots based on optimum input parameters. ............................................. 126
Figure 93: Heating plots based on optimum input parameters .................................................. 127
Figure 94: Mach No. v/s drag coefficient of the supersonic parachute ..................................... 128
Figure 95: Supersonic trajectory plots ....................................................................................... 128
Figure 96: Altitude (km) v/s Density (kg/m
3
) at Titan .............................................................. 129
Figure 97: CAD models of the quadcopter ................................................................................ 129

11
Figure 98: Power required by the quadcopters to hover ............................................................ 130
Figure 99: PID control system for quadcopters [62] ................................................................. 130
Figure 100: CAD model of the RTPV ....................................................................................... 131
Figure 101: Subsystem block diagram ....................................................................................... 134
Figure 102: Landing Ellipse [67] ............................................................................................... 134
Figure 103:
244
Cm
2
O
3
core configurations with Beryllium as the thermal capacitor. ............... 138
Figure 104: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 1. ........................................................................... 139
Figure 105: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 2. ........................................................................... 140
Figure 106: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 3. ........................................................................... 140
Figure 107: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 4. ........................................................................... 141
Figure 108: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 5. ........................................................................... 141
Figure 109: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 6. ........................................................................... 142
Figure 110: Circuit Diagram of Pacific Monolithics PM2105 Power Amplifier ...................... 142
Figure 111: Illustration of the pinger device’s BPSK modulation scheme ............................... 143
Figure 112: Data downlink transmittal time (hours) v/s picture count ...................................... 144

12
Chapter 1

Bi-Modal Radioisotope Power & Propulsion System

I. CONCEPT

The following study was performed under the supervision of Dr. Nathan Jerred (Principal
Investigator) of the Center for Space Nuclear Research (CSNR) under his NIAC Phase I award.
The proposed bi-modal radioisotope-based power and propulsion system concept would leverage
the high specific energies [J/kg] associated with radioisotope materials and enhance their
inherent low specific powers [W/g]. This is accomplished by accumulating thermal energy from
nuclear decay within a central core over time, which allows for significant amounts of power to
be transferred to a flowing gas over short periods of time. In the proposed configuration the
stored energy can be utilized in two ways:
1. with direct propellant injection through the core, the energy can be converted into thrust
using of a converging-diverging nozzle and
2. by flowing a working fluid through the core as a subsequent Brayton engine, energy
within the core can be converted to electrical energy.
The first scenario achieves moderate ranges of thrust, but at a higher Isp than traditional
chemical-based systems. The second scenario allows for the production of electrical power,
which is then available for electric-based propulsion. Additionally, once at location the
production of electrical power can be dedicated to the deep-space communication for data
transfer. Ultimately, the proposed dual-mode propulsion platform capitalizes on the benefits of
two types of propulsion methods – the thrust of thermal propulsion ideal for quick orbital
maneuvers and the specific impulse of electric propulsion ideal for efficient interplanetary
travel. Overall, the system is functioning as a radioisotope thermal rocket (RTR).

The functionality of the overall system relies on the integration of several key
components – the energy source, thermal storage media, an insluation scheme, gas flow design,
energy conversion and propulsion system. At the center of the propulsion system is the

13
radioisotope source. For safety and retention, the radioisotope fuel will be encapsulated within a
tungsten-based matrix. The resulting fuel rods will be integrated within a central core material.
The ideal core material must be capable of storing thermal energy, acting as a thermal
capacitor, and then dissipate that energy to a flowing gas. Several materials have been identified
elsewhere as being capable of achieving this task relying on their specific heat capacities, e.g.,
beryllium and boron tetra-carbide [1]. Instead, in this study the use of silicon as a thermal
capacitor material is being considered. Silicon undergoes a latent heat of fusion (ΔH
fusion
= 50.2
kJ/mol) at 1685 K [2]. By taking advantage of silicon’s storable energy, when gas is flowed
through the silicon core its phase transforms from liquid to solid. This in turn, dissipates energy
from the core to the gas at a constant core outlet temperature, yielding a constant chamber
temperature or turbine inlet temperature depending on the operational mode. For heat rejection,
turbine exhaust gases will be passed through flow channels in a solid lithium block. Having a
high heat capacity, the lithium block absorbs the thermal energy from the gas, which is then
allowed to dissipate slowly between pulses. This method has the potential to deliver a low mass,
compact heat rejection subsystem [3].
II. CORE DESIGN

As mentioned the system concept relies on the decay energy from radioisotopes.
Radioisotopes in general exhibit very high specific energies [J/kg], however, they have poor
specific powers [W/g]. Several radioisotopes have the potential to be used for the system
concept. Table 1 tabulates several potential radioisotopes and their properties.
Isotope Specific Power [W/g] T
1/2
[yrs]
238
Pu*

0.392 87.7
90
Sr†

0.254 28.8
244
Cm‡

2.269 18.1
241
Am
§
0.094 432.7
*Assumes 80% isotopic purity and 88% compound mass in
238
PuO 2
†Assumes 57% isotopic purity and 48% compound mass in
90
SrTiO 3
‡Assumes 90% isotopic purity and 91% compound mass in
244
Cm 2O 3
§
Assumes 98% isotopic purity and 88% compound mass in
241
AmO 2
Table 1: Tabulated values of radioisotopes [4, 5].
Compared to the other radioisotopes presented in the above table
238
Pu has a good
specific power and a long half life. Additionally, this plutonium isotope has a long historical

14
use in NASA and is already flight qualified, being used for numerous NASA deep space
missions, e.g., New Horizons, Curiosity, Cassinni, etc.
Housing the radiosotope can be accomplished by directly encapsulating it within a
tungsten – rhenium matrix to form the radioisotope heat source (RHS) for the system. This
encapsulation concept relies on the radioisotope-of-choice to be fabricated in to microspheres
(dia. ≈ 100 µm), which are then directly sintered in to a tungsten-based matrix. In the design of
the system to properly evaluate masses, this study utilized an isotope loading within the
tungsten-rhenium matrix of 50% by volume. Where the encapsulation matrix was comprised of
tungsten – 25 at. % rhenium metals.
A solid, tough, high-temperature tungsten-rhenium matrix can be formed to encapsulate
radioisotopes commonly used for power production [6]. The idea is that this tungsten-based
matrix would be robust enough and provide the required strength to prevent the dispersion of
the radioisotope inventory in launch abort scenarios, atmospheric re-entries and planetary
impacts in the case of a failed in-situ probe deployment. By using this encapsulation method,
the proposed fuel forms exhibit an increased energy density that is nearly five times greater
than the traditional GPHS units.

Figure 1: Early NTP fuel concept [7]
The primary component of the conceptual system is the thermal capacitor, whose
functionality drives the entire RTR concept. As previously described the thermal capacitor
accumulates thermal energy from the radioisotopes over time. Then the stored thermal energy
is extracted quickly by a flowing gas. Depending on the gas used the extracted thermal energy

15
can be converted to thrust by use of a converging-diverging nozzle or converted to electrical
power through the use of an energy conversion system. In determing an adequate material to
act as the thermal capacitor several qualifications must be met:
1. The material must have high thermal storage capabilities
– to accumulate a large amount of energy within a given volume
2. The material must have a high thermal conductivity
– to dissipate stored thermal energy quickly to a flowing gas
3. The material must have a high melting temperature
– allows for a high operational temperature increasing the systems performance
Thermal storage can primarily be accomplished through two methods – sensible heat
storage and latent heat storage. Sensible heat storage is the energy stored in a material over a
certain temperature range and is described by the material’s specific heat capacity [C
p
]. Several
materials have excellent sensible heat storage and the graph given in Figure 2 shows the heat
capacity of several materials plotted over a given temperature range.

Figure 2: Heat capacity vs temperature of several materials [8 – 13]
Beryllium is seen to be an excellent candidate material for sensible heat storage (C
p
=
1.83 J/g-K, T
melt
= 1551 K & k = 201 W/m-K) allowing for an operational temperature of 1200

16
K. It has the potential to store over 2 MJ/kg over the temperature range of 500 – 1200K.
Additionally, boron would also make a good thermal capcitor material (ΔH
fusion
= 2.1 MJ/kg,
C
p
= 1.03 J/g*K, T
melt
= 2348 K & k = 27 W/m*K) [14].
Boron allows for a higher operational temperature, which in turn allows for a greater
amount of energy storage due to the larger operational temperature. But Boron wasn’t chosen
because its use required an operational temperature greater than 2000 K in order for its thermal
storage potential to exceed that of beryllium. Thermal cycling at those temperatures could
present a significant challenge to the system and reaching those temperatures using
radioisotopes would be equally challenging.

Latent heat thermal storage is the energy stored in a material through its phase change
and is described by a material’s latent heat of fusion [ΔH
fusion
]. Several phase change materials
(PCM) can be utilized depending on the application. Terrestrial based systems using latent heat
thermal storage typically use molten salts as their PCM. However, their low melting
temperatures and their low potential to store thermal energy do not make them ideal for this
application. Silicon was determined to be the ideal PCM (ΔH
fusion
= 1.8 MJ/kg, T
melt
= 1687 K
& k = 148 W/m-K) matching the storage performance of beryllium [2]. And because melting
silicon is the primary goal, using it allows for an operational temperature approaching 1700 K.
Several PCM materials are tabulated in Table 2 for comparison.
Material ΔH
fusion
[MJ/kg] T
melt
[K] K [W/m-K]
Silicon 1.80 1687 148
Boron 2.09 2348 27
LiF* 1.04 1121 --
LiH* 2.58 956 --
80LiOH + 20LiF* 1.16 700 --
*molten salts
Table 2: PCM Materials [2, 14, 15]
Sensible heat storage systems exhibit non-isothermal behavior as they discharge their
stored energy. This equates to a continually decreasing core exit gas temperature through the
blowdown process. In turn this means the chamber temperature of the propellant gas or the
turbine inlet temperature is constantly changing, complicating the design of these subsystems.
In general latent heat storage systems are favorable because their temperature is held relatively

17
constant at the phase change temperature as they accumulate and discharge energy. This
isothermal behavior simplifies the system design and limits its thermally cycling. Figure 3
shows a simplified example of sensible vs latent heat storage over a temperature range.

Figure 3: Sensible v/s latent heat storage
For the concept presented here the thermal capacitor was determined to be silicon.
Silicon, which simplifies the system’s design, exhibits a high energy storage potential,
operational temperature and thermal conductivity. The major technical challenges in using
silicon as the thermal capacitor are (1) containing its liquid phase and (2) volume change of the
silicon thermal capacitor while handling its liquifying – freezing cycle.

During a blowdown sequence as energy is dissipated and the core re-solidifies uneven
freezing can form void spaces, which in turn can introduce stresses into the insulation layers
surrounding the central core. To deter possible insulation fracturing the thermal capacitor can
be first contained within a structural canister that can withstand core volume fluctuations
through the phase change cycling. At the proposed operational temperatures this housing will
most likely be a refractory metal or alloy, e.g. a molybdenum or tungsten based alloy
exhibiting ductile behavior.

The silicon core canister would be fabricated as a shell with tubes running axially acting
as the flow channels. The canister wall thickness will be several millimeters in the periphery
but with a minimal wall thickness at the flow channels. The distance between flow channels
(web thickness) has to be minimal to minimize hot spots and to ensure all stored energy can be
extracted. However, these parameters could also affect the stresses associated with the silicon
freezing cycle. One thought was to increase the web thickness to allow more expansion for

18
silicon before encountering another flow channel tube. This might also minimize the applied
stresses to the canister outer wall. Computational fluid dynamics (CFD) modeling was carried
out to obtain the ideal flow channel size and web thickness (discussed later).

When silicon liquefies its volume decreases by up to 8% leaving void spaces within the
canister. The formation of these voids also has the potential to create a loss in conductive
pathways to the walls of the canister and flow channels. Modeling alone will not be sufficient to
study the effects of these voids to the thermal hydraulics of the core. It has to be experimentally
investigated in order to gain a better understanding of the volume change phenomenon.
A. Radioisotope Fuel Selection
Three different radioisotope fuels were thermally analyzed to size the core for a thermal
power output of 1200 W. These fuels were: (1)
238
PuO
2
, (2)
241
AmO
2

and (3)
244
Cm
2
O
3
. The
densities, volume and mass (with matrix) of the radioisotopes fuel oxides are tabulated in Table
3 below for the thermal output of 1200 Watts.

Radioisotope Density (g/cm
3
) Volume (cm
3
) Mass with matrix (kg)
238
PuO
2
11.50 521.8 8.139
241
AmO
2
11.68 2142 33.61
244
Cm
2
O
3
11.60 88.68 1.392
Table 3: Radioisotope options for a thermal power output of 1200 W

Figure 4: From left to right:
238
PuO
2
,
241
AmO
2

and
244
Cm
2
O
3

2-D thermal analysis of six different core configurations for each radioisotope fuel types
was investigated. So in total 18 unique core designs were evaluated. Figure 5 - Figure 7 show
the core configuration designs for
238
PuO
2
,
241
AmO
2

and
244
Cm
2
O
3
each respectively. In these

19
figures the larger circles filled with temperature color maps represents the fuel rod elements and
the smaller white circles represent the flow channel elements. Table 31 - Table 33 in the
appendices, lists all the cores specifications for
238
PuO
2
,
241
AmO
2

and
244
Cm
2
O
3
.
In each of the figures below it can be seen that the core reaches the operational
temperature of 1683 K following the meltdown of the thermal capacitor, in a certain number of
hours. The ideal core for each fuel type is chosen based on the configuration which reaches the
operational temperature quickest.

Core 1 – 17 Hours Core 2 – 17 Hours

Core 3 – 19 Hours Core 4 – 16 Hours

20

Core 5 – 14 Hours Core 6 – 15 Hours
Figure 5:
238
PuO
2
core configurations with Silicon as the thermal capacitor.

Core 1 – 19 Hours Core 2 – 18 Hours

Core 3 – 19 Hours Core 4 – 19 Hours

21

Core 5 – 19 Hours Core 6 – 17 Hours
Figure 6:
241
AmO
2
core configurations with Silicon as the thermal capacitor.

Core 1 – 21 Hours Core 2 – 19 Hours

Core 3 – 21 Hours Core 4 – 22 Hours

22

Core 5 – 19 Hours Core 6 – 18 Hours
Figure 7:
244
Cm
2
O
3
core configurations with Beryllium as the thermal capacitor.
From the figures above it can be seen that core 5 for the
238
PuO
2
fuel is the best core
configuration for
238
PuO
2
core with Silicon as the thermal capacitor. Similarly, core 6 for the
241
AmO
2
fuel is the best core configuration for
241
AmO
2
with Silicon as the thermal capacitor.
Also core 6 for the
244
Cm
2
O
3
fuel is the best core configuration for
244
Cm
2
O
3
with Beryllium as
the thermal capacitor.

Even though the
244
Cm
2
O
3
core is the lightest, it still takes four hour more than the
238
PuO
2
core to reach operational temperature. Also
244
Cm
2
O
3
has a way shorter half life
compared to
238
PuO
2
and has high energy gamma signatures which makes it very difficult to
handle or fabricate. Due to the sparcity of
238
PuO
2
in the United States,
241
AmO
2
as an alternative
was studied above. The above analysis suggests that the power & propulsion system can use an
Americum Oxide core but it will be a much heavier and voluminos core compared to the
238
PuO
2
core. Nevertheless it would still take three hours longer than the
238
PuO
2
core. Thus
238
PuO
2
core
was considered as the ideal candidate for this power & propulsion system.

Silicon thermal capacitor in the core has 280 flow channels (d = 5 millimeters). 1D heat
transfer analysis was conducted by discretizing along the length of a single flow channel and
assuming constant surface temperature (T
S
= 1683 K, melting point of Si).

23

Figure 8: 1-D Heat Transfer Analysis through the flow channel of the core

à Equation 1

à Equation 2

Thermal power transferred from core to fluid is:

à Equation 3

And fluid carries away:

à Equation 4

Equating the two allows to solve for T
OUT
:

à Equation 5

Where, à Equation 6

Heat transfer coefficient h is a function of fluid's thermal conductivity k, and Nusselt number:

à Equation 7

Gdielinski correlation of Nusselt number was used for turbulent flow. For laminar flow
Nu = 3.66. Flow in the core is firmly laminar. Reducing the number and diameter of flow
channels would allow turbulent flow, which would improve heat transfer to fluid. However,
forcing turbulent flow isn't necessary for the given power requirement, and tinkering with core
geometry would change its melting characteristics.
T
MEAN
= (T
IN
+ T
OUT
)/2
T
FILM
= (T
S
+ T
MEAN
)/2
˙
Q
IN
= hA
s
(T
S
− T
FILM
)
˙
Q
OUT
= ˙ mC
p
(T
OUT
−T
IN
)
T
OUT
=
K (T
S
/2− T
IN
/4)+ T
IN
1+ K/4
K=
hπDdL
˙ mC
p
h=
Nuk
D

24
B. Thermal Subsystem Modeling

During designing the model of the core, the thermal and radiative losses were somewhat
problematic, as working with such high temperatures allowed the heat to escape through the
mounting systems or by radiation. To combat both effects, an insulating material was included
on the outer walls of the core. This allowed the core to stay hot enough to melt the silicon, while
keeping the outside walls of the insulation cool enough to limit radiative losses. It also doubled
as a structural mounting material along with being a conductive loss limiter. Based on thermal
and structural properties, Zirconia (ZrO
2
) was selected to be the primary insulator. Carbon
Aerogel was chosen as a secondary insulator to be used in areas where no stresses will be
experienced. Table 4 shows the material properties of the insulation materials that were used in
the modeling.
Insulating Material Specific Heat [J/kg*K] Thermal Conductivity [W/m*K] Density [kg/m
3
]
Zirconia 400 3 5700
Carbon Aerogel 754 0.03 2230
Table 4: Material properties of insulation materials used in modeling [2, 14, 15]

For the thermal propulsion mode, hydrogen was chosen as the ideal primary propellant.
In using hydrogen the zirconia insulation may be placed in a reducing environment, inadvertently
degrading the insulating material. Because of this, boron nitride and zirconium carbide were
examined as alternatives. However, no other option had the same thermal insulation
characteristics while maintaining the structural integrity necessary to support the core. Thus the
use of a hydrogen-compatible material, such as boron nitride or tantalum would be used as a
possible cladding to form a protective barrier for the zirconia. The cladding’s thermal expansion
was also considered in order to ensure stability through thermal cycling. The design for the core
was concluded with the following assumptions:
(1) Zirconia insulation would incorporate a protective cladding from the hydrogen,
(2) The metallic housing for the molten silicon had negligible thermal effects and
(3) The silicon expansion could be overcomed.
With these aspects in mind, the core is radially insulated by a zirconia sheath with a
thickness of 5 cm and axially insulated by a zirconia cap (top and bottom), each having an 18.5
cm diameter and 20 cm length. Around that assembly is a carbon aerogel secondary insulation

25
layer, having a 40 cm diameter and 70 cm length. Four tantalum rods of 0.4 cm diameter
attached the zirconia insulation to the housing of the unit to act as a support structure for the
overall core assembly. With such small core dimensions, conductive losses at these points can be
significant. A diagram of the preliminary core design can be seen below in Figure 9.

Figure 9: Model of core geometry
C. Final Core Design Summary & Conclusion

Core 5’s thermal analysis shows that fuel rods of varying sizes radially spaced within the
silicon core provided the most equal dissipation of thermal energy into the thermal capacitor.
The 2-dimensional model did take into consideration the phase of the material at different
temperatures and the absorption of the input heat by melting. At the operating temperature the
thermal models predict that 815 W
t
of thermal energy would escape through the insulation due to
conductive and radiative losses. This leaves 365 W
t
from the total 1200 W
t
provided by the
radioisotope to melt the core.

While the model did take in to account the length of the core for the purposes of
calculating exposed areas, it did not examine the changes to the melting profile near the edges. It
was assumed for this profile that edge effects could be neglected. As the model ran, the silicon
material absorbed the input energy, as it went through the phase change. Only once the material
had melted, was the temperature allowed to increase. This created a melting profile that slowly
extended away from the hot fuel rods, until the entire core had become liquid.

26
The core is powered by 3.79 kg of
238
PuO
2
, having a power density of 0.392 W/g, for a
total of 1.2 kW
t
of input heat. The total mass of the loaded fuel rods will be 8.139 kg, and the
core uses 15.58 kg of silicon, giving the core a total mass of ≈
23.72 kg. The core is designed to
operate at 1683 K continuously, and it would alternate between totally molten to totally solid.
The total meltdown of the thermal capacitor inside the core will take approximately 14 hours to
complete. The silicon thermal capacitor was capable of storing 28 MJ of energy, which if
discharged over 360 seconds (6 minutes) and converted to electricity at 24 % efficiency, would
provide 18.67 kW
e
of electrical power. A mass breakdown of the radioisotope core is shown in
Table 5 below:
Core Component Mass [kg]
Silicon PCM

15.58
PuO
2
Loaded Fuel Rods

8.139
ZrO
2
Insulation 108
Mounting Structure 2.1
Total Mass 133.82
Table 5: Mass or the total core and the various components

III. OPERATIONAL MODES

The concept relies on the function of two modes to accomplish the overall goals of a
mission. At the center of the operation is the thermal capacitor, discussed above. The thermal
capacitor accumulates thermal energy that can be made available for different operations and / or
functions of the entire system. The two operational modes discussed in greater detail below are
the thermal mode and electrical conversion mode. The thermal mode takes advantage of the
stored thermal energy transferring it to propellant injected in to the core. In turn, the now
energized propellant flows through a converging-diverging nozzle creating thrust.

The second operational mode is converting the stored thermal energy to electrical power
to be used for electric propulsion, communications, etc. There were two primary energy
conversion methods identified that could be used with this system – thermal photovoltaic (TPV)
or a Brayton cycle. Each system provides power in two drastically different manners.

27
A TPV system can be designed to utilize thermal radiation from the core and
convert it to useable electrical power. This adds an element of complexity in designing the
overall core system. In one hand the thermal capacitor is insulated to reach a certain operational
temperature, however, on the other hand radiative losses are needed to provide electrical power.
A TPV system is a solid-state conversion method that can provide continuous power, but would
need a capacitor bank when the system requires bursts of high power.

A Brayton-based conversion system is a dynamic cycle and when pulsed can produce the
bursts of higher power that may be needed by a communication subsystem. The primary choice
of the Brayton system was due to the high power requirements initially identified by the
communication system and its ability to be paired with the thermal capacitor. The operation of
the Brayton system comprises of passing a working fluid through the thermal capacitor,
extracting the stored thermal energy and converting that to electrical power through the use of a
turbine and alternator. Figure 10 shows the two flow schematics representing each operational
mode and how they may be integrated in to the system. The electrical conversion mode will be
discussed in detail below, whereas the thermal mode and its operation is discussed in greater
detail later.

(a) (b)
Figure 10: The flow schematics for the two operation modes where (a) is the thermal operation
mode and (b) is the electrical conversion operation mode.

28
A. Power Conversion Mode

For power generation, a Brayton cycle was selected based on its high efficiency, low
waste heat, and ability to utilize the thermal capacitor of the core effectively. A Rankine cycle,
for comparison, needs to reject large amounts of heat to return the working fluid to the beginning
liquid state. Instead, a Brayton cycle’s major loss mechanism is the power taken by the
compressor to return the fluid to the starting state.

While both cycles are quite efficient, rejecting waste heat in space usually requires the
use of massive radiators that contribute greatly to the size and weight of the craft. To tackle the
heat rejection, fluid in the Brayton cycle will be passed through an absorber material to collect
the remaining waste heat. After the 360 seconds blowdown, the core would slowly melt again
and the absorber could safely radiate the waste heat to space without the need for large fins. It
will be ready for the next cycle after 14 hours.

Figure 11 shows the schematic of the Brayton cycle power generation system. The
objective was to design a Brayton cycle engine, which could provide up to 25 kW of electrical
power for six minutes in every 15 hours. Also, the turbine inlet temperature was not to exceed
1100 K to limit thermal deterioration of the blades overtime.

Figure 11: Schematic of the Brayton cycle for power generation.

29
Absorber and working fluid temperatures increase during the six minute power
generation period (blowdown), because heat is stored in the absorber to be radiated to space
later. This makes the whole cycle time-dependent. Core temperature is constant over 1683 K
(melting point of silicon), so increasing fluid temperature reduces heat transfer at the core, and
thus power extracted at the turbine. Therefore absorber needs to have a sufficient thermal mass to
prevent power output from dropping too much (< 1 kW) towards the end of the blowdown.
Compressor and turbine were modeled using isentropic flow equations with efficiency
corrections. Outlet temperatures are given by:

=

!

à Equation 8

=

!

à Equation 9
And shaft powers by,

=

(−

)
!

à Equation 10

=

(

!

!)

à Equation 11
Pressure ratio of compressor and turbine was chosen to be r
p
= 4, which is typical for
Brayton cycles. Compressor, turbine and generator were all assumed to have efficiencies of η
T

=
η
C

= η
G

= 0.9. A typical power plant pressure loss parameter, β = 1.05, was assumed.

The thermodynamic analysis of the Brayton cycle resulted in Helium as the best choice of
working fluid. Table 6 presents the studied power cycles including the necessary core and
absorber masses. It turned out in all cases that nominal amount of Plutonium (3.3 kg) was not
enough to provide energy for six minutes blowdown to produce 25kW
e
. Mass of the rest of the
core was scaled up linearly with increasing Pu mass.

Only the length of the core was adjusted to provide sufficient heat transfer to working
fluid. Tinkering with number of flow channels (N = 280) and their diameter (d = 5 mm) would
change melting characteristics of the thermal capacitor.

30
Fluid m
Pu
[kg] m
CORE
[kg] L
CORE
[m] m
ABSORBER
[kg] Power conversion efficiency
He

4.0 159 0.24 200 24%
H
2
5.5 219 0.32 200 18%
Water 5.6 222 - 230 17%
Table 6: Comparison of power cycles with different working fluid.

Input Variable Value Units
Fluid

Helium
Mass Flow

32.2 g/s
Pressure Ratio 4
Compressor Efficiency 85 %
Turbine Efficiency 88 %
Alternator Efficiency 98 %
Motor Efficiency 98 %
DC to AC Conversion Efficiency 97 %
AC to DC Conversion Efficiency 97 %

Table 7: Input variables for Helium as the working fluid.

Figure 12: Illustration of Brayton cycle parameters for both working fluids during operation.

31
This configuration provided 24.7 kW of electrical power to be used for powering
communications systems or electrical propulsion systems at 31.38% total conversion efficiency.
The analysis of components was based largely on a report by the CSNR in 2010 for a similar
pulsed power system [3]. The analysis used a turbine / compressor combination in a similar
fashion, but was sized to produce only 10 kW. The results of that analysis were scaled linearly
to approximate the masses of the turbine and compressor combination used to produce 25 kW
and match the thermodynamic cycle properties used. The resulting masses for the turbine and
compressor were estimated to be 54.1 kg. A single, 30 kW peak power off-the-shelf alternators
provided the mass estimate for the electrical conversion system, which had a total mass of 15.9
kg [16]. A budget of 10 kg was allowed for miscellaneous components such as piping. In total,
the mass of the power conversion system with the core would be 229 kg. Table 8 tabulates the
system components and Figure 13 shows an artistic rendering of the main engine of the concept
with the dual Brayton engines.
Cycle Component Mass (kg)
Core

132.12
Turbine and Compressor

54.1
Absorber Mass 16.9
Alternator 15.9
Housing 10
Total Mass 229.02

Table 8: Mass budget of the power conversion system.

Figure 13: Artistic rendering of the system engine and the dual Brayton engines

32
Idea of using water as propellant means that power generation becomes Rankine cycle if
same turbomachinery is to be used for both modes. Due to complications in calculating heat
transfer when vaporizing water, the cycle was analyzed by keeping track of enthalpies. It
necessitates assumption of constant rate heat input in the core and constant rate heat rejection in
the absorber:
à Equation 12
Where, E
core
is initial thermal capacitor energy content in Joules
à Equation 13
Heat rejection rate follows from restricting the temperature rise of the absorber to (T
Li max

– T
Li initial
) = 50 K. In effect average rates are used throughout the simulation instead of actual,
time dependent values. Another consequence of not performing heat transfer analysis is that flow
channel geometry of the absorber cannot be determined. Thus it becomes a separate task to
determine a geometry that provides assumed heat transfer rate with allotted absorber mass.

Larger pressure ratios are possible in Rankine cycles than Brayton cycles. However,
going below P
min
= 0.5 bar, made the cycle very sensitive to the absorber mass. At a certain mass
of the absorber water wouldn't condense completely, and a fraction of a kilogram lower, resulted
to water freezing in the absorber. P
max
= 10 bar was selected by keeping in mind weight of the
plumbing. Pump, turbine and generator efficiencies were assumed to be 0.9. Absorber was
required to cool water below boiling point, so that it wouldn't vaporize and cavitate in the pump.

Figure 14: Rankine Cycle with water as the working fluid
˙
Q
IN
=
E
core
t
blowdown
˙
Q
REJ
=
(T
Li max
−T
Li initial
)m
Li
C
p Li
t
blowdown

33
B. Propulsion Mode

Propulsion configuration is depicted in Figure 15. Propellant (hydrogen) is pumped from
tank to the core, heated and then expelled through the nozzle. Nozzle was assumed to have
typical space nozzle thrust coefficient C
F
= 1.78.

Figure 15: Thermal Propulsion Mode
Performance parameters given in Table 9 for hydrogen as the propellant shows the
maximum specific impulse, which could be achieved. Compressor is driven by an electric motor
that draws it power from spacecraft batteries. As long as they can provide sufficient power, Isp
can be traded for higher thrust by increasing mass flow, if required.
Cycle
Mass flow rate
(g/s)
P
COMP

(kW)
T
CORE_OUT

(K)
Isp (s) F
th
(N)
T
burn
(s)
H
2
4 9 1664 697 27 641

Table 9: Propulsion Performance.
C. Heat Rejection System

The heat rejection system being employed in the concept is an enabling technology to the
concept that leads to a compact, low mass system. Because the Brayton engines are pulsed, heat
rejection does not have to be instantaneous, but instead can be carried out over long periods of
time when the cycle is not operating. Normal space radiators reject waste heat by radiative heat
transfer, and their effectiveness depends on the acceptable operating temperature of the system.
Because many systems need to be run at much lower temperatures than the waste heat of a power
system, these radiators are often very large and heavy. However, because of the pulsed nature of
this power system, the energy can be radiated over long periods of time when the cycle is not
running. This allows for the radiators to be much smaller, to the point that they can be
incorporated into the housing of the unit without the need for fins. Therefore, a thermal capacitor
can also be used to absorb the waste heat from the working fluid through the blowdown sequence
and then dissipate that stored thermal energy while the cycle recharges.

34

Many materials were considered for the absorber, such as lithium, beryllium, boron, and
molten salts. However, many molten salts had operating temperature ranges well above what
was required, and they would not effectively cool the exhaust of the cycle. Boron and beryllium
had acceptable specific heats at 1.03 kJ/kg-K and 1.83 kJ/kg-K, but beryllium is hazardous to
handle and can make manufacturing quite difficult. Boron remains a viable contender but its
thermal physical properties were not quite as impressive as lithium.

Lithium is an almost ideal candidate for waste heat absorption because of its high specific
heat capacity (C
p
= 3.58 kJ/kg-K). However, it has a melting temperature of only 453 K [17].
The turbine outlet temperature of the Brayton cycle is estimated to be about 796 K. This means
that the leading edge of the lithium absorber could likely melt. However, a similar housing being
employed for the thermal capacitor would be able to contain the molten lithium and maintain the
integrity of the flow channels. Calculations and models show that the absorber, even if
temporarily molten, would be able to absorb the amount of waste heat produced and radiate back
to starting temperatures over the course of 20 hours.

1D heat transfer approach was used to evaluate fluid temperature profile as it flows
through the lithium absorber. In this case surface temperature of the flow channel is not constant,
as the temperature of the absorber increases with time. This translates to fluid exit temperature
increasing with time. Number of flow channels and their diameter can be determined by studying
their effect on thermal conductance of the absorber. Parameter of interest here is absolute
thermal conductance per unit length, which is give by the following expression:

=
!
!
=
!!
!
!
=
!!"#$
!
= ℎ
!
!"
à Equation 14
Where, N is the number of channels.

It describes how many Watts of thermal power is transferred (per unit length) from fluid
to absorber for a given temperature difference between the two. Fixing mass flow rate makes h
dependent on viscosity (which is known) and channel diameter d only, so c can be plotted as N
vs. d, as shown in Figure 16 (next page). Decreasing the channel diameter is an effective way to
increase heat transfer, but there is a practical limit on how narrow the channels could be
manufactured.

35

Figure 16: Effect of number and diameter of flow channels on absorber thermal conductance.
Increasing the number of channels is effective only to a certain point. Going beyond that,
the flow eventually enters the laminar region, where heat transfer reduces drastically (flatlands in
the above figure). Selected parameters were d = 4 mm and N = 60. This puts the design safely to
turbulent flow region. Once N and d are fixed, length of the flow path can be studied. Figure 17
below shows fluid temperature as it flows through the absorber.

Figure 17: Variation of fluid temperature through the absorber.

36
After 2 meters fluid temperature is already close to absorber temperature, and at 3 m the
minimum achievable temperature is reached. For storing all waste heat accumulated during the
six minute blowdown, required lithium absorber mass was determined to be 200 kg. Dividing it
into six cylinders gives L = 1m and diameter D
ABS
= 0.28 m. 3 meter flow path is then achieved
by passing the fluid through three cylinders linked together. Figure 18 below illustrates one of
the cylinders.

Figure 18: One of the six lithium cylinders comprising the absorber.
Described heat transfer model assumes even temperature distribution throughout the
absorber, which in reality is not the case. It depends on the Biot number, which is defined as:

=
!"#$%&'("# !" !"#$%
!"#$%&'(#&) !" !"#$%
=
!!
!
à Equation 15

If Biot number is less than 0.1, heat transfer inside the solid is much faster than heat
transfer into the solid. Biot number is 1.23 and 0.96 for hydrogen and helium respectively, so in
reality there are both axial and radial thermal gradients inside the absorber material. Flow
channel surface temperature would be higher than predicted by the current 1D model, which
would reduce the heat transfer rate. A 3D finite element analysis would be necessary to obtain a
more accurate answer. It should be noted, however, that the selected 3 m flow path length
provides more heat transfer area than necessary.

Once the 6-minute cycle is over, the absorber must return to its initial temperature (250
K) by radiating heat to space. Assuming emissivity ε = 0.8 and that 75% of the absorber area
sees cold space, this is achieved in 12 hours, which is less than time needed to recharge the core.

37
IV. MISSION ARCHITECTURE
The above concept design was based around Enceladus Orbiter mission architecture. The
mission design incorporated an evaluation of the science objectives laid out in the decadal survey
for an Enceladus mission. An assessment of potential instruments that can meet those objectives,
which would fit within the payload space envelope, was also investigated. An in-depth trajectory
analysis and optimization was carried out to take advantage of the dual-mode propulsion system
scheme to enable the mission.
A. Design Approach

From a mission perspective the architecture design is scrutinized in three ways. Figure
19 below shows an illustration of the importance of the three ways a mission design is
scrutinized. These are in order of most critical (bull’s eye) to least critical.

Figure 19: Scrutinized hierarchy for the mission design
a) Can we get there? – This is the primary concern in the mission architecture design. This
requires to obtain the most optimized trajectory to Enceladus for the required payload mass,
power & propulsion system mass (Brayton Engine), communication system mass (Antenna
& Electronics) and the minimum propellant mass. A Matlab code was written to develop and
optimize a trajectory design to Enceladus. The trajectory design was segregated into two
different portions: (1) Earth – Saturn Cruise Phase & (2) Saturnian Moon Tour – Enceladus
Orbiter.

38
b) Can we talk to the Satellite? – This concern in the mission architecture design requires a
communication system which we be able to achieve a reasonable data rate and a decent
signal to noise ratio from Enceladus. Uplink data transmission from the Deep Space Network
(DSN) to the spacecraft at Enceladus is not a major issue, since high kW class of power can
be transmitted with from the DSN.
c) Can we survive there? – This concern requires the determination of shielding specifications
or the duration of the mission against harmful radiation environments in space. The mission
architecture incorporates the concept of phasing maneuvers to escape Earth’s sphere of
influence from Low Earth Orbit (LEO). Hence an estimation of the total radiation dosage in
the radiation environment of Earth caused due to Earth’s magnetosphere has to be predicted.
The high energetic trapped electrons and ions in Earth’s Van Allen Radiation Belt are the
most hazardous environment for electronics and can cause multiple Single Event Upsets
(SEUs). Also the radiation environment of Saturn has to be considered for making sure there
is enough shielding to counter during the primary scientific mission life at Enceladus.
Estimating the radiation dosage and the energetic charged particle fluxes along the entire
trajectory of the spacecraft will help optimize the trajectory to a threshold limit of dosage in a
reasonable allowed tolerance. This tolerance defines the shielding thickness, which drives the
cost and mass.
There are four types of missions that can be adopted for Enceladus. These are: (1)
Enceladus Flyby, (2) Enceladus Rendezvous, (3) Enceladus Bon Voyage and (4) Enceladus
Sample return. Figure 20 in the next page illustrates the different concepts / strategies of
Enceladus mission architectures that can be adopted for the demonstration of the dual-mode
radioisotope propulsion technology. The red boxes in Figure 20 represents the selected mission
architecture option that has been worked on, for demonstrating the application of the bi-modal
radioisotope power and propulsion system.

39

Figure 20: Mission Architecture Trade Tree [19]
B. Science Objectives

Apart from all the exquisite beauty of Saturn, some of the most interesting discoveries of
all were found in Saturn’s myriad moons. Saturn has 62 confirmed moons, which is the most of
any planet in the solar system. Figure 21 in the next page shows some of Saturn’s moons. Most
of these moons are small and less than 30 miles across. But the larger moons are some of the
most intriguing extra-terrestrial bodies in the solar system.
Enceladus
Flyby

Saturn
Orbiter
with

Enceladus
Flybys

(Example
-‐
Cassini)

Single
High
Speed

Flyby.

Through
the

Enceladus
plumes
at

the
south
polar

region

Enceladus

Rendezvous

(Orbiter
only)

Simple
Orbiter

Multiple

Engineering
and

Scientifc
Flybys
of

Rhea
and
Titan

Selected
Science

Goals
-‐
shorter

operations
at

Enceladus

High
Performance

Orbiter

Multiple
Dlybys
of

Rhea,
Mimas,

Iapetus,
Dione,

Tethys
and
Titan

All
primary
science

goals
-‐
longer

operations
at

Enceladus

A
penetrator
type

lander
onto
the

surface
of
Enceladus

Enceladus
Bon

Voyage

Titan
Orbiter
with

Enceladus
Dlybys

Titan
Orbiter
which

will
deploy
a
Titan

lander

After
deployment

of
lander,

continue
orbiting

Titan
to
collect

lander's
science

data
until
lander's

EOL.

After
Titan
lander's

science
mission
EOL

escape
Titan
to

perform
EOI
as
a

simple
orbiter

40

(a) (b) (c)
Figure 21: Saturn’s smaller moons – (a) Epimetheus [20], (b) Pandora [21] and (c)
Telesto [21].
For example Mimas (shown below in Figure 22) is recognizable from its massive crater
that spans more than 80 miles in diameter. The crater is the remnant of an impact so violent, it
nearly split Mimas into two. Iapetus is Saturn’s third largest moon and seems to have almost a
split personality, with one side a soft white in color like snow and the other side a dark and
tarnished surface. Running along the equator of this moon is a mountain ridge more than 800
miles long, 12 miles wide and reaching more than 42000 feet high (higher than the Himalayas).
The bizarre looking Hyperion was the first non-spherical moon to be found. Its irregular shape,
chaotic rotation and strange sponge like appearance remain unexplained.

(a) (b)

(c) (d)
Figure 22: Saturn’s larger moons – (a) Rhea [22], (b) Mimas [23], (c) Iapetus and (d) Hyperion
[24].

41
But among the smaller icy moons of Saturn, none has generated more excitement and
fascination than Enceladus. It is smaller than our own moon but is still one of the brightest
objects in the solar system. Its frozen surface reflects nearly 100 % of the sunlight that hits on it.
It was quite surprising for the science community when Cassini detected a hot zone at
Enceladus’ South Pole. Closer inspection revealed a very active surface geology, with cracks and
fissures continually forming and reforming in the icy crust.

Cassini made several very close flybys of Enceladus and scientists were astounded by the
discovery of huge plumes of water vapor and ice crystals continuously venting out into space
from this southern hot zone. It soon became clear that these geysers dubbed “cold faithful” where
actually the material source for Saturn’s immense yet diffused “E-ring”. But even more
significantly, they suggest that a liquid ocean warmed by volcanic activity may exist beneath the
frozen surface of Enceladus, making it one of the most promising candidates for harboring
microbial life in our solar system. Figure 23 (a) shows the icy moon Enceladus and (b) shows
Enceladus’ interaction with Saturn’s e-ring.

(a) (b)
Figure 23: (a) Saturn’s ICY moon – Enceladus & (b) Enceladus’ interaction with Saturn’s “e-
ring” [25]

Currently navigating around Saturn, Cassini is a very successful mission, which was built
in the legacy of past missions such as Voyager and Galileo. In time, Cassini will exhaust its
remaining fuel and when that time approaches, mission navigators have devised a plan that they
hope will thread Cassini at the small space between the inner most ring of Saturn and the planet
itself. Here Cassini will observe Saturn in unrivalled detail for 22 orbits, before gravity finally
draws the spacecraft down into the clouds of Saturn.

42
To build on Cassini’s revelations of Enceladus, an Enceladus orbiter mission architecture
was hence proposed that would utilize the dual mode propulsion system for power and
propulsion. The science objectives developed for this architecture are those based on NASA’s
Planetary Science Decadal Surveys. As the Decadal Survey alludes to, the South Pole plumes are
the most important in scientific interest. This is because it is believed that the plume may contain
the basic necessities for biotic material, including the elements H, C, N, O and possibly liquid
H
2
O. Figure 24 below shows the data from Cassini’s Composite Infrared Spectrometer (CIRS)
instrument. It shows the plumes in the South Polar Region are associated with elevated
temperatures. The understanding of the source of heat driving the plumes, their molecular
composition and the physical & temporal characteristics of the plume’s dynamics are the three
most essential scientific goals of their study.

Figure 24: Cassini’s CIRS instrument data of Enceladus South Pole [19].

The primary proposed science objectives of a mission architecture for Enceladus are as
follows: (1) Entering an orbit around Enceladus to map gravity and magnetic field, (2)
Measurements of the molecular composition of the plume’s macro particles, (3) Measurements
of the temporal and spatial variation of the plumes, (4) Slower flybys for plume sample &
surface mapping and (5) Potential sample collection and return for analysis [19]. Based on the
restriction in the space & mass envelope of the payload section and the mission architecture type,
a selected list of science objectives in Table 10 (next page) are listed from the total list of all
science objectives laid out in the decadal survey. This will be the list of science objectives, which
this proposed mission architecture design would work towards in achieving. All the science
objectives in Table 10 can be accomplished between five instruments. They are: (1) Medium
Angle Camera (MAC), (2) Thermal Imaging Radiometer (TIR), (3) Dust Analyzer, (4) Mass
Spectrometer (MS) and (5) Radio Science (RS).

43
Objective Investigation Instrumentation
Physical
conditions at the
plume source &
origin of south polar
surface features

1. Topography & thermal output
2. Plume vent shape
3. Surface strength & roughness
4. Particle size distribution and speed
5. Ice temperature
Medium Angle
Camera (MAC),
thermal imager, dust
analyzer, Mass
Spectrometer (MS)
Chemistry of the
plume source and
chemical clues to
Enceladus’ origin and
evolution.
1. Chemical inventory of the plume gas
2. Isotropic ratios
3. Isotopic and elemental analysis of plume
gases and dust grains
MS, dust analyzer
Presence of
biological activity
1. Organic molecules inventory
MAC, MS, dust
analyzer
Plume dynamics
and mass loss rate
1. Plume structure
2. Ejection rates
3. Particle size, density, composition and
velocity
MAC, MS, dust
analyzer
Internal structure
1. Static gravity
2. Potential Love numbers
3. Magnetic field
Radio science,
magnetometer,
imaging
Presence, physics
and chemistry of the
ocean
1. Magnetic induction
2. Plume chemistry
Radio science,
magnetometer, MS,
dust analyzer
Tidal dissipation
rates and mechanisms
1. Long wavelength global thermal emission
2. Bolometric albedos
MAC, thermal
imaging radiometer
Nature and origin
of geological features
and geologic history
1. Geology
2. Topography
3. Stratigraphy
MAC, radio science
Plasma and neutral
clouds
1. Spatial distribution and composition
2. Time variability of neutral clouds and its
correlation with plume activity
MS, MAC to
monitor plume
activity
E-ring
1. Variation in composition and relation to
Enceladus’ activity
Dust analyzer, MAC
for E-ring structure
Table 10: Science objectives, investigation and instrumentation [19]
The five science instruments, which will be used in
this mission, could be housed inside a 6U CubeSat envelope.
Figure 25 on the right shows an artistic rendering of the 6U
CubeSat casing, which could house the entire instrumentation
payload, including the supporting sensors and a power source.

Figure 25: Artistic rendering of the 6U toroidal cage.

44
V. CONCLUSION

A TPV energy conversion system was also considered as an alternative power generation
method to the Brayton cycle. Coupled with batteries or capacitors, a TPV system could provide
bursts of electrical power that could reach 25 kW
e
as well. However, based on calculations done
on off-the-shelf super capacitors the mass of the capacitor bank would be extremely high. A
TPV / capacitor system was evaluated using VHC 2R3 807 QG capacitors from VinaTech [18].
These capacitors require 2.3 V, have a mass of 0.94 kg and provide 800

F of capacitance each. In
order to provide the 9 x 10
6
J needed to reach 25 kW
e
the capacitor bank would have a mass
burden approaching 4,000 kg. This mass need of the capacitor bank greatly exceeds the mass
estimates of the Brayton engine, but also that of the entire propulsion system.

Furthermore, using TPV system as a secondary power source in conjunction to the
Brayton engine was also investigated. The exposed surfaces of the core are at much lower
temperatures than that of the thermal capacitor. In order for a TPV system to function, it needs
an exposed hot surface that can transfer energy through radiation. The exposure of the hot core
would require the removal of the insulation layers, and reduce the maximum operating
temperature of the core. This in turn, negates the thermal storage potential of the silicon thermal
capacitor, greatly affecting the thermal management of the overall system. Also, if the
photovoltaic panels were placed at any point along the path of the working fluid or propellant,
the stresses and temperatures they would encounter would quickly degrade them to the point of
rendering them nonfunctional.

The reliability of the components over the mission lifetime is always a big technical
challenge of using a Brayton engine. However, because the system is pulsed, the cycle operates
for a significantly shorter amount of time over the entire mission. The current design of the
Brayton system is to be pulsed once a day for 6 minutes. Over a 15-year mission, the cycle will
only operate for 547.5 hours. Additionally, due to the operational temperatures (T
in
> 1000 K),
the turbine can be at risk of thermal creep. However, the use of ceramic materials, such as Si
3
N
2
,
can drastically increase turbine lifetime and operational temperatures.

45
VI. REFERENCES

[1] Jerred, N. D., S. Cooley, R. C. O’Brien, and S. D. Howe., Proceedings of AIAA Space 2012
Conference, Pasadena, Paper 5152.
[2] Gaskell, David R, Introduction to the Thermodynamics of Materials, 4th edition, New York:
Taylor & Francis, 2003, Print.
[3] Morgan, S., B. Manning, N. Addanki, M. Trubilla, S. Howe and J. King., 10kW Radioisotope
Powered Pulsed Brayton Cycle for Space Applications, Proceedings of Nuclear and
Emerging Technologies for Space 2011, Albuquerque, Print. Paper 3303.
[4] Howe, S. D., R. C. O'Brien, R. M. Ambrosi, B. Gross, J. Katalenich, L. Sailer, M. McKay, J.
C. Bridges, and N. P. Bannister, The Mars Hopper: An Impulse-driven, Long-range, Long-
lived Mobile Platform Utilizing in Situ Martian Resources, Journal of Aerospace Engineering
Special Issue Paper (2010): 144-53. Print.
[5] Baum, E. M., H. D. Knox, T. R. Miller, Nuclide and Isotopes: Chart of Nuclides, 16th
Edition, Lockheed Martin (2002).
[6] O’Brien, R. C., R. M. Ambrosi, N. P. Bannister, S. D. Howe, H. V. Atkinson, Spark Plasma
Sintering of Simulated Radioisotope Materials in Tungsten Cermets, Journal of Nuclear
Materials, 393 (2009) 108 – 113.
[7] O’Brien, R. C., Radioisotope and Nuclear Technologies for Space Exploration, PhD Thesis,
University of Leicester, UK (2010).
[8] Kelley, K. K., The Specific Heats at Low Temperatures Of Crystalline Boric Oxide, Boron
Carbide And Silicon Carbide, Journal of the American Chemical Society, 63 (1941) 1137-9.
[9] Kantor, K., P. B. Krasovitskaya, R. M. Kisil, O. M. Fiz., Determining The Enthalpy And
Specific Heat Of Beryllium In The Range 600-2200, Phys. Metals and Metallog. 10 (6)
(1960) 42-4. Mcl-905/1, Ad-261792.
[10] Booker, J. Paine, R. M. Stonehouse, A. J. Wright, Investigation Of Intermetallic
Compounds For Very High Temperature Applications, Air Development Division (1961) 1-
133. Wadd Tr 60-889, Ad 265625.
[11] Pankratz, L. B. K. K. Kelley, Thermodynamic Data for Magnesium Oxide, U S Bur
Mines, Report 1-5 (1963), Bm-Ri-6295.

46
[12] Kandyba, K., V. V. Kantor, P. B. Krasovitskaya, R. M. Fomichev, E. N. Dokl,
Determination Of Enthalpy And Thermal Capacity Of Beryllium Oxide In The Temperature
Range From 1200 – 2820, Aec-Tr-4310, (1960) 1 – 4.
[13] Hedge, J. C., J. W. Kopec, C. Kostenko, J. I. Lang, Thermal Properties Of Refractory
Alloys, Aeronautical Systems Division, (1963) 1-128 (Asd-Tdr-63-597, Ad 424375).
[14] Metals Handbook, Vol.2 - Properties and Selection: Nonferrous Alloys and Special -
Purpose Materials, ASM International 10th Edition 1990.
[15] English, R., Technology for Brayton-Cycle Space Powerplants Using Solar and Nuclear
Energy, NASA Technical Paper 2558, 1986.
[16] MotoEnergy, ME1115, URL: http://www.motenergy.com/me1115motor.html [cited 13th
May 2014]
[17] NETZSH, Periodic Table of the Elements, NETZSCH-Geratebau, URL: www.netzsch-
thermal-analysis.com [cited 13th May 2014]
[18] Microtech Product details, PR05479, Ultra Capacitor, Part No: VHC 2R3 807 QG, URL:
www.microtech-hk.com/product_details.php?id=PR05749 [cited 13
th
May 2014]
[19] Spencer J., Niebur C., Mission Concept Study: Planetary Science Decadal Survey JPL
Rapid Mission Architecture (RMA) Enceladus Study Final Report, April 2010, URL:
http://sites.nationalacademies.org/cs/groups/ssbsite/documents/webpage/ssb_059319.pdf.
[20] Wikipedia, Epimetheus (moon), URL: http://en.wikipedia.org/wiki/Epimetheus_%28moon%29.
[21] Wikipedia, Pandora (moon), URL: http://en.wikipedia.org/wiki/Pandora_%28moon%29.
[22] Wikipedia, Telesto (moon), URL: http://en.wikipedia.org/wiki/Telesto_%28moon%29.
[23] Wikipedia, Rhea (moon), URL: http://en.wikipedia.org/wiki/Rhea_%28moon%29.
[24] Wikipedia, Mimas (moon), URL: http://en.wikipedia.org/wiki/Mimas_%28moon%29.
[25] Wikipedia, Iapetus (moon), URL: http://en.wikipedia.org/wiki/Iapetus_%28moon%29.
[26] Wikipedia, Hyperion (moon), URL: http://en.wikipedia.org/wiki/Hyperion_%28moon%29.
VII. BIBLOGRAPHY

[1] Jerred N., Howe T., Howe S., Rajguru A., Dual Mode Propulsion System Enabling CubeSat
Exploration of the Solar System, NASA Innovative Advanced Concepts Phase I: Final
Report, URL: https://www.nasa.gov/sites/default/files/files/Jerred_2013_PhI_DualModeProp
.pdf, Cited on [July, August and September 2015].

47
Chapter 2

Commercializing C-Class Asteroids in the main Asteroid Belt

I. ASTEROID COMMODITIES AND POTENTIAL MARKETS
A. Water

Water can be electrolyzed into LOX and LH
2
or it can directly be used as a propellant in
a nuclear propulsion system (Chapter 1). There is abundant evidence that water is present in
carbonaceous chondrites, which forms the C-Class asteroids in the main asteroid belt. There is a
strong belief that water is also available in the form of regolith ice and chemically bound
sources. Water is vital for manned space programs as it can be used for drinking water,
agriculture and radiation shielding. Approximately 90 % of this water can be carried from Earth
and recycled within a life support system [1]. But for a very long duration human presence in
missions beyond the asteroid belt, refueling water from the main asteroid belt for sustainability
can be an enabling aspect of such a mission. Hence the value of information for water maps on
the main asteroid belt can be a very lucrative market for futuristic space missions.
B. Aluminum, Iron, Nickel, Silicon and Titanium

Information on the availability and distribution of Al, Fe, Ni, Si and Ti can be vital for
commercial mining interests and the evolution of private deep-space companies.
C. Platinum Group Metals (PGMs)

PGMs can potentially be a valuable commodity at Earth. The PGMs consist of
Ruthenium (Ru), Rhodium (Rh), Palladium (Pd), Osmium (Os), Iridium (Ir) and Platinum (Pt).
Platinum’s non-allergenic and oxidation resistant property makes it ideal for jewelry. As an alloy
with Iridium, its brilliance is enhanced. Platinum is also used in catalytic reforming, which is a
process to upgrade the octane content of gasoline. Palladium, which is more expensive than gold,
has a wide variety of applications. It is used in coating multi-layer ceramic capacitors (MLCC),
which stores electric energy in the form of an electrostatic field. Palladium is also used in the

48
connectors of hybrid integrated circuit. Magnetic platinum-cobalt alloys constitute vital
components in computer hard disks. Rhodium is very useful as an electrical contact material due
to its lower electrical resistance. It is also highly resistant to corrosion [2].

Figure 26: Platinum (top left), Palladium (top right) and Rhodium (bottom). Platinum has high
melting point (1772
0
C) and is very stable at high temperatures. Palladium and Rhodium are both
a silvery white metal. Rhodium is used as electrodes for aircraft spark plugs [2].
However, the extraction and refining of PGMs involves very complicated metallurgical
processes on Earth. Hence, implementing an ISRU process for the extraction of PGMs in Space
poses a great challenge. In the future, if any game changing applications of PGMs creates a large
demand on Earth, there could be a possibility in formulating profit making business models of
PGMs mining on Asteroids. Planetary resources estimate that a single 30-meter long platinum-
rich asteroid could contain more than $25 billion worth of Platinum [3]. Hence mapping such
asteroids in the main asteroid belt could be a valuable information asset.

49
D. Regolith

During Mars Science Laboratory’s 253 days cruise to Mars, the Radiation Assessment
Detector (RAD) absorbed about 50 rem (Roentgen equivalent man) [1]. On average human
beings receive 0.1 rem in a year and one rem carries with it a 0.055 % chance of developing
cancer [4].

Hence 50 rem can be very damaging to a human mission. In the future if a human
mission to the outer planets has to be foreseen, then the strategic importance of a base station in
the main asteroid belt becomes crucial. To build such a habitat will be a very expensive process.
Hence an alternative solution to reduce the cost would be to employ ISRU of asteroid regolith
and attach the fabricated regolith tiles to the exterior of the habitat. This is a better option for
radiation shielding, as it does not consume the volume of the habitat unlike water. Also, regolith
will be excellent in protecting against Galactic Cosmic Rays (GCR), micrometeoroids and
enhancing the thermal stability of the habitat [1]. Thus collecting asteroid regolith samples and
studying the composition of this regolith material will have great value. Information on asteroid
regolith material composition could boost the 3D printing market, because the demand for
producing a space based 3D printer that can use regolith to build habitat structures will be high.

Figure 27: Artist’s impression of ISRU based robotic construction technologies using
asteroid regolith. This picture is just an illustration of a futuristic space based 3D printer using
asteroid regolith material to fabricate habitat structures [5].

50
II. WATER DETECTION

The Mars rover Curiosity used the Sample Analyzer for Mars (SAM) instrument in order
to detect water on the subsurface of Mars. It successfully detected a large amount of water, about
1 liter per every cubic foot [6]. This equates to about a 2% water composition of the soil.
Asteroids are suspected to have a higher concentration. It is predicted that an average hydrated
C-Class asteroid can contain 5 – 12 % of frozen water, with an additional 10 – 20 % trapped in
molecular bonds within various complex metal ions constituting the soil [7]. Table 11 below lists
expected elemental surface and subsurface composition of C-Class asteroids in the main asteroid
belt. The last two rows with bold & italic font at the bottom of the table are the water-bounded
molecules. This bounded water can be somewhat easily extracted, simply by heating the soil to
the right temperatures.
Element / Molecule Min % Max % Element / Molecule Min % Max %
Fe-Ni granular 5 35 Co 0.1 0.5
Iron-oxide (Magnetite, Silicates) 15 30 Na
2
O 0.3 0.9
Magnetite Traces 9.7 K
2
O 0.04 0.1
Silica (SiO
2
) 28 40 P
2
O
5
0.23 0.28
Magnesia (MgO) 20

25 Clay matrix - -
Aluminum (Al
2
O
3
) 2 3 Olivine 7.2 -
Calcia (CaO) 2 - Mg - Olivine with FeO - -
Iron Sulfides (FeS) 6 - Pyroxene - -
Water (Clay matrix) 10 - Troilite 2.1 -
Sulphur 1 5 Pyrrhotite 4.5 -
Carbon 1 5 Saponite - Serpentine - 71.5 71.5
Water (Epsomite) 5 15
Ferrihydrite 5 -

Sodium & Magnesium Salts 10 -
Table 11: Predicted surface composition of C-Class asteroids in the main asteroid belt.
A. Concept of using a neutron source for water detection

Neutrons can be used to accurately detect the relative abundance of water on a target
body. The likelihood of an incident neutron interacting with the nucleus of a water molecule
depends on the neutron cross section of the hydrogen atom in the water molecule. Since the
neutron cross section of hydrogen is a universal parameter, the post analysis of a neutron hitting
the nucleus of a water molecule is predictable.

51
The after effects of a neutron bombarding with the nucleus of another atom can result in
the following three outcomes: (1) Elastic and inelastic scattering, (2) neutron capture and (3)
fission. Elastic scattering is simply a neutron bumping into a nucleus, imparting some of its
energy to that nucleus, and continuing on a new path with no absorption. Inelastic scattering
involves the neutron being absorbed by the atom and then re-emitted, along with a photon.
Neutron capture involves the neutron being absorbed permanently. This is typically followed by
a rearrangement of the atom’s nucleus to a lower energy state and the emission of a photon.
Fission occurs when a nucleus is large and unstable, such that the energy of the bombarding
neutron is enough to break the large nucleus apart into two smaller molecules [8].

Elastic scattering is the primary after effect of a neutron collision with a hydrogen atom.
This is due to the simplicity of the hydrogen atom, which is made up of a proton and an electron.
Having the nucleus composed of just a proton does not allow for much room to absorb or capture
a neutron when they interact. The probability of a neutron to collide with a hydrogen atom is
shown in Figure 28 below. In the plot below, the small region in the low energy range where the
cross section is decreasing, represents the 1/v dependency. “v” represents neutron energy. The
cross-section of a hydrogen atom is the probability for a neutron to collide with a hydrogen atom.

Figure 28: The probability of interaction of a neutron with another atom’s nucleus based on
neutron energy & cross section [9].

52

As the energy of the neutron increases, the probability of interaction decreases. Most
elements show this relation at low neutron energies. However, hydrogen cross-section is unique
due to the leveling off of this dependency as shown in Figure 28 above. Most elements will
continue this decrease until a region of resonance, which represents the absorption energy of the
neutron, is reached. The leveling of the elastic interaction in Figure 28 shows a constant
interaction probability even when the neutron energy is increased. This serves as an ideal
situation for our application. As stated earlier, the lack of the resonance region in Figure 28
shows that a hydrogen nucleus is unlikely to absorb a neutron.

It is required not to capture or absorb neutrons, as this would decrease the number of
neutrons that could possibly be detected. Having a large neutron culture not only increases the
statistical data but also allows for a greater detection distance between the detector and the target
body. An added benefit of neutron-hydrogen elastic scattering is their equal size. When the two
collide, much of the neutron energy is imparted to the hydrogen atom. This occurrence is known
as neutron moderation. Hydrogen will effectively keep the neutron culture at a maximum, while
shifting the neutron energy spectrum to a lower energy. By knowing the energy spectrum of the
emitted neutrons, it is possible to calculate or simulate the shift in the energy spectrum based on
the amount of hydrogen present in the target body.

Water, having two hydrogen atoms for one oxygen atom, is known to be an effective
moderator in reducing neutron energy. This is why water is used around nuclear reactors. Figure
29 in the next page shows an example of neutrons interacting with a surface containing various
percentages of water. The primary constraint on the efficiency of detecting water-using neutrons
is the neutron cultures that can be send in the direction of the target body. Neutron production is
very easy in comparison to directing them towards a desired target. This is due to the fact that
neutrons are neutral particles with minimal magnetic moment and hence cannot be influenced by
an electric or magnetic field. The plot in Figure 29 shows the neutron energy spectrum from
emitted neutrons being elastically scattered back towards the detector. The surface reflecting the
neutrons had the water concentration varied in each of the three simulations as shown in Figure
29. The x-axis of the plot in Figure 29 is the neutron energy in MeV.

53

Figure 29: Neutron spectra returned as a function of water content [10].
B. Neutron Generator Concepts

There are two concepts that were considered. The first concept was re-evaluated from Dr.
Steven Howe’s NASA Innovative Advanced Concepts (NIAC) phase 1 progress report, about a
Micro Asteroid Prospector Powered by Energetic Radioisotopes (MAPPER). This concept
involved using a neutron generator assembly, where a “gun” will shoot a lithium ion through an
electrostatic acceleration column directed at a hydrogen gas target. When the lithium ion collides
with the hydrogen, the two elements will fuse together and create a beryllium atom while
emitting a neutron at 2 MeV. The neutron will pass through the surrounding material unaffected.
The vast majority of the lithium will pass through into the electrostatic deceleration column,
decreasing its energy by 0.2 MeV, and into a lithium absorber. This absorber will recover most
of the energy required to accelerate the lithium ions. The schematic of the neutron generator
assembly is shown in Figure 30 below [10].

Figure 30: Mapper spacecraft after the deployment of lithium acceleration and deceleration
electrostatic columns [10].

54

There are a few problems with this concept. Creating a large concentration of neutrons in
this process is unlikely. As only one neutron per lithium-hydrogen interaction is obtained, and
getting the emitted neutron to go in a single predictable direction is very difficult.

An alternative second concept was to use isotopes that can fission spontaneously,
releasing multiple neutrons per spontaneous fission as shown in Figure 31. Three requirements
guided the selection of the material. These requirements were: (1) an isotope that is readily
available for use, (2) an isotope that releases the greatest amount of neutrons per fission and (3)
an isotope that has a high probability of decaying through spontaneous fission. Out of the usable
isotopes
250
Cm and
252
Cf (Californium – 252) have a higher spontaneous fission probability and a
higher neutron per decay count. However, the availability of
250
Cm is limited and
252
Cf has a
short life of 2.6 years. Hence
244
Cm was chosen as a suitable neutron generator source and was
used in all MCNP (Monte Carlo N – Particle Transport Code) simulations. The average neutron
yield per fission of
244
Cm is approximately 2.69 [11]. For the MCNP simulations the
radioisotope source composition used was 50 %
244
Cm, 25 % Tungsten and 25 % Rhenium.

Figure 31: Spontaneous fission of a large unstable molecule [12].
In order to check, whether the neutron source has been properly simulated or not in
MCNP, the neutron energy spectrum for the spontaneous fission of
244
Cm was found and plotted
in Figure 32. The neutrons released from the neutron source generated in MCNP, were tallied
and plotted in Figure 33 for comparison to plot in Figure 32.

55

Figure 32: The actual energy spectrum of neutrons from spontaneous fission of
244
Cm [13].

Figure 33: Energy spectrum of neutrons from the spontaneous fission of the generated
244
Cm
source in MCNP.
Although the y-axis of the two graphs has slightly different units, it can be seen from the
shape of the graphs that the energy spectrum is well represented in MCNP. It is the shape of the
graph that primarily matters. Additional information about
244
Cm that was used to generate the
MCNP graph is given in Table 34 in the appendices.
0 2 4 6 8 10 12 14
0
0.5
1
1.5
2
2.5
3
3.5
x 10
5
Energy (MeV)
neutron count (neutrons / sec.Watt)
Cm −244 Spontaneous Neutron Energy
0 2 4 6 8 10 12 14
0
0.5
1
1.5
2
2.5
3
3.5
4
x 10
10
Energy (MeV)
neutron count (neutrons / second)
Cm −244 Neutron Source in MCNP

56
C. Radioisotope Core of a Radioisotope Thermal Rocket as a Neutron Source Design

Two different design concepts were assessed that were capable of producing a directed
neutron beam towards a target body. The first idea was to use the radioisotope core of a
Radioisotope Thermal Rocket (discussed in chapter 3) as a neutron source. The second concept
was to employ a neutron gun as a separate system on the spacecraft. Again, for both the concepts
the fuel element chosen was
244
Cm, because of its excellent neutron emission property due to
spontaneous fission. Also for both the concepts MCNP was used to evaluate the flow of reflected
neutrons across a simulated hydrated C-Class asteroid detector surface.

In order to direct the flow of the neutrons towards the target body, shielding was used
around the neutrons. An ideal shielding material will have an elastic interaction with the neutrons
similar to the way hydrogen interacts. However, a heavier element has to be used in order to
minimize the exchange of the neutron into the shielding atoms. When the neutron has an elastic
scattering interaction with a heavy element, the neutron will simply bounce off of the nucleus
and retain most of its energy.

Figure 34 shows the neutron cross section for Beryllium. The plot also shows a similar
pattern as to that of Hydrogen in Figure 28 but with a lower cross section. There is a constant
elastic cross section at lower neutron energies, but Beryllium shows a small resonance region at
the higher neutron energy levels. The cross section of beryllium is the probability that a neutron
will interact. The first task to simulate the nuclear core as a neutron source was to generate 6
different core configurations in MCNP, each having a different number of fuel rods of various
radii. Other than the rod
configuration, the rest of the
system is exactly the same.

Figure 34: Shows the
interaction probability of a
neutron with a beryllium
nucleus based on neutron
energy dominated by elastic
scattering [9].

57

Figure 35: MCNP plot of a radioisotope thermal rocket using tungsten encapsulated
244
CmO
2

fuel rods embedded inside Beryllium acting as a thermal capacitor and neutron reflector.

Figure 35 shows the MCNP configuration of the radioisotope core, Beryllium cladded
nozzle, neutron detector and the simulated hydrated C-Class asteroid surface. The Beryllium core
(rendered as blue) surrounds all of the fuel rods (rendered as pink). A Beryllium cap is then
placed on top of the core to keep neutrons from escaping in that direction. Now the neutrons can
only escape through the bottom of the core through the nozzle. The design of the nozzle was
based on the expansion ratio required for thermal propulsion. The shape of the nozzle hence does
not serve as an ideal muzzle for a neutron beam. With the nozzle constricting and then widening,
the emitted neutrons will be allowed to disperse rather quickly from the exit.

The interior of the nozzle material was Beryllium cladded to minimize the escape of
neutrons through the nozzle wall. The neutron detector wraps around the exit of the nozzle and is
made up of pressurized hydrogen. The asteroid surface (yellow) is composed of a 10 cm thick
top surface layer, which is made up of 8% water, 46% Calcium Sulfate and 46% Gypsum.

58

MCNP was used in simulating the generation of neutrons from
244
Cm due to spontaneous
fission. The Beryllium shielding helped most of the neutrons to escape through the nozzle of the
system. Once emitted from the nozzle, the neutrons collide and interact with the asteroid surface.
The hydrogen in the asteroid thermalizes the neutron energy spectrum and reflects the neutrons
back towards the core. As the distance between the core and asteroid increased, more of the
neutrons were spread out and missed the detector surface. This limited the operational distance
between the neutron detector and the asteroid surface. In this simulation, the entire neutron
spectrum was tallied to find the maximum distance at which a neutron signal can be detected.

The goal of this simulation was to determine which core configuration allows the furthest
detection distance. Figure 36 & Figure 37 shows how the fuel rods are arranged in the
Beryllium core, the diameter of the fuel rods and the number of fuel rods. In core configuration
5, the diameter of the 17 small fuel rods is 0.263 cm. The diameter of the 6 medium fuel rods is
0.394 cm. The diameter of the 6 large fuel rods is 0.526 cm. Figure 104 to Figure 109 in the
appendices show the number of neutrons per energy bin that are reflected off from the asteroid
surface and registered into the detector surface for each core configuration. It also shows the
maximum detection distance as well.

Core configuration 1 Core Configuration 2 Core Configuration 3

Figure 36:
244
Cm with Beryllium encapsulated core configurations 1 – 3 designed in MCNP.

59
Core configuration 4 Core Configuration 5 Core Configuration 6

Figure 37:
244
Cm with Beryllium encapsulated core configurations 4 – 6 designed in MCNP.

There was a slight anomaly with the output data for determining the maximum distance
of detection in order to decide the best configuration. When comparing the minimal detection
distance, it can be seen that the minimum neutron count varies between each plot as shown from
Figure 104 to Figure 109 in the appendices. Ideally, these simulations could be carried out up to
a distance where the neutron count falls below 25 neutrons / second, which is the minimum
detection limit for most neutron detectors. The distance between the core and the asteroid were
increased as far as possible until MCNP no longer gave a signature for more than one energy bin.
The reason why this MCNP output signature goes to zero is because MCNP normalizes the
output neutron count to the single starting source particle. The number is then manually scaled
up depending on the amount of fuel in the system and the size of the detector, which limits the
ability to detect small counts of neutrons. However, it is assumed that this is giving correct,
usable data as the neutron count is decreasing with the increasing distance up to a point where
neutrons are no longer detected as expected.

Although this is less than ideal, some conclusions can still be drawn from the data. It
appeared that the maximum distance that will fall below the 25 neutrons / second count was
approximately 15 meters. It could be safely assumed that core 1 is the worst choice out of all the
core configurations, as it drops below a 1000 neutron/sec count at 4-5 meters. The other 5 cores
have very similar output energies. In order to determine which core configuration is most likely
the best choice, the 7-meter distance was evaluated. Also, only the energy bin from 0 – 0.4 MeV

60
was examined since this contained the thermal neutron energies used to detect water. Table 12
gives the neutron count for last 5 configurations in the lowest energy bin at a 7-meter distance.

Distance Core 2 Core 3 Core 4 Core 5 Core 6
7 meters 1.26 x 10
4
1.15 x 10
4
1.40 x 10
4
7.40 x 10
3
1.08 x 10
4

8 meters

7.02 x 10
3
8.27 x 10
3
5.48 x 10
3
7.40 x 10
3

9 meters

5.84 x 10
3
3.67 x 10
3
4.97 x 10
3

10 meters

3.04 x 10
3

2
nd
Best

4
th
Best Best Worst 3
rd
Best
Table 12: Neutron Counts in the Energy Bin 0 – 0.4 MeV.
It can be seen that core configuration 4 has the largest count for all heights. Hence, this is
the suggested configuration to use for neutron detection. Core 2 has the second highest count at a
7-meter distance, but did not show any counts at further distances. From this study it can be
concluded that the minimum distance required for detecting water on the asteroid surface is on
the orders of a few meters. This is still very impractical for a spacecraft instrument and an
alternative design (Neutron Gun – see below) was investigated where the neutron flux beam can
be more directional.
D. Neutron Gun as a Neutron Source Design

A MCNP model of the Neutron Gun design, which was analyzed, is shown in Figure 38.
The MCNP simulation helped to determine the minimum amount of Beryllium required while
keeping the neutron culture at a maximum. However, a larger overestimated amount than the
minimum amount was chosen to ensure no neutrons would escape. The output data is shown in
Figure 39. Straightaway, it can be seen from the output data that the neutron gun has a much
greater potential at detecting neutrons from a further distance away than any of the radioisotope
core configurations mentioned earlier. Again, a nice trend of a decreasing neutron count can be
seen relative to the distance of the detector. Although MCNP output was limited to 40 meters, it
can be assumed that the detection distance limit is beyond 40 meters since there is a one million-
neutron count at that distance. If the trend of the graph continues, it could be roughly
approximated that a distance of just under 100 meters could possibly give a detectable neutron
count, but this should be investigated in an actual experiment.

61

Figure 38: Neutron Gun Design.

Figure 39: The detected reflected neutron spectrum from the neutron gun with varying distance
from the surface of the asteroid.

This reason for the large difference in counts for the neutron gun versus the radioisotope
thermal rocket core is the amount of Curium used, the location of the Curium and the design of
the nozzle allowing the neutrons to disperse. When the volume of the Curium used in the neutron
0 2 4 6 8 10 12
4
5
6
7
8
9
10
Energy (MeV)
neutron count (neutrons / second) in logarithmic scale
Neutron Count vs Neutron Energy at Varying Detector Distance

2 meters
4 meters
8 meters
12 meters
20 meters
28 meters
40 meters

62
gun is changed to match the volume used in the RTR cores, the neutron count at 40 meters is still
100K less than 1 million, which indicates that the design of the nozzle is a critical governing
factor. Knowing the most effective detection system now, a few more simulations were
performed in MCNP. The variation in the count of thermal neutrons versus water content was
investigated. For this simulation the layer of ice over the asteroid surface was removed and the
concentration of water in the asteroid surface were varied to different values for each simulation.
Figure 40 shows the MCNP results of thermal neutron spectrum versus water content.

Figure 40: Increasing neutron count with increasing water content in the asteroid composition.

There is a very nice trend of an increasing neutron count versus the amount of water
contained in the asteroid. The amount of reflected neutrons in the lower energy region is
sensitive to the amount of water. However, the energy above 0.06 MeV begins to have an
opposite trend. The 100% water line drops below the other lines at this point. Thus, when
attempting to detect water, it should not be necessary to examine the energy ranges above the
0.06 MeV, however this should be experimentally proven.

The last simulation performed was to look into the way the neutron count changes with
the diameter of the asteroid. The two plots in Figure 41 (next page) shows the neutron count for
different asteroid diameters at 5 meters distance (top plot) and 10 meters distance (bottom plot)
respectively.
−4 −3.5 −3 −2.5 −2 −1.5 −1
4.5
5
5.5
6
6.5
7
7.5
8
8.5
9
Energy (MeV) in logarithmic scale
neutron count (neutrons / second) in logarithmic scale
Neutron Counts vs Neutron Energy With Varying Water %

0 %
2 %
4 %
6 %
8 %
10 %
12 %
14 %
16 %
100 %

63

Figure 41: (Top plot) 5 meters distance away from asteroid.
(Bottom plot) 10 meters distance away from asteroid.

Figure 41 shows larger neutron detection for a larger sized asteroid at the same distance
away. The 5 meters distance plot quickly increases the neutron detection rate with asteroid size.
For this distance an asteroid greater than 2 meters will give similar results to any larger asteroid.
0 1 2 3 4 5 6
5.5
6
6.5
7
7.5
8
8.5
9
Energy (MeV)
neutron count (neutrons / second) in logarithmic scale
Neutron Count With Varying Asteroid Diameter (5 m Distance)

1 meter
2 meters
5 meters
10 meters
50 meters
0 0.5 1 1.5 2 2.5 3
5
5.5
6
6.5
7
7.5
8
8.5
Energy (MeV)
neutron count (neutrons / second) in logarithmic scale
Neutron Count With Varying Asteroid Diameter (10 m Distance)

2 meters
3 meters
5 meters
10 meters
20 meters
50 meters

64
The bottom plot for 10 meters distance shows a slower increase of neutron detection rate as the
asteroid size increases. Also, it takes a diameter of 20 meters to get to the point of approximately
a maximum neutron return. It can also be seen through comparison of both the plots that the
closer detection range has a higher neutron detection rate than the longer range. These were all
expected outcomes and have now been confirmed. This simulation can be repeated with
increased detector distances to find how large an asteroid has to be at further distances.
E. Neutron Detector

The main concept of a neutron detector is dependent on the way the neutron interacts
with matter. Since neutrons are neutral particles, it is impossible to directly measure their energy.
However, neutrons will interact with various isotopes in different ways. Initially, the goal is to
use isotopes that will elastically scatter the neutrons in order to keep the neutron culture at a
maximum. Typically when a neutron is absorbed by a nucleus the atom is excited to an unstable
state. In order to return to a more stable state, the excited nucleus will undergo some sort of
radiative decay. This is typically in the form of alpha decay (emitting a helium atom without the
electrons), beta decay (emitting an electron and an electron antineutrino), or gamma decay
(emitting high energy) out of the nucleus. Neutron detectors will take advantage of these
reactions and indirectly detect neutrons by detecting the energy of the emitted decay product.

The easiest type of decay to detect is the alpha particle as it is large (2 protons and 2
neutrons) and has a positive charge associated with it. An effective detector is a one that uses an
isotope, which has a high probability of capturing a neutron and also alpha decay. Currently, the
most common type of neutron detector is the scintillation detector. A typical Scintillation
detector uses Helium, Lithium or Boron to interact with the neutron. The neutron cross-sections
for these elements are given in Figure 42. This figure shows that helium has the highest
probability of interacting with a neutron, followed by boron and lithium. The probability
interactions plotted in Figure 42 are for the following nuclear interaction equations, which are
the most probable interaction for each element:
+
!
→
!
+
!
+

+
!
→
!
+
!
+

+
!"
→
!∗
+
!
→
!
+
!
+

65

Figure 42: The primary neutron interactions with helium, lithium and boron [9].

Each detector type has its own benefits and drawbacks. In either of the cases, the way the
scintillator works is basically the same, and is summarized in Figure 43. Detectable radiation
enters Sodium-Iodide crystal, producing light. The light is enhanced in photomultiplier tube and
detected by a measuring device.

Figure 43: Scintillation detector. [14]

66

Figure 43 shows the emitted radiation from the neutron interaction entering a Sodium-
Iodide crystal, which excites the atoms of the Sodium-Iodide crystal. The excited atoms emit
light in order to relax back to a lower energy state [15]. The light emitted is sent into a
photomultiplier tube, which enhances the signal through the photoelectric effect. This is
continuously done until the signal reaches an anode, creating an electric signal. The strength of
this electric signal can then be related back to the initial energy of the neutron. Figure 44 shows
a typical scintillation detector. In order to maximize the detection area, an array of these
detectors could potentially be set up in a circular pattern, representing the detection disk that was
simulated in MCNP.

Figure 44: A typical neutron scintillation detector [10].

i. Helium-‐3
Scintillation
Detector

A helium-3 scintillation detector is predominately used to detect moisture content in soil.
The primary benefit of using helium-3 is its sensitivity to thermal neutrons, making it well suited
for measuring substrates high in hydrogen. Also another critical benefit of using helium-3 is
having a byproduct of hydrogen. Since hydrogen is good at scattering neutrons, it gives the
detector a higher efficiency, as the neutrons are less likely to escape without being detected. A
secondary benefit is the voltage required to power the detector is low, relative to a boron
scintillation detector. From Figure 42 it can be seen that helium-3 has the highest probability of
undergoing the desired neutron interaction, especially at lower neutron energies, which are the
energies of most interest in detecting water [16]. The only drawback is that there is a short
supply of helium-3 and it would soon be difficult to get hold of.

67
ii. Lithium-‐6
Scintillation
Detector

Lithium 6 would be another possible element that can be used for detection. But, from
Figure 42 it can be seen that it has the lowest probability of undergoing the desired interaction
with a neutron when compared to helium-3 and boron-10. The only reason to use lithium-6
detector would be in the case of not being able to get hold of helium-3 and boron-10 detectors.

iii. Boron-‐10
Scintillation
Detector

The boron-10 scintillation detector actually uses Boron Trifluoride (BF
3
) as the material.
The boron in BF
3
undergoes the neutron interaction. Although BF
3
counters require more voltage
and are less sensitive than the helium-3 detector, the decay of boron-10 gives this detector a
longer life span. Once the boron-10 absorbs a neutron and undergoes alpha decay, it becomes a
lithium atom. This is lithium-7 instead of lithium-6, but is still fairly effective at absorbing a
neutron and emitting radiation. It is not effective enough to use lithium-7 just by itself but it will
allow for an extended use of the BF
3
detector.
F. Ground Penetrating Radar

Ground Penetrating Radar (GPR) can be used to unveil subsurface features such as the
composition, structure and depth using low frequency electromagnetic (EM) waves. Subsurface
features of an asteroid are composed of various materials that have different electric and
magnetic permittivity, which may be structured randomly in layers. Different pockets of material
will respond to EM waves accordingly. This technique has been proven and widely used to
interface materials beneath the surface (TRL 9). This technology has been used in MARSIS on
board the Mars Express Orbiter that is orbiting Mars at 250 – 300 km. The instrument MARSIS
uses the ground penetrating radar technique to locate water or ice below the surface and the
anticipated depth for observation is on the order of a few kilometers. The echoed signal would
include information on the electrical properties of the reflected surfaces, which is related to the
composition of that surface, and the depth of that surface. The power of the bounced radar EM
wave can be expressed from the following expression:

!
=
!
+2
!"#
+10
!"
!
!
!!
!
+
!"
+
!
à Equation 16

68
A typical hydrated C-Class asteroid surface can be comprised of the following materials
listed in Table 24. Figure 45 below shows the system architecture of the ground penetrating
radar.
Asteroid Soil Material Chemical Formula Composition (%) Relative Permittivity
Epsomite MgSO
4
·7H
2
O 5 – 15 0.23 [17]
Sodium Thiosulphate
+ Borax
Na
2
S
2
O
3
+
(Na
2
B
4
O
7
.10H
2
O)
10 6.55 [18]
Ferrihydrite (Fe
3+
)
2
O
3
.0.5H
2
O 5 30
Water / Ice H
2
O 10 80.4 / 3.15
Andesite (Na,Ca)Al
1-2
Si
3-2
O
8
TBD 5.83
Table 13: Hydrated C-Class asteroids soil composition and their relative permittivities.

Figure 45: System Architecture

G. Conclusion

In this chapter, various detection methods of water on asteroids were investigated. From
the study, it is concluded that the use of a ground penetrating radar to detect water on asteroids is
much more plausible from a far distance than any of the nuclear techniques discussed earlier.
However data processing could be complex and there could be calibrating issues. If a specific
asteroid target is identified for a human mission and is considered for retrieval and mining, it
won’t be a bad idea to use the neutron gun system coupled with a helium-3 scintillation detector.
Within a minimum distance of 90 meters, depending on the size of the asteroid, the neutron gun
can be used to determine the depth and concentration of water at specific locations. This will
allow locating the best spots for water harvesting and a landing spot for the humans to setup
machinery.

69
III. WATER EXTRACTION
Identifying the presence of water and determining its concentration will be a great source
of information for Asteroid Mining missions to harvest water. However mining asteroids for
water has immense engineering challenges that have to be worked on. This section of the chapter
provides an insight to such challenges and potential solutions.

Extraction of free water ice can be relatively simple and easy as it can be sublimed and
captured on a cold surface. However water extraction from hydrous minerals can be a daunting
task as it might require more heat and a complex ISRU system to tackle potential subsidiary
reactions between liquid, vapor, or gaseous water from the parent material. Hydrated
Carbonaceous Chondrites on asteroids can be Gypsum (CaSO
4
.H
2
O). To separate the bounded
water from Gypsum, it has to be heated to approximately 200
0
C under normal atmospheric
conditions [1]. The process of removing water from whatever molecule it is bounded to will be a
simpler process on a planet, such as Mars, due to its size and gravity. The extraction process
becomes complicated on asteroids where there is no gravity and no atmospheric pressure. Figure
46 below shows the phase chart of water based on temperature and pressure. The pressure on the
surface of Mars is 0.087 atm. From the figure below it can be seen that if water is extracted from
the Martian subsurface during Martian daytime, then it will be in the liquid form.

Figure 46: Phase chart of water depending on the temperature and pressure of the surroundings.

70
But, water located on most moons and asteroids, however, will be under a pressure less
than 0.006 atm. From Figure 46, it can be seen that water will not be available in liquid state
under 0.006 atm regardless of the temperature. Thus operating in the vacuum of space will be
very difficult. In general, the extraction process can be categorized into three levels of difficulty.

The easiest water for extraction is pure water that is unbounded to any other minerals.
Whether in liquid, ice, or vapor form, this will be the easiest water to extract from mars, moons
or asteroids. The next more difficult extraction process of water are from molecules, such as
clays or salts, that have bonded with water through Van Der Waals forces. This is more difficult
because the bond between the water molecules and the clay or salt molecule is typically stronger
than the bond between two water molecules. However, most of the water molecules in this
configuration can be removed through direct heating. The most difficult water to remove will be
the water that has been chemically altered or trapped inside of a molecule when it bonds or forms
another molecule. This can result in the generation of hydroxyl molecules.

A. Water Extraction and Collection Concepts

A couple of processes for extracting water off of an asteroid have been conceptually
considered. The first process involves the direct heating of the surface of an asteroid using
microwaves. Microwaves at the right wavelength interact with water, causing the water
molecules to vibrate. This vibration of the water molecule will heat up the molecules nearby.
Once the temperature of the asteroid surface has reached 0.01
0
C, water will begin to sublimate.
Once this happens, the water vapor will begin to migrate and expand in all directions. The
spacecraft will have to collect this vapor as it expands away from the asteroid. Operating at the
vacuum of space, it would be impossible to use a suction vacuum to intake the water vapor.
However a metal collector can be used to collect the water vapor. The water vapor leaving the
asteroid will collide with the metal collector, which will be at a lower temperature than 0.01
0
C in
order to cause the water to re-solidify on the surface.

71
One issue with this design is the low amount of water vapor that will be collected. When
the ice is heated, it will sublimate and spread out in all directions. If the collector of the
spacecraft is not directly against the surface of the asteroid then some of the water will escape
and be lost to space. Another problem is that, once the vapor has been collected on the collector,
the water must be transported from the surface of the collector and into a containment vessel.
This would require some sort of system to scrape the ice off of the collector, gathering it inside
of a vessel that can close, re-heat and encapsulate the water.

Only the frozen water on the surface of the asteroid will be expected to expand
specifically in the direction of the spacecraft. This is due to the density of the asteroid versus the
emptiness of space. When the water is heated it will vibrate off of the molecules surrounding it
and travel in the direction of least resistance, which is away from the asteroid. In this concept, a
fraction of water will be able to escape in the direction of the metal collector on the spacecraft.
This severely limits the amount of water that the spacecraft can collect at any given location on
the asteroid. However, it will be ideal to collect subsurface soil in a chamber and then microwave
it. The soil sample collected from the asteroid surface containing water molecules will be
surrounded by matter in all directions. This eliminates a path of least resistance and will cause
the water molecules to slowly expand in all directions. This idea is shown in Figure 47 below.

Figure 47: Illustration showing most of the water molecules escaping near the surface and the
rest of the water molecules spreading out in all directions [19].

72
Another idea is to use drilling equipment to collect the frozen surface water and the water
trapped in the soil. This process will use drill bits connected to each landing leg of the spacecraft.
The drill bits will drill into the surface and remove the water and soil samples. These samples
will be pulled, with the drill bit, into a containment vessel. This is where the heating of the water
will take place. One way of heating the water inside the container is using microwaves again.

However, using microwaves inside of a container with a metal drill bit could lead to a
charge build up on the drill bit. If the charge were to discharge on a nearby surface it is possible
that some of the electronic systems could get damaged. An alternative way to heat the water
would be using electrical heating, similar to that of a stovetop. This option would not be too
difficult to implement and the water could be heated up to the desired temperature rather quickly.
Simply by lining the drill bit with a high resistance metallic coil would be enough to allow for
heating. The primary drawback to using the drilling concept is the total weight of the system.
Typically, drilling systems are quite heavy, which makes them a costlier option.
B. Drilling Equipment

The most important part of the drilling equipment is the drill bit. There are multiple
possibilities on the design of the drill bit, but the two ideas thought to be the most suitable are
outlined. The first drill bit type is to use an auger type of drill piece, which will collect soil along
the outside helical threads of the bit and pull the soil into a housing unit. The second type is to
use a hollow drill piece, which will collect the soil sample on the inside of the bit. These two
designs are shown in Figure 48.

Figure 48: (Left) Auger drill bit [20] and (Right) Internal sample collecting drill bit [21].

73
i. Auger
Drill
Bit

The determining factor in the type of drill bit to be used will depend on the weight of the
two types. The drill bit that weighs the least while collecting the same amount of soil will be
selected. The Auger drill bit has a thick metal pole going down the middle of it with large
threading wrapped around. The threading will be the collecting plane for this bit. The size of the
threading will be calculated based on a 1000 cm
3
removal volume. For a six legged asteroid
lander spacecraft, each leg will have to drill and collect a soil sample of approximately 166 cm
3

per drill. However a total extraction volume of 188 cm
3
was chosen to give a contingency of just
over 10 % to help ensure enough soil is removed per drill. Table 14 below shows the possible
dimensions and weight of the total auger style drill bits, along with the casing for the drill bits.

Auger Drill Bit Auger Drill Casing
Parameter Quantity Parameter Quantity
Thread Layers 6 Number of Housing Units 6
Inside Radius 2 cm Radius 3.5 cm
Outside Radius 3 cm Length 13 cm
Collected Soil Depth 2

cm

Width 1 cm
Volume of Center Pole 188 cm
3
Volume of Housing 327 cm
3

Thread Thicknes 0.1 cm Volume of Bottom Panel 64 cm
3

Threading Volume 9.4 cm
3

Total Volume 198 cm
3
Total Volume 390 cm
3

Density of Aluminum 2.7 g/cm
3
Density of Aluminum 2.7 g/cm
3

Number of legs 6

Number of legs 6
Weight of all 6 Drill Bits 3.21 kg Weight of Casing 6.32 kg
Table 14: Dimensions and total weight of 6 Auger drill bits with housing.
The total weight of drill bit and casing combined = 9.53 kg. Figure 49 shows the possible
Auger design. The thread thickness is 2 mm. The housing for the drill bit is required in order to
create an enclosed area for water such that it does not escape when heated. As can be seen from
Table 14, the weight of the enclosure far outweighs the weight of the drill bits themselves. The
dimensions of the Auger drill bit casing are given in Figure 50.

74

Figure 49: The dimensions of the described Auger drill bit (not to scale).

Figure 50: The dimensions of the required housing for the Auger drill bit (not to scale).

ii. Hollow
Drill
Bit

The next idea is to use a hollow drill bit that will collect the soil sample inside of the drill
bit. This idea comes from the Honey Bee rover, which takes core samples from the Martian
surface [22]. This design is ideal because a separate housing is not required for the drill bit. The
collected soil is already inside of an enclosed tube. Water evaporated within the enclosed tube
can be channeled to any desired location. The dimensions and weight of this drill bit is given in.

Hollow Drill Bit
Parameter Quantity
Collection Depth 10 cm
Inside Radius 2.45 cm
Soil Volume Collected 188 cm
3
Length of Drill Bit 12 cm

Outside Radius 3.45 cm
Volume of Bit 223 cm
3
Volume of Bottom Panel 37.4 cm
3
Total Volume 261 cm
3
Number of Bits 6
Weight of all 6 Drill Bits 4.2 kg
Table 15: Dimensions and weight of 6 internally sample collecting drill bits.

75
When comparing Table 14 and Table 15, it can be seen that the auger drill bit is lighter
than the internally sample collecting drill bit. This is due to the fact that the auger has its primary
mass in the center of the design, minimizing material used. The internally collecting drill bit has
to cover the collected soil, which requires more material. However, since the internally collecting
drill acts as both the drill bit and as the enclosure to encapsulate the sample, it does not need a
housing unit. When the housing weight is added to the auger drill bit weight, it ends up being
approximately 2.25 times heavier than the hollow drill bit. Hence from a weight and feasibility
perspective, the hollow drill bit was selected for the water extraction. The dimensions of the
internally collecting drill bit are shown in Figure 51 below. The thickness of the drill bit is 1 cm.

Figure 51: The dimensions of the hollow drill bit [21].

iii. Water
Evaporation

Once the soil is collected inside the hollow drill bit, the base of the drill bit must be
closed off to keep water from escaping. The soil sample will then be ready for heating. By lining
the inside of the drill bit with a coil, electrical heating can then be employed. The dimensions of
the electric coil that would be laid along the inside of the hollow drill bit are shown in Figure 52.

Figure 52: The dimensions of the electric coil.

76
By knowing the possible asteroid composition, the energy required to remove water can
be estimated. The most likely molecular compounds to contain water are Saponite, Ferrihydrite,
Hexahydrite, Epsomite, Sodium Thiosulfate, and Borax [7]. Saponite has a varying molecular
mixture of elements and hence it was difficult to find information on extracting the water bonded
to it. But, for the other 5 compounds, the temperature required to remove the water is given by
the following equations:
Ferrihydrite

!!
!

!
∙0.5
!

!"# !

!

!
+0.5
!
à Equation 17 [23]

Hexahydrite

!
∙6
!

!"! !
4 ∙
!
+5
!
à Equation 18 [24]
*More removal requires molecular decomposition

Epsomite

!
∙7
!

!"" !

!
+7
!
à Equation 19 [25]

Sodium Thiosulfate

!

!

!
∙5
!

!"! !

!

!

!
∙3
!
+2
!

!"# !

!

!

!
+5
!
à Equation 20 [26]

Borax

!

!

!
∙10
!

!!! !

!

!

!
∙5
!
+5
!

!"! !

!

!

!
∙2
!
+8
!
à
Equation 21 [27]
The next task was to calculate the energy required to extract the water, based on the
temperature requirement shown in the above chemical dissociation reactions. The expressions
shown in equations below were used to calculate the required energy.

!
!
=
!
∆ à Equation 22

=
!
!
×
!"#$%&#$
× %
!"#$%&#$ !" !"#$
×
!"#$
à Equation 23

Where, Q à Energy,
m à Mass,

77
C
p
à Heat Capacity,
ΔT à Change in temperature for water dissociation,

!"#$%&#$
à Density of molecular compound,
%
!"#$%&#$ !" !"#$
à Suspected percentage of the specific molecule in soil and
Vol
soil
à Volume of the soil to be heated = 1000 cm
3
Using the above equations, Table 16 is created to show the energy required by the heating
apparatus to remove water from the soil. The temperature of the asteroid was assumed at 200 K.

Molecules
C
p

(J/mol.K)
MW
(g/mol)
T
water

Dissociation (K)
ΔT
(K)
Q/m
(J/g)
% in
soil
Density
(g/cm
3
)
Q
(kJ)
Ferrihydrite 105 [28] 169 598 398 247 5% 3.8 46.9
Hexahydrite 268 [29] 228 363 163 191 10% 1.57 30.0
Epsomite 255 [30] 246 400 200 207 10% 1.68 34.7
Sodium
Thiosulfate
361 158 413 213 486 10% 1.67 81.1

Borax 381 [31] 381 393 193 193 10% 1.73 33.4

TOTAL
1,324
226
Table 16: Energy Required to Remove Bonded Water Molecules in Various Molecular
Compounds.
Apart from the asteroid temperature assumption of 200 K [32], a few other assumptions
were made as well. They are: (1) the heat capacity values in the table above represent the average
heat capacity value over the given temperature range, (2) each drill pulls in 166 cm
3
of soil, (3)
the molecular compositions of the hydrated compounds are the true compositions that would be
found on the asteroids and (4) heating apparatus to remove water is 100 % efficient.

From the table above, it will take approximately 230 kJ of energy to heat the soil to a
temperature that will remove all of the water that can be removed from simply heating the soil.
In order to remove any additional water, the soil would have to undergo a chemically driven
extraction process. From the expected average water concentration of 8 % by volume, along with
the possibility of 20 % by volume of water content in the soil, 12 % by volume was considered
as a good estimate such that the equipment is not undersized. It was approximated that 30 % of

78
the volume of soil extracted is water weight. This equates to a volume of 300 cm
3
of water per
extraction. This value will be used to size the equipment used to process the extracted water.
C. Propellant Production

i. Introduction

The water extracted can be used as either a propellant or it can be used in a closed loop
Rankine cycle engine (discussed in Chapter 1 earlier) to produce electricity for electrical and
telecommunication equipment onboard the spacecraft. However, water is not an ideal fuel to use
for this Rankine cycle engine operation due to its high corrosive nature. However, water can be
broken up into its constituting elements to form hydrogen and oxygen, which are much easier to
handle and less harmful to the operating equipment. This section in the chapter will describe how
the water is separated into hydrogen and oxygen from the condensing container attached to the
internally collecting drill bit.

ii. Water
Transportation

As the water begins to evaporate, it needs somewhere to go. By having another chamber
attached to the back end of the drill bit chamber, the water can evaporate and travel into this
additional chamber. Having this chamber at the proper temperature and pressure will allow for
the water entering this chamber to condense and turn into liquid water. This system is shown in
Figure 53 below.

Figure 53: Water condensation system

79
The driving force that will cause the evaporating water to continue to travel into this
condensation chamber will be the equilibrium of water vapor in the evaporating soil chamber to
that of the water vapor in the condensation chamber. As the water vapor in the condensation
chamber condenses, more water will be allowed to travel from the evaporation chamber. Once
the water has been removed from the soil, the condensation chamber will be closed off from the
drill bit chamber. From this point there are two ways to transport the water up the legs of the
spacecraft and into a holding chamber, where electrolysis can take place. The first idea was to
use a pump and push the liquid water up into the next chamber. This would be the simplest way
to move the water, but using an additional pump will add weight to the spacecraft. Another idea
was to use capillary action, where the surface tension of the water will transport the liquid up the
legs and onto the holding chamber. Tiny tubes can be used that will allow the water to migrate
upwards. The equation to calculate the distance that water will travel up a thin tube is given by
the following expression:

ℎ=
!! !"#$
!"#
à Equation 24 [33]
h à Height of the capillary action (~ 1.5 m),
γ à Surface tension of liquid (0.0756 N/m for water)
ρ à Density of liquid (1 kg/m
3
for water)
g à Gravity. Here the gravity term in the equation is irrelevant if the condensation
chamber is pressurized. If the condensation chamber is pressurized to 1000 kPa then,
r à Radius of capillary tube = 0.1 mm.

iii. Electrolysis

The capillary action mentioned above will pull the water from the condensing chamber
and transport it into a holding chamber, where electrolysis will take place. The holding chamber
will have two large tubes, attached by a third tube. Inside the large tubes from the bottom of each
tube will be a metal conducting surface, typically platinum (electrodes). These electrodes will be
hooked up to a D.C electric supply, the positive end being the cathode and the negative end

80
being the anode. Once the water fills the chamber, the electrical charge can be switched on. This
will cause an excess of electrons to form on the surface of the anode and a deficient amount of
electrons on the surface of the cathode. In order to make the water more conductive, a salt, acid
or some other form of ions can be added, but this is not necessary. Salt will just make the water
react faster. The electrolysis system is shown in Figure 54.

Figure 54: Possible set up the electrolysis system that could be used on the spacecraft.
The volume of water in the electrolysis container needs to hold 6 times more water than
in the condensation chamber. The radius of the electrolysis holding container will be 2.5 cm with
a height of 20 cm. The tubes of the electrolysis system will have a radius of 1 cm and an
assumed height of 50 cm (the exact height will have to be re-calculated later when the distance
from the electrolysis system and the gas container is known)

Water is a dipolar molecule, meaning the hydrogen ions have a slight positive charge.
Oxygen ions on the other hand have a slighter negative charge due to the fact that oxygen is
more electronegative than hydrogen. Hence, the hydrogen ion of the water molecule will be
attracted to the anode and the oxygen ion will be attracted to the cathode. Once the ions make
contact with the electrode they are attracted to, electrons will either be given to the atom, in the

81
case of hydrogen, or taken away from the atom, in the case of oxygen, making the ions
electrically balanced. When this happens, a bond is no longer required to stabilize the ions and
the balanced atoms can come off as pure oxygen gas or pure hydrogen gas [34]. The production
of hydrogen and oxygen is shown in the following overall electrolysis equation:

2
!
−→ 2
!
+
!
à Equation 25
However the separate reactions going on at the anode and the cathode are represented by
the following molecular equations:
Oxidation at Anode à 2
!
−→
!
+4
!
+4
!
à Equation 26
Reduction at Cathode à 4
!
+4
!
−→ 2
!
+4
!
à Equation 27
As it can be seen from the above reaction equations for every oxygen molecule produced,
two molecules of hydrogen are produced. The relation between the number of oxygen and
hydrogen molecules produced depending on the current provided by the electric source is shown
in the following expression below:

!
!
=
!"#$
!"#
à Equation 28
R à Universal gas constant = 8.314 J/mol.K,
I à Current (Amps),
T à Temperature (K),
t à Time (seconds), which is assumed to be 5
0
C above the triple point temperature,
F à Faraday’s constant = 96485 C/mol,
P à Pressure (Pa), which is assumed to be 5 Pa above the triple point pressure,
z à Number of excess electrons (2 for H
2
and 4 for O) and
ε à Efficiency ≈ 70 % (common industrial electrolysis efficiency) [35]
From the above expression plots of gas production in volume v/s current and the mass
flow rate of the gas produced v/s current were plotted and shown in Figure 55 and Figure 56. As
the gas bubbles off it will be directed into a storage tank for later use as propellant. An
approximate weight of this electrolysis system is estimated to be around 1.27 kgs.

82

Figure 55: Hydrogen gas and oxygen gas production by volume v/s electric current.

Figure 56: Mass flow rates of hydrogen gas and oxygen gas produced v/s electric current.

0 5 10 15 20 25 30 35 40
0
100
200
300
400
500
600
Current (Amperes)
Production (cm
3
/ second)
Electrolysis Production of Hydrogen and Oxygen

Hydrogen
Oxygen
0 5 10 15 20 25 30 35 40
0
0.05
0.1
0.15
0.2
0.25
0.3
0.35
0.4
Current (Amperes)
Mass (grams / second)
Electrolysis Production of Hydrogen and Oxygen

Hydrogen
Oxygen

83
IV. CHALLENGES OF AN ASTEROID HOPPING MISSION

Asteroids are classified based on their spectral variance. The Spectral color corresponds
to an asteroid surface’s composition. Smaller asteroids do not have large variation in their
surface composition. From Figure 57 (left), it can be seen that the main asteroid belt lies outside
of Mars orbit and inside of Jupiter’s orbit between 2.1 AU and 4.2 AU (some massive C-Class
asteroids are highly eccentric). NASA JPL’s Solar System Dynamics database has catalogued
541,746 asteroids in the main asteroid belt. Half of them are plotted in Figure 57 (right). Figure
57 (right) also shows the donut shaped like feature of the main asteroid belt.

Figure 57: Asteroids in the Solar System (left) and half of the known asteroids in the main
asteroid belt (right) [36].
The epoch of the half of the known main asteroid belt asteroids plotted in the right is 23
rd

May 2014. The torus feature was added for clarity. The origin (intersection of the three axis
shown in red) is the center of the sun and grid spacing represents 1 (AU). The width of this donut
shaped belt above and below the ecliptic plane is approximately 0.214 AU. The orbital elements
of these asteroids were used, as a starting point to evaluate the average ΔV required for hopping
from one asteroid to another. This is a strong function of transit time (time of flight between the
hops). After preliminary iterations a fixed transit time of 100 days was taken as the basis of the
analysis. This resulted in an average ΔV of 1.5 km/s, which is a reasonable ΔV for the amount of

84
propellant required. The total mass of propellant that would be required for this ΔV is dependent
on the Isp of the propulsion system, the total dry mass of the spacecraft (including the power
system) and the total mass of water that needs to be extracted. The average ΔV of 1.5 km/s was
estimated using the following approach:
A. Lambert’s Problem

A random asteroid from the 541,746 asteorids was chosen. The heliocentric position and
velocity of this asteroid served as the initial heliocentric position and velocity of the spacecraft.
The assumption made here is that the spacecraft is in the vicinity of the asteroid a few kilometers
away from it’s surface, moving with the asteroid in the heliocentric frame of reference. This
distance from the center or the surface of the asteroid makes negligible difference in the
calculations of the trajectory parameters to the next target asteroid. Using the 100 days transist
time as a fixed parameter, the “Lambert’s problem” for the transist trajectories to all the
remaining 541,745 asteroids was solved. Lambert's problem assumes two impulsive burns, one at
departure and one at arrival. ΔV of each maneuver is the magnitude of the velocity vector needed
to switch from the asteroid orbit to transit orbit (or vice versa). Continuous or number of impulse
thrust maneuvers > 2 were not considered for this study.
The one with the smallest ΔV requirement was chosen as the next target asteroid, and the
process was repeated for 100 hops. This method is opportunistic in the sense that it always picks
the smallest ΔV option. Therefore, it is not a trajectory optimization algorithm, but serves well
for gathering ΔV data needed to hop around randomly. The illustration of the algorithm used in
solving the Lambert’s problem is described below:
Δθ is the angle between the initial (r
1
) and final (r
2
) heliocentric position vectors of the
spacecraft. Δθ can thus be calculated from the following equation:

∆θ=
!
!
.!
!
!
!
!
!
à Equation 29 [37, Eq. 5.23]
Depending on the inclination (“i”), two cases can be considered: prograde trajectories (0
0

< i < 90
0
) and retrograde trajectories (90
0
< i < 180
0
). Retrograde trajectories can be too ΔV

85
intensive and hence for our analysis, only prograde trajectories were considered instead. For
prograde trajectories if (r
1
x r
2
)
z
≥
0 then the above equation holds good. But if (r
1
x r
2
)
z
< 0,
then Δθ from then the above equation has to be subtracted from 360
0
. Knowing the time of flight
Δt, Lambert’s problem is finding the trajectory joining P
1
and P
2
as shown in Figure 58 below.
P
1
and P
2
represent the initial and final asteroids respectively. “M” represents the Sun and “m”
represents the spacecraft. “i” is the inclination of the transit orbit w.r.t the ecliptic plane.

Figure 58: Illustration of Lambert’s problem [37, Fig. 5.3].
Figure 59 below shows a typical trajectory for 100 hops. The spooling nature of
the trajectory owes to the fact that changing the direction of spacecraft velocity is more
expensive than changing the magnitude alone. Therefore asteroids with wildly different orbital
planes (inclination) and non-circular orbits (large eccentricity) are unlikely targets when seeking
the minimum Δv transit. The trajectory algorithm
targetted asteroids with inclination around i = 3
0

and eccentricity e = 0.08. In addition to these
values being close to circular orbits in the ecliptic
plane, it also has to do with the abundance of
asteroids with these orbital elements.
Figure 59: Trajectory for 100 hops.

86
There are of course many more unknown asteroids in the middle belt than found so far. In
order to estimate the average ΔV depending on whether there are 10 million, 100 million or 1
billion asteroids, more asteroids had to be predicted. An intelligent approach was adopted to
predict this. The distribution of all the orbital elements of known asteroids were considered.
Figure 60 shows distributions of orbital elements of the known middle belt asteroids. These
distributions were used to generate the required number of asteroids. An example of 10
7
asteroids
is shown in Figure 61.

Figure 60: Orbital elements of all known asteroids in the main asteroid belt.

Figure 61: Prediction of 10
7
asteroids not discovered and logged yet.

87
As the number of asteroids increases, the computational load to computer the trajectory
algorithm becomes heavy quickly. Lending to the fact that only asteroids with small inclination
and eccentricity are likely targets when ΔV is of premium, most asteroids in the cloud can be
ignored. Figure 62 shows the same asteroid cloud (originally 10
7
) sliced to contain only those
with inclination 3
0
± 0.5
0
degrees. By restricting eccentricity close to circular orbits, further
reduction in number of irrelevant asteroids is possible. This was utilized in the case of 10
8
and
2.5 x 10
9
asteroids, whose databases contained 6 million asteroids.

Figure 62: 10
7
asteroids reduced to 600,000 to handle computation load.
This study was decided to contain only asteroids with diameter > 50 meters. Figure 63
(next page) shows the total number of asterioids with a specific size, present in the main asteroid
belt, depending on the diamater of the asteroid. The red line in the figure is a power law fit to
approximate the distribution. From this figure, there are 2.5 x 10
9
that fall into the category of
asteroids with diamater > 50 meters. Table 17 below tabulates the average ΔVs for t
transit
= 100
days for various numbers of Asteroids. Figure 64 (next page) shows the typical ΔV distribution
of 100 best hops from 2.5 x 10
9
asteroids choices.

Asteroids
ΔV
avg
Standard Deviation
580,000 1.5 km/s 0.4 km/s
10
7
1.1 km/s 0.3 km/s
10
8
0.7 km/s -
2.5 x 10
9
0.5 km/s 0.15 km/s
Table 17: Tabulation of average ΔV for various numbers of Asteroids in the main asteroid belt.

88

Figure 63: Main belt asteroid size – frequency distributions [38].

Figure 64: ΔVs of 100 best hops from 2.5 x 10
9
asteroids choices.

89
V. ACQUIRING OWNERSHIP OF AN ASTEROID
The article II of the 1967 Outer Space Treaty states that – “Outer space, including the
moon and other celestial bodies, is not subject to national appropriation by claim of sovereignty,
by means of use or occupation, or by any other means” [39]. This implies that private sector
space pioneers need not apply for permission to use space-based resources from their national
governments as such authorizations are disallowed under the treaty. One interesting exception is
orbital slots for GEO, which is recognized as de facto property in space. However, GEO Orbital
rights are always subject to negotiations. Another example is exclusive seabed mining. Licenses
are granted and recognized multilaterally for a limited area for a limited period of time.

The exact form of commercial rights for Asteroid mining may depend on how various
nations agree. Long before private space companies travel into space in search of wealth, many
more legal areas have to be negotiated and written into arguments. A proposed amendment to the
treaty should be – those who have the ability to exploit space resources, without harmful
contamination, for peaceful exploration & commercial purposes should have the right to do so.
This section of the chapter conceptualizes a scheme on how to claim ownership of an asteroid
and lease it to future spacecrafts for mining. When the water mapping spacecraft recognizes a
suitable asteroid containing water it collects water map data through the instrumentation. This
information will be for sale for other parties that would like to seek mining opportunities. But
since “our” spacecraft was first to get there and map the asteroid, it can drop a beacon / pinger
device on the asteroid surface to claim ownership.
A. Pinger Design
The pinger would be a transmitter, which would continuously transmit the Identification
(ID) information of an asteroid, such as its name. The transmission will be over a certain large
range, so that an incoming mining spacecraft seeking it can have enough time for orbital
corrections / maneuvers in order to head towards its sphere of influence. Also, during the
mapping phase of the asteroids, it will aid the mapping spacecraft in acknowledging that the

90
specific asteroid has been tagged. The estimation of the data that needs to be transmitted is
crucial in determining various factors of this beacon device such as the data rate, power and
frequency of transmission.
i. Data
Type

The data that the pinger needs to transmit is the ID information of the asteroid, in other
words the name of the asteroid. According to the naming convention of an asteroid, every
asteroid that has been named till date has nine characters in its name. For example, an asteroid
named as 1999AN101. The first four characters denote the year in which it was discovered and
the subsequent alphanumeric characters denote when in a year it was discovered. These
characters are first converted to an ASCII decimal representation [40]. Those decimal numbers
are then represented as binary digits, with each character represented by 7 bits. These bits add up
to a total of 63 bits. Adding another seven bits to represent the start an end of the message would
make the total message size to be 70 bits at most for every asteroid. Table 18 below shows the
data breakdown of the Asteroid named – 1999AN101.

Character ASCII Decimal Code

Binary Conversion Number of Bits
1
49 0110001 7
9
57 0111001 7
9
57 0111001 7
9
57 0111001 7
A
65 1000001 7
N
78 1001110 7
1
49 0110001 7
0
48 0110000 7
1
49 0110001 7
Total number of bits 63
Table 18: 1999AN101 asteroid’s name’s data breakdown.
ii. Pinger
Schematic

The pinger device will have all the components of a general communication transmission
system like source compression, channel coding modulation and amplification. However, there is
no requirement to compress data and introduce redundancy for channel coding as the message
itself is very short and is continuously transmitted.

91
If one transmission received has an error, the subsequent transmission received can lead
to the correction. The data of 70 bits is modulated using Binary phase shift keying (BPSK)
modulation. This modulation is simplest digital modulation scheme, which is extensively used in
satellite communication. In BPSK two phases of the same sinusoidal carrier frequency to
represent the logic level 1 and level 0 respectively, is transmitted. The 0
0
phase signal represents
bit 1 and the 180
0
phase shifted signal represents bit 0 [41] [42] [43]. Mathematically BPSK
modulation can be described by the following expression:
= ∗ = ∗cos 2
!
=±cos 2
!
à Equation 30 [44]
= cos 2
!
à Carrier Signal,
m(t) = +1 or –1 (0 ≤ t ≤ T
b
) à Message Signal, where -1 is the logic 0 and +1 is logic 1,
T
b
à the bit duration in seconds,
A à is the amplitude of the carrier frequency in Volts and
f
c
à frequency of the carrier signal in Hz
The functional block diagram of the BPSK modulation is shown in Figure 65 below.

Figure 65: Pinger device’s communication system architecture.
The off-the-shelf electronics components were identified for implementation of the
functional blocks shown in Figure 65 above. These components are listed in Table 19 below.
The electronic circuit diagram is shown in Figure 110 in the appendix.

Component
Type
Off-the-shelf
component

Functionality Image
Power
Consumption
Microcontroller
Texas
Instruments
MSP430 [45]
This is an ultra-low power
microcontroller to store the
message bits. It will perform
Non-Return to Zero (NRZ)
encoding on the message.
7.2

W

92
Frequency
mixer
Linear
Technology
LT5560 [46]
This double balanced modulator
performs the multiplication of
the data signal and the carrier
frequency to generate the
modulated BPSK signal. The
frequency range of LT5660 is
from 0.01 MHz to 4 GHz.
0.067 W
Band Pass
Filter
Triquint
Semiconductor
885024 [47]
The Band Pass filter is used to
eliminate unwanted frequencies
that are generated because of
multiplication of the signals in
the frequency mixer. It allows
only the carrier frequency to be
transmitted, which reduces the
noise in the transmitted signal.
N/A
Power
Amplifier
Pacific
Monolithics
PM2105 [48]
The power amplifier boosts the
signal to a high power in order
to enhance the signal to noise
ratio and achieve a more
compact antenna design.
0.032 W
Table 19: Off-the-shelf electronics components for the pinger device.
Figure 111 in the Appendix shows an illustration of the pinger device’s BPSK modulation
scheme.
iii. Pinger
Antenna
Design
from
link
budget

A helical antenna, which would be used as an Omni-directional antenna, was chosen to
propagate the electromagnetic waves from the pinger device. This is because it will propagate in
all directions. The height, diameter, number of turns and the distance between the turns are the
important design parameters of the helical antenna. A communication link budget was created to
iterate and optimize the sizing of these parameters. Figure 66
shows these physical parameters of a typical helical antenna. A
sphere of influence of ~17000 km was chosen as a reasonable
range of the signal with a Signal to Noise (SNR) ratio of 10.
Figure 66: Typical physical parameters of a helical
antenna.

93
A power source of 1mW was selected as a reasonably deliverable power source for the
transmission signal with a frequency just under 2 GHz (higher L-Band frequency). Fixing these
parameters helped to iterate and optimize the antenna sizing parameters. The final link budget,
which shows the final iterated antenna design parameters are represented in Table 20 [49].

Item Symbol

Units Quantity Expression
Frequency f GHz
1.96
Input
Wavelength λ m 0.1531 Speed of Light / frequency
Transmitter Power Pt mW 1 Input
Number of turns N N/A 10 Input
Transmit Antenna Diameter D cm 2.94 Input
Transmit Antenna
Circumference
C m 0.0922
2πD
Transmit Antenna Height H cm 3.5 Input
Spacing between turns s m 0.0035 H/N
Transmitter Losses Lt dB -1.10 Cassini input
Transmit Antenna Gain (net) Gt dB -0.5101 10.3 + 10
log
!"
(
!
!
!
!
!
)

Propagation Path Length S km 16900 Input
Space Loss Ls dB -182.8529 147.55 - 20log
10
(1000 x S) - 20log
10
(f)
Receive Antenna Diameter Dr m 1.5 Input
Receive Antenna Gain Gr dB 27.1711
20log
10
(π) + 20log
10
(D
r
)
+ 20log
10
(f) + 10log
10
(0.55) - 20log
10
(c)
Data Rate R bps 70 Input
Required SNR
Req.
Eb/No
- 10 Input
Table 20: Communication link budget of the pinger device.
B. Pinger Device’s Power Source
If a spacecraft will be visiting multiple asteroids in its mission to search, map and tag
them for water then multiple pinger devices had to be carried on board. Hence the mass of the
pinger device has to be as low as possible. Also the power supply of the pinger device has to be a
constant and long lasting power source. This is because it is unexpected to be replenished.
Ordinary Li-ion batteries would not be a good option as they are very fragile and need heavy
protection cases. The asteroids will be very harsh environments with extreme temperature
variations. The Li-ion batteries are not very strong in withstanding such extreme temperatures.

94
Hence a more reliable and lighter option of
beta-voltaic cell was considered. The beta voltaic
cells currently manufactured by City Labs are in
the nano-watt range.
Figure 67: City Labs’ NanoTritium™
beta-voltaic power source [50].
They have a lifetime of 20 years, current range of 50 – 350 nano amperes, voltage range
of 0.8 – 2.4 volts and a power supply of 840 nW [50]. However, the pinger requires a power of 1
milli watt. Hence designing a custom radioisotope beta-voltaic battery is more appropriate for the
pinger device. The working of a beta-voltaic cell is similar to that of a solar cell. Except that
instead of a photon striking on a p-n diode and generating electron-hole pairs, an electron strikes
the p-n diode to generate the electron-hole pairs. The beta-voltaic cell construction consists of a
conducting plate coated with an electron-emitting radioisotope material. The radioisotope
candidates considered for the custom radioisotope beta-voltaic cell of the pinger were Nickel –
63 (
63
Ni), Strontium – 90 (
90
Sr), Cesium – 137 (
137
Cs) and Technetium – 99 (
99
Tc). An illustration
of the construction of a beta-voltaic cell is shown in Figure 68.

Figure 68: Construction of the radioisotope beta-voltaic cell [51].
The sizing of the beta-voltaic cell is done using the expression shown below. The surface
area of the receiving plate and the corresponding beta-voltaic material’s surface area is given by
“A
s
” as shown in the equation below obtained from Child Langmuir’s law [52]:

!
=
!!
!"#
!!
!
!
!
!!
!
!
!
!
!
!
!
à Equation 31

95
P
req
à Power required by the pinger = 1 mW,
m
e
à Mass of an electron = 9.10938356 × 10
-31
kilograms,
q
e
à Charge of an electron = 1.60217662 × 10
-19
coulombs,
ε
0
à Permittivity of free space = 8.854187817 × 10
−12
F/m,
L à Plate gap in centimeters (iterated in MCNP)
E à Total energy deposited on the receiving plate in Volts (obtained from MCNP)
Keeping the width of the beta-voltaic surface and the width of the receiving plate fixed at
1 cm, the surface area obtained from the above equation can be used to obtain the length of these
rectangular surfaces. Figure 69 below shows the energy density per cm
2
v/s the gap distance
between the beta-voltaic material surface and the receiver surface of the p-n diode. The plots are
for the four beta emitter radioisotopes –
63
Ni,
90
Sr,
137
Cs and
99
Tc considered for this application.

Figure 69: Energy flux v/s distance between the beta-voltaic material and the receiver plate.
From the above figure it can be seen that
137
Cs gives has the highest energy deposition,
whereas
63
Ni has the lowest energy deposition (three orders of magnitude lower than
137
Cs).
Unfortunately
137
C also has some gamma signatures, which could be very detrimental for the
electronics. Shielding them would make the device very bulky.

96
However,
90
Sr, which is the second highest in energy deposition (same order of
magnitude as of
137
Cs), does not have any intense gamma signatures that would require heavy
shielding. It is in essence the purest in beta emission. Figure 70 below shows the length of the
beta-voltaic surface and the length of the receiver p-n diode plate with respect to the gap between
these surfaces. It can be seen from the plot below that
63
Ni is the most impractical candidate for
this application. The length is in meters for this isotope. However in case of
90
Sr the length is in a
reasonable range. A length and width of 1 cm and a plate gap of 1.5 mm will deposit enough
energy from the beta-voltaic cell to provide the required power of 1 mW for the beacon. Hence
90
Sr was chosen as the best radioisotope beta emitter candidate.

Figure 70: Length of beta-voltaic surface and the receiver p-n diode plate v/s plate gap.
C. Pinger Device’s final assembly and mass budget
There are two options for deploying the pinger device on the asteroid surface: (1) If the
spacecraft lands on the surface then drill the pinger into the surface or (2) if the spacecraft’s
primary mission is water detection and gathering water map information, which does not include
landing on the surface then the pinger device could be loaded in a penetrator probe.

97
The penetrator probe will be designed to impact the surface and penetrate the subsurface.
Either way a general assembly of the pinger device on a probe is shown below.

Figure 71: Pinger device’s penetrator probe.
The radiation shielding protects the communication electronics from any radiation from
the beta emitter radioisotope power source. Due to high velocity impact, a long structural part in
the front of the penetrator probe is conceptualized to take the shock and protect the pinger
device. When this probe is realized from the spacecraft, it will be given an initial angular
momentum to enable the drill bit at the impactor tip to drill through the asteroid surface, which in
turn will enable a stronger grip onto the target body. The helical antenna on the penetrator probe
will be sticking out to the open space for transmitting the beacon signal.

98
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101
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103
Chapter 3

Applications of Radioisotope ThermoPhotoVoltaic Power Source

I. ONE WAY DEEP-SPACE LASER COMMUNICATION FROM A CUBESAT
(EUROPA ORBITER MISSION)
A. Abstract

With NASA’s renewed interest in exploring the gas giants and their moons, cost
reduction is on the forefront of concern. Concurrently, the surge in successful CubeSat launches
to Near-Earth Orbit, has spawned interest in examining their use for deep space applications. In
current CubeSat flight heritage, the communication systems use RF signals on lower bands, milli
Watt class of power and compact omnidirectional Antenna system. But in order to close the
communication link for distances greater than 10 AU, kW class of power is required to overcome
the >250 dB space loss. Subsequently, the antenna diameter size goes into the order of a couple
of meters. This results in the communication system to be bulky and spacious, increasing the cost
of launch and operations into a billion-dollar tier mission. The study discussed in this chapter
targets the reduction of this generic RF communication system’s mass and size, by replacing it
with laser communication technology, which can fit in a 6U CubeSat constraint. Achieving this
miniaturization can lower the cost of deep-space missions to the order of a million-dollar tier
making it more accessible to small budget organizations such as university research labs.
B. Data Requirement

In order to demonstrate the concept of deep-space laser communication architecture,
some sample mission architectures were evaluated from NASA’s Decadal survey - Vision and
Voyages. These sample mission architectures are as follows: (1) Enceladus (Saturnian moon)
Orbiter, (2) Jupiter-Europa Orbiter, (3) Io (Jovian Satellite) Observer, (4) Ganymede (Jovian
Satellite) Orbiter, (5) Trojan Tour (L4 / L5 Asteroids of Jupiter and Sun), (6) Titan Lake Lander
and (7) Uranus Orbiter. Out of all the missions mentioned above, the Europa Orbiter Mission
was formulated in detail.

104
The current state of the art near-Earth laser communication system supports about 10
Gbps link from GEO, around 100 bps from Mars and a mere 0.25 bps from Pluto. But the current
Radio Frequency link performance of Cassini’s 20-W X-band TWT, 4-m HGA is approximately
14 kbps [1 – Fig 2.2, Table 2.1]. Hence in order to compete with RF technology, the laser
communication technology requires significant improvements (> 50 dB) in optical link
performance. In order to achieve this large performance over the current state of the art, the
amount of signal power delivered to the receiver has to be improved by increasing the transmitter
power and antenna gains and / or installing a receiver telescope beyond the earth’s atmosphere
[1]. This receiver telescope could be placed on the moon or in a LEO / GEO orbit around Earth.

This chapter lays out the design parameters and power source requirements of the laser
communication transmitter system for transmitting downlink data only. Due to the extremely
large space loss and difficulty in maintaining pointing accuracy from ground station to a
spacecraft of size in the order of 6U CubeSat, comprehending an uplink system using the same
technology has been omitted for now and recommended as future work. Hence, in order to close
the uplink data communication link, existing RF technology will be used. However, a new
deployable Passive Microwave sensor (Dish Antenna reflector type) that can be integrated in a
6U CubeSat platform has been discussed. This serves as a supporting sensor for the laser
communication instrument because in existing CubeSat communication systems only deployable
omnidirectional antenna systems are being used for two-way communication.

i. “Off-‐the
shelf”
science
instrumentation
for
deep
space
CubeSat
missions

An off-the shelf instrumentation package list mostly common between the 7 missions
mentioned above has been listed below in Table 21. This instrumentation list focuses in
achieving limited science objectives and can be assembled to fit within the constraints of
multiple 6U CubeSat platforms. Some of the instruments are shown in Figure 72 and Figure 73
(next page).
Science
Instrument Type
Science Instrument
Mass
[kg]
Volume
[cm
3
]
Power
[W]
TRL
Rad
(kRad)
Thermal
Limits
X-Ray / Gamma
Ray Detector
X-123 CdTe [2] 0.18 [2] 175 [2] 2.5 [2] 7 2 – 10
- 40
0
C à
40
0
C
Infrared
Spectrometer
Argus Infrared
Spectrometer [3]
0.23 [3] 180 [3] - 9 2 – 10
- 40
0
C à
40
0
C

105
Surface Camera
NanoCam C1U
(High Resolution
Camera) [4]
0.166
[4]
501 [4]
0.66
[4]
8 2 – 10
- 40
0
C à
40
0
C
Mass Spectrometer LVGEMS [5] 0.25 [6] 32 [5] 0.5 [5] 7 -
- 40
0
C à
40
0
C
Radar Altimeter Mini-SAR

3.1 2888 40 9 2 – 10
- 40
0
C à
40
0
C
Radar Sounder MicroMAS < 1 < 1000 10 6 2 – 10
- 20
0
C à
65
0
C

Thermal Imager HIBRIS [6] 7.1 [6]
22 x 26 x
21 [6]
8 [6] 8 2 – 10
- 40
0
C à
40
0
C
Magnetometer Multiple 0.23 0.4 < 1000 9 2 – 10
- 50
0
C à
85
0
C
Dust Analyzer LAMBDA - - - 8 2 – 10
- 40
0
C à
40
0
C
Table 21: A list of instrument packages for the 7 mission architectures.

Figure 72: X-123CdTe (X-ray and Gamma-ray detector system [2], Argus Infrared
Spectrometer [3] and NanoCam C1U (High Resolution Camera) [4]

Figure 73: Laser Anemometer and Martian Dust Analyzer (LAMBDA) [7], Low Voltage
Gated Electrostatic Mass Spectrometer (LVGEMS) [5] and Highly Integrated Micropayload for
Broadband Infrared Spectrometry (HIBRIS) [6]

106

ii. Science
Instrumentation
Data
Budget
(Example
Mission
–
Europa
Orbiter)

Figure 74 below emphasizes on the strength of the data rate to portray the quality of
information that can be obtained. And Table 22, lists the respective data rate requirements of the
science instruments on board for all the sample missions considered from the decadal survey.

Figure 74: Data rate requirement to portray the quality of information [1 – Fig 1.1]
Instrument Average Science
Data Rate with contingency
Enceladus
Orbiter
(kbps)
Jupiter – Europa
Orbiter (kbps)
Io
Orbiter
(kbps)
Ganymede
Orbiter
(kbps)
Uranus
Orbiter
(kbps)
Medium Angle Camera 3120 1065 - 6400 4.2 (WAC)
Thermal Imaging Radiometer 3120 15 1150 - 0.18
Mass Spectrometer 4.42 2 1.53 1.51 0.064
Dust Analyzer 0.39 - - - 0.05
Magnetometer 0.39

4 1.61 0.88 0.15
Laser Altimeter - 2 - 28 -
Ice Penetrating Radar - 140 - 45 -
Narrow Angle Camera - 10700 3900 25000 1.05
UV Spectrometer - 10 - 403.2 0.02
VIS – IR Spectrometer - 11400 - 5000 7.8
Particle and Plasma Instrument - 2 0.09 80.99 0.25
Table 22: Instrument Average Science Data Rate with contingency for the various sample
science missions considered for this study. [8 – Table 3-1], [9 – Tables 3-1 – 3-5], [10 – Tables
3-1 – 3-12] and [11 – Tables 3-1 – 3-10]

107
The data rate available for dumping data through the downlink channel governs the
mission lifetime. Currently operating deep space crafts such as Cassini or Mars Reconnaissance
Orbiter are limited to dumping high definition videos. This is due to the small data rate available
to them through the RF channel on which they are operating. From Table 22, the average data
rate required to be transmitted, was approximated to 3 Mbps. This value for data rate was used as
the target data rate requirement while designing the link budget of the laser com. system.

Except the Titan Lander mission (discussed later), all the other sample decadal survey
missions reviewed, are orbital missions. Hence to determine the optimum parking orbit at target,
the surface image resolution required for the science objective, plays as the crucial parameter.
For example to study the gravity field variations in Europa, which has a surface area of 3.09 x
10
7
km
2
[12], a Doppler tracking device coupled with a surface camera has to be used. The
surface camera would require a resolution of a minimum 15 meters / pixel. Table 36 in the
appendix shows the tabulation used in obtaining the optimized parking orbit, for the required
image resolution. From Table 36, the optimal parking orbit altitude required to accomplish the
mission is 121.875 km. A total of 43657 pictures would be required to map the entire surface.

The high-resolution camera – NanoCam C1U mentioned in Table 21 earlier has a color
depth of 10 bits, which implies that each color (red, blue and green) is represented by 10 bits.
Therefore, 30 bits are required to represent three channels of color red, green and blue. Also a
header size of 54 bytes has to be assigned to each image. Thus the size of an uncompressed
bitmap image in bytes can be calculated by the following expression:

=
!∗!
!
+ à Equation 32

P à Total number of pixels = 2048 x 1536
B à Number of bits per pixel = 30 bits
H à Header file of each image = 54 bytes
From the above equation, the size of an uncompressed bitmap image = 11796534 bytes or
11.25 MB. The number of images required to cover the entire surface area of Europa from a
specific orbital altitude is multiplied by this size of 1 uncompressed bitmap image to obtain the
total uncompressed data size of all the images. The total time it would take to dump all that data
with a 3 Mbps data rate is 45.48 hours. For a data rate of 65 kbps it will take 87.45 days.

108

Figure 75: Orbital Altitude (km) v/s Picture Count for Jupiter – Europa Orbiter Mission

Figure 75 shows a plot of parking orbit altitude (km) v/s average number of pictures
required to complete the science objective for the Europa Orbiter Mission. Figure 76 shows a
plot of parking orbit altitude (km) v/s ground image area (km
2
) for the same Europa Orbiter
Mission. Figure 112 in the appendix shows the data downlink transmission time (hours) v/s the
number of pictures being sent to Earth for a data rate of 3 Mbps. The parking orbit has to be
precisely optimum in order to meet the science objectives and most importantly be able to send
the data back within a tolerable mission duration time. Longer duration implies more radiation
shielding, especially for a Europa mission where the radiation environment due to Jupiter’s Van
Allen belt is too harsh (discussed in the next section – trajectory analysis for Europa Orbiter).
Hence a laser communication downlink will benefit hugely in minimizing the radiation shielding
by speeding up the mission, which is achieved with a high data rate transfer.
C. Re-creation of an existing Europa mission trajectory design for radiation analysis
A detailed trajectory design for the Europa Orbiter Mission was borrowed and re-created
to evaluate the mission duration, estimate rough launch & operational costs, and evaluate
radiation dosage during Europa Orbit Insertion. This trajectory design (Table 37 – Appendix)
was referred from the paper – “Jovian tour design for orbiter and lander missions to Europa”
[13].

109

Figure 76: Parking orbit altitude (km) v/s ground image area (km
2
) for Europa Orbiter.

Figure 77: Top view of Earth – Jupiter Cruise Phase (Refer Table 37 in the Appendix)

110

Figure 78: Jovian Moons Tour Phase (Refer Table 37 in the Appendix)
Based on the trajectory, the GRID 2 Jovian Radiation Model was used to predict the
radiation dosage that the spacecraft would encounter in its lifetime. The Grid 2 model is a
“numeric implementation of the GIRE 2 Jovian Radiation Model for estimating trapped
radiation” [14]. A FORTRAN code was written to extract the information from the GRID 2
model. Figure 79 shows the combined electron and proton fluence output obtained from the
GRID 2 model based on the trajectory design borrowed for the Europa Orbiter Mission. From
Figure 79, it can be seen that there is a big spike in the radiation level during November 2029,
when the Europa Orbit Insertion is in effect. This is expected due to flying through the intense
Van Allen belt of Jupiter in which Europa orbits around Jupiter. From Figure 80 and Figure 81,
it is estimated that for the example trajectory considered, a 7 mm Aluminum shielding would be
able to protect the electronics for a period of roughly 18 days after circularization around Europa
is achieved. The circularization around Europa happens at an altitude of 6800 km, which is still
pretty high for the surface cameras. Hence a solid rocket motor for the CubeSat is proposed as a
solution to lower the altitude to the 120 km range.

111

Figure 79: Combined Electron and Proton Fluence Output for the Europa Orbiter Mission.

Figure 80: Calculated dosage based on 7 mm Aluminum shielding.

112

Figure 81: Fluence (cm
-2
) v/s Energy (MeV).
D. Sensor Requirements

As mentioned earlier, apart from the uplink RF communication system (supporting
secondary instrument for command, data and handling), the primary laser communication
instrument also requires a state of the art, robust Sun Sensor. This sensor has to be implemented
on a 6U CubeSat platform for Attitude determination and Control in order to achieve the highest
pointing accuracy when the downlink data is being transmitted from the CubeSat to Earth. Along
with the Sun Sensor, another supporting instrument required is a jitter / inertial sensor for jitter
isolation and rejection [1]. To achieve a stable line of sight, the laser communication terminal
must properly isolate and reject the spacecraft platform jitter and spacecraft attitude control
errors. This is accomplished using a combination of vibration isolators and a pointing
stabilization control loop as shown in Figure 82. The design and the working operation of this
sensor is recommended as future work.

113

Figure 82: Block diagram showing jitter isolation and rejection for a laser communication
terminal [1 – Fig 2.10]

i. Observations
Required

The Laser Communication downlink system is required to transmit data only. This
primary instrument pairs with a ground receiver station to complete the laser communication
architecture. The Ground receiver has to observe the incoming laser transmission.

It is required to collect the optical signal, filter out the noise photons and convert it into
an electrical signal to interpret the data from the signal. The secondary instrument on the
spacecraft, which is a passive microwave sensor (Dish Antenna), has to capture the transmitted
RF Command, Data and Handling (CDH) signal from the Deep Space Network (DSN). The sun
sensor as a tertiary sensor observes the sun or collects the photons from the sun, to measure the
orientation of the spacecraft to aid the pointing accuracy of the Laser beam towards earth.

ii. Quality
of
the
Data:
Pointing
Accuracy

One of the main advantages of laser communication over RF communication is that, laser
transmitters compared to RF transmitters can achieve very low Beamwidth angles for a signal
transmission. This implies that the directive gain that can be obtained is very high. Figure 83
below illustrates that.

114

Figure 83: Comparison of RF and optical beam spreads from Saturn [1 – Fig 1.2].
The approach that is used to design the Acquisition, Tracking and Pointing (ATP) is to
use the Earth’s reflectance of the sun as a beacon. The subsystem architecture of this approach
within the laser transceiver is shown in Figure 84 (next page). The visible light spectrum lies
between 0.4 to 0.7 micrometers. This spectrum will be used to receive the image of Earth by the
convex telescope as shown in Figure 84. This will be the same optical telescope, which will be
used to transmit the laser signal towards the Earth ground station.

Figure 84: ATP subsystem in the laser transceiver using Earth as the beacon [15 – Fig 3.11].

115
The external jitter, which is measured by the jitter sensor, will provide the vibration
signal that would be used to compensate for the spacecraft’s vibrations. When the laser
transmitter is transmitting, the fine beam steering mirror shown in Figure 84, will be adjusted by
a servo system inputted with the error signal from the external jitter sensor. The accelerometer
unit shown in Figure 84 above is the representation of the jitter sensor. The fast tracking signal
on the beam steering mirror would be connected to the CubeSat’s attitude control system (3 axis
reaction wheel system) to adjust its attitude. A band interference filter would be placed over the
CCD to mitigate noise photons from the Sun [15].

iii. Transmission
Power
Requirement
for
Europa
Orbiter

The diameter of the laser transmitter telescope’s convex lens was fixed to 10 cm. This is
because it is intended to fit inside a 1U space of the 6U CubeSat. The equation below computes
the required transmission power for a data rate of 3 Mbps.
à Equation 33 [15 – Eq. 2.26]
Where,
- Diameter of the transmitter telescope = 10 centimeters
- Diameter of the receiver telescope = 5 meters
- Quantum Efficiency x Signal losses in transmitter and receiver = 0.2
- Wavelength of the transmitting laser light = 1.064 µm
The wavelength chosen above for this laser communication downlink is the same for
Nd:YAG – Neodymium-doped yttrium aluminum garnet crystal. Various wavelengths of
commercially available lasers are shown in Figure 85 below as an illustration.
- Approximate furthest distance between Europa and Earth = 8.817 AU = 1.3191 x 10
12
meters

- Signal counts per decision interval = 6.2
- Planck’s constant = 6.62607004 x 10
-34
kg.m
2
/s
- Frequency of the transmitting laser signal
=
!
= 0.5 bits per pulse
- Data rate = 3 Mbps
From the above equation, transmission power P
t
required to close the laser downlink
from Europa = 443.57 W ≈ 0.44 kW.
P
T
={(4πR)(λ)/(πD
T
)(πD
R
)}
2
{(hν)(D
rate
)(S)/αQF}
D
T
D
R
Q.F
λ
R
S
h
ν
D
rate

116
Figure 85: Wavelengths of commercially available lasers [16].

10 W is a reasonable amount of power that can be potentially generated in a 6U CubeSat
platform using a Radioisotope ThermoPhotoVoltaic (RTPV) power source. For a data rate of 65
kbps, the transmission power P
t
required to close the laser downlink from Europa will be 9.6 W.
That is still a commendable data rate for transferring low-resolution pictures but to transfer the
science data required to accomplish the science objectives, it would take roughly 1120 hours of
transmission time. Based on the harsh radiation environment that the spacecraft would be
operating in, 1120 hours is beyond the lifetime of the spacecraft. After the data has been
collected, a second orbital maneuver could be possibly performed to leave the Van Allen belt
region of Jupiter. This would buy more time for the on-board laser communication system to
dump the entire data using the lower power source.

iv. Laser
Transmitter

Figure 86: Top view of the laser transmitter layout [1 – Fig 5.15].

117
The laser source type chosen for the laser transmitter is an Nd:YAG Solid-State Laser
diode. This laser diode changes incoherent laser light from laser diodes to coherent light. The
output light from several high-power pump laser diodes with inherently poor spectral and spatial
quality beams are summed up and converted into a single beam of light. Thus the output is a
bright, highly coherent light with excellent spectral and spatial beam properties [17]. A top view
of the laser transmitter layout is shown in Figure 86 earlier (previous page).

v. Ground
Station
Optical
Receiver:
Pin-‐Photodiode
Design

For the ground system receiver, pin-photodiode type detector is considered to convert the
incoming optical radiation into an electrical signal. A photodiode is operated as a reverse-biased
diode, where incident photons are absorbed to generate electron-hole pairs, leading to a current
in the electrical circuit connected to the diode’s electrodes [18]. The material used for the pin-
photodiode is InGaAs because it has a typical operating range between 1 – 1.6 micrometers and a
cutoff wavelength of 1.65 micrometers. HgCdTe diodes are not used or opted for this design
because it requires cooling to 77 K that will be difficult to achieve in a 6U CubeSat platform,
where volume is extremely constrained. The cross-section and bias supply of the pin-photodiode
is shown below in Figure 87 (a). The equivalent circuit diagram of the pin-photodiode is shown
in Figure 87 (b): I
D
– dark current, G
D
– conductance, C
D
– junction capacitance, C
M
– further
capacitance, L
M
– series inductance and S
P
– photo induced current.

Figure 87: (a) Left – Cross-section and bias supply of the InGaAs pin-photodiode [18 – Fig 7.6]
(b) Right – Equivalent circuit of the pin-photodiode. [18 – Fig 7.8]
The power received by the optical detector can be obtained from the following
expression:

à Equation 34 [15 – Eq. 2.25]

P
R
= (D
rate
/α)(S /Q)(hν)

118
From the equation above P
R
≈ 0.0035 nanowatts, which is extremely low power for the
pin-photodiode to work effectively. The dynamic range of the InGaAs photodiode is typically
several orders of magnitude, extending from the noise level up to diode currents of a few
milliamperes (optical input power of a few milliwatts) [18]. The only way to increase the
received optical power is to increase the transmission power or increasing the receiver optical
telescope’s diameter to a couple of 100s of meters.
If the transmitted power is pulsed at 25 kW and everything else is kept the same as
before, the data rate increases to 163.42 Mbps but the P
R
≈ 0.0002 microwatts, which
unfortunately is still six orders of magnitude less than the required. Hence unfortunately from the
analysis so far, there is no other way to accomplish this mission unless a ground optical detector
system could be established with a ridiculously large receiver telescope! Since the received
power is so small, the receiver station at the ground becomes the worst receiver station option.
The losses due to the atmosphere will be immense and the final received power might just be
noise with no data. An assumption is made here that the received power obstacle mentioned
above, can be achieved after the innovation of high-tech and high efficient pin-photodiodes in
the future. Hence this is included in future work.

There is rapid progress in research growing in the innovation of high-tech and high
efficient pin-photodiodes. The low received power obstacle mentioned above, will be tackled
adequately. Increasing the received power to the order of milliwatts would imply that the data
rate is in the order of several Gbps. The table below shows the typical parameters of pin-
photodiodes made from Indium-Gallium-Arsenide (InGaAs) for data rates of several Gbps.
Pin-Photodiode Parameters InGaAs
Operating Range (micrometer) 1.0 – 1.6
Quantum Efficiency, η 0.95 at 1550 nm
Rise time, fall time (ps) 50
Capacitance (pF) 0.7
Dark Current (nA) @ 290 K 1

Bias voltage (V) -5
Diameter of active area (micrometer) 80
Pin-Photodiode Parameters InGaAs
Table 23: InGaAs pin-photodiode parameters. [15 – Table 7.2]

119
II. ANALYSIS OF ENTRY, DESCENT AND LANDING (EDL) OF AN AMPHIBIOUS
QUADCOPTER SWARM FOR THE EXPLORATION OF TITAN
A. Abstract

There have been various proposals for a successor to the Huygens lander to further
explore the Earth-like nature of Saturn’s moon Titan. Each proposed mission attempts to further
understand Titan’s methane cycle, its nitrogen rich atmosphere and hydrocarbon lakes, which are
the Solar Systems’ only observed liquid bodies outside of Earth. The ideal science target of Titan
is its lake, Kraken Mare. The aforementioned missions each recommend various mid to large
mass exploration spacecraft (> 100 kg) to accomplish this task. The following study proposes a
low mass and cost effective alternative, the quadcopter. This modified quadcopter concept will
be able to survive and operate in the Saturnian moon’s atmospheric and surface environment.
The study provides an in-depth model and analysis of the entry, descent, and landing of the
mission architecture. It also outlines the scientific objectives and instrumentation that the probe
will carry as a payload to Titan and deploy for experimentation and measurement quantification.
B. Introduction

This mission design is conceived and designed based upon mission architecture for the
exploration of Titan using a swarm of amphibious quadcopters and focused predominantly on the
entry, decent and landing (EDL). The architecture will cover liquid lake and sea based science
objectives due to the amphibious nature of the quadcopter swarm. The telecommunications,
power budgets, mass budgets, and the payload for the mission are also discussed.
C. Science Objectives

The science objectives, investigations / experiments and corresponding instrumentation
were chosen from the list of necessary science objectives in the “Planetary Science Decadal
Survey JPL Team X Titan Lake Probe Study Final Report” [19]. These objectives were crossed
referenced with available payloads for the quadcopter design (see Landing section for details) to
narrow down the selection of objectives possible for the mission parameters. This narrowed
down list was then finalized by choosing the objectives that were most associated with the
exploration of Titan’s seas and liquid content. The final science objectives and corresponding
instrumentation can be found in Table 24 (next page).

120
Objective Investigation Instrumentation
Characterize the
amount of liquid on
Titan surface
Determine depth of lake at landing site Echo Sounder
Determine the surface area of the lake DISR (Descent Imager / Spectral
Radiometer) Cameras
Quantify total major organic inventory
present in lakes and seas
FTIR Spectrometer, Turbidimeter,
LPP, Hi-Res GCMS
Characterize major
processes
transforming the
surface throughout
time
Characterize origin of major surface
features, including effects of liquid
flow, tectonic, volcanic, and impact
events.
DISR (Descent Imager / Spectral
Radiometer) Cameras
Determine the
existence of sub-
surface liquid water
ocean and whether
Titan has an intrinsic
magnetic field
Determine depth of lake at landing site
over Titan’s 16-day orbit.
Echo Sounder, Turbidimeter, LPP

Determine induced magnetic field
signatures to confirm subsurface liquid
and place constraints on conductivity
and depth of liquid
Magnetometer
Table 24: Science objectives, investigation and instrumentation [19]
LPP – Lake Properties Package
Hi-Res GCMS – High Resolution Gas Chromatography Mass Spectrometer
The objective in the last row of Table 24 above is the one science objective that could be
fully completed in the CONOPS of this mission architecture. The other two science objectives
were only partially completed due to the lack of available payload mass in the quadcopter design
(required instruments to complete these objectives are in blue).
D. Mission Architecture

The objective of this mission is to explore the Titan’s Kraken Sea, a liquid methane sea
that exists due to extremely cold surface temperatures (~ 94 K). The proposed concept is for a
swarm of amphibious quadcopters to float on the surface of the liquid and hover in the air, to
conduct scientific experiments in each realm. The swarm will consist of 5 daughter quadcopters
and 1 mother quadcopter. Each quadcopter will carry a scientific instrument and the mother
quadcopter will only house a Radioisotope ThermoPhotoVoltaic power source to serve as a
charging station for the daughter quadcopters. The swarm will fly as a unit and will use their
designed buoyancy property to float on the lake surface while charging their batteries.

121
The entry, descent and landing (EDL) architecture of the mission is designed to enter the
atmosphere of Titan from an orbiting platform in a ballistic nature. The reasons for choosing a
ballistic entry and orbital entry are listed below:
1. The area of the landing ellipse over Kraken Sea is very large (400 – 800 km).
2. The primary goal of the mission is to implement a low-cost mission and since a lifting body
entry has a higher overall heat load, the structural aspect of the thermal protection system
requirement becomes more robust and complex hence increasing the mission cost.
3. No Guidance, Navigation and Control required for the vehicle since the required landing
ellipse is very large which reduces the complexity of the EDL mission.
4. The quadcopters are deployed at very high altitude (60 km) and hence can fly to a more
precise landing point over the Kraken Sea.

Figure 88: CAD model of the entry vehicle
Figure 88 above shows the CAD model of the entry vehicle, which will deliver the
quadcopter swarm into Titan’s atmosphere. The entry vehicle is dropped from a 1000 km
circular orbit and enters the hypersonic phase around 180 km. The hypersonic regime lasts for a
drop of 140 km and the regression rate of the Thermal Protection System (TPS) material is as big
as 6 mm / km at the earliest part of the regime. Thus achieving a shorter duration in the
hypersonic phase enables for a thinner heat shield and lighter TPS layer reducing the entry mass
and mission cost. When the Mach number reaches 1.8, a supersonic drogue parachute is
deployed to carry the vehicle through the supersonic phase, which lasts for about 32 seconds. At
Mach number of about 0.1, when the entry vehicle has achieved the terminal velocity, the
subsonic / landing sequence of the EDL is initiated by separating the TPS system and the
parachute from the entry vehicle. After the separation of the TPS from the entry vehicle, the 6
quadcopters are deployed. The swarm drops freely and takes approximately 300 seconds to reach
the surface. The entire EDL process is achieved in roughly 10 minutes.

122

Figure 89: EDL Concept of Operations (CONOPS)
E. Entry, Descent and Landing

i. Input
Parameter
Methodology

To properly size the system as well as determine the entry conditions, a Monte Carlo
Simulation model was designed to generate random combinations of the selected input variables
within the set boundary conditions of each parameter. Each of these combinations then went
through the EDL calculations to produce a very wide spread distribution of performance
parameters. In order to select the proper input parameters, the hypersonic and supersonic
governing equations were analyzed to determine which input parameters were top level variables
that were not based upon any other inputs. These variables are entry velocity (V
e
), entry altitude
(h
e
), entry mass (m
e
), Flight Path Angle (FPA), Solid Cone Angle of the entry vehicle (SCA) and
the diameter of the heat shield (D
HS
).

123
The boundary conditions used for the input parameters in the Monte Carlo Simulation
model are listed in Table 25 below.
Parameter V
e
(km/s) h
e
(km) m
e
(kg) FPA SCA D
HS
(m)
Boundary Condition 4.5 – 13.5 80 – 200 150 – 200 -80
0
à -6
0
20
0
– 70
0
0.5 – 1.5
Table 25: Boundary conditions set for the input parameters to the Monte Carlo Simulation.

Figure 90: Monte Carlo Simulation Results showing Altitude (km) v/s Peak Heat Flux (W/cm
2
)

Figure 91: Monte Carlo Simulation Results showing Altitude (km) v/s Peak Deceleration (g)

124
The Monte Carlo Simulation model ran 100,000 different combinations of these
variables. The results from the Simulation were then narrowed down to one optimized set of
input parameters. The optimization of the heat shield sizing and entry conditions were
determined by satisfying both the deceleration and heating limits of the EDL vehicle. The
optimum values of the entry parameters are listed in Table 26 below.
Parameter V
e
(km/s) h
e
(km) m
e
(kg) FPA SCA D
HS
(m)
Optimum Values 4.5 190 350 -55
0
63.8
0
1.48
Table 26: Optimum values of entry parameters obtained from the Monte Carlo model.
ii. Hypersonic
Entry

The entry, descent and landing (EDL) phase of the exploration mission to Titan begins
with the hypersonic entry of the aero shell into Titan's atmosphere after being released from the
orbiter. During this portion of the descent, the aero shell is traveling at speeds > 5 km/s and is
where the majority of the energy dissipation occurs during the EDL phase (> 95%) of the
mission. Large amounts of heating occur so the thermal protection system must be designed
accordingly to handle larger deceleration and heating loads.

A ballistic and orbital entry was chosen as the entry method for the aero shell. A ballistic
entry was chosen because of the large landing ellipse, reduction of total heat load with shorter
EDL time and the high altitude deployment of the quadcopters (> 50 km). The high altitude
deployment of the quadcopters will allow for correction to the deviations from the initial flight
path. An orbital entry was chosen because the exploration mission architecture includes an
orbiter, which allows for pre-assessment of atmosphere conditions on Titan before EDL is
initiated. It also reduces any interplanetary trajectory errors that would creep into EDL process.

For hypersonic descent, the governing equations for a ballistic entry are derived from
Newtonian Aerodynamics. This makes the assumption that after the particles impinge the
surface, the normal momentum is lost while the tangential momentum is conserved, an
assumption only valid at hypersonic speeds [20]. The equations that explain the trajectory are as
follows:
ℎ =
!,!"#$%

!!!
à Equation 35
Where,
à Density, A à Scale height and h à Altitude

125
ℎ =
!,!"#$%
!
!"#$%
!
!"#$%
!!
!
à Equation 36
Where,
g à acceleration due to gravity and R à Radius of the target planet / moon
ℎ =
!

!!
!!!
à Equation 37
Where,
v à Velocity and v
e
à Entry Velocity
=
!(!)
!!"#$%(!"#)
à Equation 38
Where,
β à Ballistic co-efficient and FPA à Flight path angle
=
!
!
!
!"#
!
!"
à Equation 39
Where,
C
dEV
à Drag co-efficient of entry vehicle
ℎ =
!!
!
!
! !
ℎ
!!!

!"
!!!
à Equation 40
Where,
n à Deceleration
∆= ∆ℎà Equation 41
Where,
Δs à Downrange

!"#
=
!
!
!
!"#$(!"#)
!!
!,!"#$%
!
à Equation 42

ℎ
!,!"#
=
!
!
ln (−2) à Equation 43

!,!"#
= 0.606
!
à Equation 44

During the hypersonic phase, the large heat load must be quantified in order to
adequately select the thermal protection system. This is done using the following equations:
"(ℎ)=
!"#$%
!(!)
!
!
(ℎ)
!
à Equation 45
Where,
q" à Heat flux and R
n
à Nose radius of the heat shield

126

!"#$%
=
!"#$%

!
!
!"
!
!,!"#$%
!!
!
!"# (!"#)
à Equation 46

"
!"#
=
!"#$%
!(!
!",!"#
)
!
!

!,!"#
!
à Equation 47

ℎ"
!,!"#
= −
!
!

!"#$%(!"#)
!!
!,!"#$%
à Equation 48

!",!"#
= 0.846
!
à Equation 49

For the EDL into Titan, Figure 92 and Figure 93 respectively, are based upon the
optimum trajectory and thermal resistance. Figure 93 also shows the TPS regression rate as a
function of altitude. In order to ensure adequate thickness for the TPS design, empirical data was
considered from the journal paper – “Post flight Thermal Protection System Analysis of
Hayabusa Reentry Capsule” published in AIAA Journal of Spacecraft and Rockets [58].
Especially for the nitridation that occurs on the TPS surface during the hypersonic entry phase in
an Earth-like atmosphere.
Figure 92: Trajectory plots based on optimum input parameters.

127
Figure 93: Heating plots based on optimum input parameters

Parameter Max h
max
(km) v
max
(m/s)
Deceleration, n -12.5 g 70 350
Heat flux, q” 66 (W/cm
2
) 170 4400
Table 27: Maximum deceleration and heat values.
iii. Supersonic
Descent

The supersonic portion of the EDL is dominated by the characteristics of the parachute
used in the supersonic descent. The supersonic DGB parachute used for this mission is the same
for the one used in the Viking mission. The change of the drag-coefficient of the parachute was
taken into consideration while designing the supersonic phase. Figure 94 in the next page shows
a plot of the Mach No. v/s drag coefficient of the supersonic parachute [59].
Parachute
Type Heritage Opening Load Factor Diameter (m)
DGB Viking 1.1 2.2
Supersonic Descent Parameters
Deployment Mach No. Altitude @ Deployment Terminal Velocity Peak Load
1.8 54.5 km 36.9 m/s 6827 N
Table 28: Parachute and Supersonic Descent Parameters.

128
Figure 94: Mach No. v/s drag coefficient of the supersonic parachute
Figure 95 shows the supersonic phase of the descent trajectory, specifically the velocity,
deceleration, and Mach No. versus altitude as well as the altitude with respect to time.
Figure 95: Supersonic trajectory plots

129
iv. Landing

At 60 km the supersonic phase will reach to an end, when the entry vehicle reaches
terminal velocity of less than 50 m/s. At this altitude the density of the atmosphere is thick
enough for the quadcopters to operate. The density at this altitude is roughly 1/3
rd
of the density
of Earth at its surface. But this density rapidly increases, as the quadcopters get closer to the
surface. Two major advantages of flying the quadcopters in Titan’s atmosphere are: 1) Titan’s
surface density is approximately 5.4 kg/m
3
, which is almost five times the density at the surface
of Earth and 2) the acceleration due to gravity is 1/10
th
that of Earth. Figure 96 shows the density
variation v/s altitude at the deployment point for the swarm from the entry vehicle. From the
figure it can be see that there is a rapid increase in the density after 100 km altitude [60].

Figure 96: Altitude (km) v/s Density (kg/m
3
) at Titan
Figure 97 below shows the schematic CAD models of the daughter quadcopter design.

Figure 97: CAD models of the quadcopter

130

Figure 98: Power required by the quadcopters to hover.
Figure 98 shows a plot of the altitude v/s power required to hover the daughter
quadcopters at a given altitude. The rotor blade radius of each rotor of the quadcopter R = 15 cm
which results in the power required to hover the quadcopter [62]. Figure 98 also shows it takes
approximately 12.5 W to fly the copter at 60 km deployment altitude and this equates to about 28
W for the mother quadcopter. Each quadcopter has been fitted with a 50 Wh battery (see
Quadcopter Power Budget section for more details), so they have more than substantial power to
travel from deployment to Titan’s surface. Figure 99 (next page) shows the schematic flying
control system model that will be adapted for each quadcopter will in flight. This is also a
standard representation of a PID controller for flying any quadcopter model. This control system
architecture is a constituent of the overall control system architecture for the swarm flying
operations [62][63].

Figure 99: PID control system for quadcopters [62]

131
F. Power Source

The central mother copter carrying the RTPV is the main power source for the daughter
copters. Table 6 shows the mass and volume parameters for potential radioisotope fuels that can
be used in the RTPV for powering the swarm. Curium (Cm) has a very high power density to
mass but it has a very small half-life and has a lot of gamma radiation. Americium (Am) is not as
good as Plutonium dioxide but the limited supply of Plutonium is definitely a big draw back for
its candidacy. However for the mission a
238
PuO
2
in Tungsten Rhenium cermet was considered
for the mother copter’s RTPV under a fair assumption that the United States will be investing in
projects trying to manufacture Plutonium again.
Curium (Cm)
Power (W
e
) Mass (g) Volume (cm
3
) Power Density (W
e
/g)
5 82.7 123.4 0.06
10 109.6 145.3 0.09
15 134.5 162.1 0.11
Americium (Am)
Power (W
e
) Mass (g) Volume (cm
3
) Power Density (W
e
/g)
5 731.7 393.5 0.06
10 1362.86 554.9 0.09
15 1986 691.2 0.11
Plutonium (Pu)
Power (W
e
) Mass (g) Volume (cm
3
) Power Density (W
e
/g)
5 249.4 226.8 0.02
10 423 296.5 0.0236
Table 29: Radioisotope fuel candidates for the power source.

Figure 100: CAD model of the RTPV

132
G. EDL Telecommunication

Noise, antenna size, propagation path length, signal to noise ratio, modulation type, data
rate and transmission power are few of the design parameters that needs to be considered while
designing communication link budget. During EDL, the communication links are established by
communicating to the orbiting satellite where the information is collected, stored and later
transmitted to earth using a laser communication system discussed earlier. Use of a Direct To
Earth (DTE) link allows us to receive information much earlier than a relaying spacecraft and
also reduces the cost of the link, however it is associated with risks of a low SNR and the need
for line of sight, higher power and complex modulation schemes. For the Huygens probe, radio
telescopes were required in order to receive the highly attenuated DTE signal at DSN [62].
An advantage of using DTE link to send carrier signals is that it allows us to determine
the status of the entry vehicle as it happens through the Doppler signature and power levels of
the signal. The ability to track rapidly changing frequency and power levels and their thresholds
has been demonstrated earlier with the MSL and MER missions. Using the Doppler signals,
critical events during the EDL such as parachute deployment, heat shield separation, payload
deployment can be acquired. This would help in understanding the performance of the EDL
system, and in the advent of a failure, this data may be the key to determine the cause [63]. The
telecom between the quadcopters plays a major role in maintaining the flight system as a swarm
and maneuvering as a unit.
A simple Link Budget calculation was carried out considering the major factors affecting
the link. The link budget was designed to the maximum distance between Titan and Earth. The
major issue with respect to designing a DTE link for a Titan mission is the space loss factor,
which is a function of the distance between the transmitter and the receiver. Major issue to
consider while incorporating these antennae on the entry vehicle is the type and thickness of the
TPS material since this can cause significant losses on the link. Another important parameter is
the possibility of a plasma blackout, which is dependent on the frequency of transmission,
velocity of the entry vehicle and the density of the atmosphere. The EDL link budget is shown in
Table 30 (next page).

133
Link DTE Entry Vehicle to Orbiter
T
x
Antenna Patch Patch
T
x
Power (W) 100 0.24
Frequency (GHz) 8.4 (X-Band) 0.45 (UHF Band)
Data type Tones Critical Data
Data rate (kbps) Carrier Signal 4
EIRP (dB mW) 56 29.97
Distance (km) 1.6 x 10
9
1300
R
x
Antenna 70 m, DSN 1 m, LGA
SNR 1.78 24.9
Margin 1.78 15.9
Table 30: EDL Link Budget
H. Payload & Thermal Control

The list of science objectives and their corresponding instruments to fulfill the objectives
were taken from Waite’s Titan decadal survey [56]. The echo sounder and the Turbidmeter were
both based on the surface science package on the Huygens probe. The echo sounder would
measure the depth of the lake and create a profile of the lake while the Turbidmeter would
measure the different densities of the Titan lake inside the submarine [64]. The descent and
surface cameras have their heritage from the MER cameras, but are currently based off of 1U
CubeSat cameras for the sake to reduce weight. The cameras have a 9.22
0
field of view, which is
used on the quadcopters [41]. The Magnetometer was also based on CubeSat designs, being used
to measure the magnetic field strengths on titan [65]. The IR spectrometer was based on the
Fourier transform infrared spectrometer that was proposed to go on the MSL mission. It would
take measurements of the lake composition inside the submarine [66]. The subsystem block
diagram is shown in Figure 101 (next page).
Radioisotope Heater Units (RHUs) were used as the primary heat source due to their
plentiful and non- electrical powered heat. In order to size how many heater units were required,
a simple heat transfer calculation was performed. The heat from the RHUs will escape the
vehicle by conduction through the insulating material, and then radiation from the vehicle wall
into the ambient atmosphere.

134
Figure 101: Subsystem block diagram
The figure above shows the Kraken
Sea landing site on the surface of Titan and the
landing ellipse. The landing site chosen is 72
0

N 310
0
W, Kraken Mare. The downrange
created in the hypersonic phase is < 450 km
and hence from a 60 km deployment altitude, a
considerable amount of this downrange can be
re-adjusted. The entire duration of the
quadcopter descent to the surface lake will not
take more than 10 minutes, which is well
within the flying capability of the quadcopter
swarm based on their power source.
Figure 102: Landing Ellipse [67]

135
III. REFERENCES

[1] Hamid Hemmati (JPL), Editor. Deep Space Optical Communications – Equation 5.4-1 &
5.4-2, Figs 2.2, 2.1, 1.1, 1.2, 2.8 & 3.38 and Table 2.1.
[2] Ametek Materials Analysis Division, X-123CdTe Complete X-Ray & Gamma Ray
Spectrometer, URL: http://www.amptek.com/products/x-123-cdte-complete-x-ray-gamma-
ray-spectrometer-with-cdte-detector/ [cited 27
th
July 2014]
[3] Thoth Technology Inc., Argus 1000 IR Spectrometer – Owner’s Manual, Issue Release 1.03,
Document Number OG728001. URL: http://thothx.com/manuals/Argus%20Owner's
%20Manual,%20Thoth%20Technology,%20Oct%2010,%20rel%201_03.pdf [cited 2
nd

August 2014]
[4] CubeSatShop.com, NanoCam C1U, URL: http://www.cubesatshop.com/index.php?page=
shop.product_details&product_id=63&flypage=flypage.tpl&pop=0&option=com_virtuemart
&Itemid=65 [cited 2
nd
August 2014]
[5] Herrero A. F., et. al., The Gated Electrostatic Mass Spectrometer (GEMS): Definition and
Preliminary Results, J Am Soc Mass Spectrom 2008, 19, 1384 –1394.
[6] Esposito M., et. al., A highly integrated micropayload for broadband infrared spectrometry
(HIBRIS), Proc. SPIE 7808, Infrared Remote Sensing and Instrumentation XVIII, 780816
(August 27, 2010); doi:10.1117/12.863728.
[7] GomSpace Profile, LAMBDA – Laser based instrument for the ExoMars mission, URL:
http://gomspace.com/index.php?p=profile-references [cited 2
nd
August 2014]
[8] Spencer J., Niebur C., Mission Concept Study – Planetary Science Decadal Survey Jupiter
Europa Orbiter Component of EJSM, Table 3-1.
[9] Turtle E., Niebur C., Mission Concept Study – Planetary Science Decadal Survey Io
Observer, Tables 3-1, 3-2, 3-3, 3-4 and 3-5.
[10] Khurana K., Niebur C., Mission Concept Study – Planetary Science Decadal Survey
Ganymede Orbiter, Tables 3-1 – 3-12.
[11] Hubbard B. W., Mission Concept Study – Ice Giants Decadal Study Uranus Orbiter,
Tables 3-1 – 3-10.
[12] Yeomans K. D., Planetary Satellite Physical Parameters, JPL Solar System Dynamics.
[13] Campagnola S., et al., Jovian Tour Design For Orbiter and Lander Missions to Europa,
23
rd
AAS / AIAA Space Flight Mechanics Meeting, Kauai, HI, 2013.

136
[14] Evans W. R., et. al., GRID 2: A Program for Rapid Estimation of the Jovian Radiation
Environment, Europa Clipper Pre-Project, JPL Publication 14 – 8, January 2014, URL:
https://solarsystem.nasa.gov/europa/docs/Grid2_JPL_Pub_14-8.pdf [cited 2
nd
August 2014].
[15] Aviv D., Laser Space Communications, Fig 3.11, Equations 3.11, Fig 2.10, 2.11, 2.25 &
2.26 and Table 7.2.
[16] Wavelengths of commercially available lasers, Weber J. M., Handbook of laser
wavelengths, CRC Press, 1999. URL: http://upload.wikimedia.org/wikipedia/commons/
thumb/4/48/Commercial_laser_lines.svg/800px-Commercial_laser_lines.svg.png.
[17] Pribil K., Hemmati H., Near Earth Laser Communications, 4 – Laser Transmitters:
Coherent and Direct Detections.
[18] Walter R. L., Peter J. W., Near Earth Laser Communications, 7 – Photodetectors and
Receivers, Figures 7.6, 7.8 & 7.11 and Equation 7.22
[19] Hunter W., Niebur C., Planetary Science Decadal Survey JPL Team X Titan Lake Probe
Study Final Report, 2010.
[20] Sengupta A., Newtonian Aerodynamics, ASTE 599: Entry, Descent and Landing for
Planetary Surface Exploration, Lecture 4.
[21] Suzuki, T., Fujita, K. Post flight Thermal Protection System Analysis of Hayabusa
Reentry Capsule, AIAA Journal of Spacecraft and Rockets. Vol. 51, No. 1, pg. 97-99.
[22] Sengupta A., Mach No. v/s Drag Coefficient, ASTE 599: Entry, Descent and Landing for
Planetary Surface Exploration. Anita Sengupta, Lecture 4.
[23] R. V. Yelle, et. al., Engineering Models for Titan's Atmosphere, 1994.
[24] Pierre-Jean Bristeau, et. al., The Role of Propeller Aerodynamics in the Model of a
Quadrotor UAV, Proceedings of the European Control Conference 2009, Budapest, Hungary,
August 23– 26, 2009.
[25] Asmar S., Huygens Wind Doppler Experiment, JPL, June 2005.
[26] Lombardi M., Communications Test Tools: A Study of the Cassini-Huygens Mission, RT
Logic, and Integral Systems Company, May 2010.
[27] Zarnecki, J.C et al., The Huygens Surface Science Package.
[28] “Magnetometer”, CubeSatShop.com, 2014. URL: http://www.cubesatshop.com
/index.php?page=shop.product_details&flypage=flypage.tpl&product_id=90&category_id=7
&option=com_virtuemart&Itemid=69 [Cited 21
st
April 2014]

137
[29] Anderson, Mark S. Fourier transforms infrared spectroscopy for Mars science, Review
of Scientific Instruments. No. 76, 2005.
[30] Strange, N., et al., Mission Design for the Titan Saturn System Mission Concept, AAS-
09-356.
IV. BIBLOGRAPHY

[1] Abid M. M., Spacecraft Sensors, Wiley – ISBN 978-0-470-86527-9, August 2005.
[2] Nieminen J., Weatherford J., Rajguru A., Thermodynamic Analysis and Radiator Design of a
Pulsed Bi-modal Radioisotope Propulsion System, Paper 5092, Proceedings of Nuclear &
Emerging Technologies for Space (NETS 2015), Albuquerque, NM, February 23 – 26, 2015,
URL: http://anstd.ans.org/wp-content/uploads/2015/07/Proceedings-of-NETS-
2015.pdf.
[3] Faler A., Rajguru A., Operations Cost Reduction of a Jovian Mission using CubeSats,
SpaceOps 2014 Conference, 5 – 9 May 2014, Pasadena, CA, AIAA 2014-1681, URL:
http://arc.aiaa.org/doi/pdf/10.2514/6.2014-1681.
V. ACKNOWLEDGEMENTS

[1] Dr. Steve Howe – Advisement on all chapters
[2] Mr. Nathan Jerred – Advisement and support on chapter 1
[3] Dr. Troy Howe – Advisement and support on chapter 1
[4] Mr. Juha Nieminen – Support in chapter 1
[5] Mr. Justin Weatherford – Support in chapter 1
[6] Mr. Joseph Santora – Support in chapter 2
[7] Pete Husemeyer
[8] Mr. Vishal Patel
[9] Dr. Mohammed Abid – Support in chapter 2
[10] Mr. Bryan Park Franz – Support in chapter 3

138
APPENDICES

Figure 103:
244
Cm
2
O
3
core configurations with Beryllium as the thermal capacitor.

Core
Type
Diameter of
fuel rods (cm)
Number of
fuel rods
Fx
(cm)
Px
(cm)
Flow Channel
Diameter (cm)
Number of flow channels
1.0 1.5 13 1.5 1 .5 280
2.0 0.9 41 0.8 1 .5 258
3.0 2.1 7 1.7 1 .5 282
4.0 1.8 9 1.7 1 .5 282
5.1 0.7 17 0.8 1 .5 280
5.2 1.1 6 1.1 1 .5 280
5.3 1.5 6 1.2 1 .5 280
6.0 1.5 14 1.3 1 .5 281

Table 31: Specifications for
238
PuO
2
core with Silicon as the thermal capacitor.

Core
Type
Diameter of
fuel rods (cm)
Number of
fuel rods
Fx
(cm)
Px
(cm)
Flow Channel
Diameter (cm)
Number of flow channels
1.0 2.9 13 1.8 1 .5 288
2.0 1.7 41 1.15 1 .5 212
3.0 4.1 7 2.6 1 .5 264
4.0 3.6 9 1.5 1 .5 236
5.1 1.3 17 1.0 1 .5 212
5.2 1.9 6 1.5 1 .5 212
5.3 2.6 6 1.7 1 .5 212
6.0 2.9 14 1.8 1 .5 299

Table 32: Specifications for
241
AmO
2
core with Silicon as the thermal capacitor.

139
Core
Type
Diameter of
fuel rods (cm)
Number of
fuel rods
Fx
(cm)
Px
(cm)
Flow Channel
Diameter (cm)
Number of flow channels
1.0 0.6 13 1.2 1 1 72
2.0 0.3 41 0.8 1 1 72
3.0 0.8 7 2.0 1 1 72
4.0 0.7 9 1.8 1 1 74
5.1 0.3 17 0.9 1 1 72
5.2 0.4 6 0.9 1 1 72
5.3 0.5 6 1.5 1 1 72
6.0 0.6 14 1.7 1 1 75

Table 33: Specifications for
244
Cm
2
O
3
core with Beryllium as the thermal capacitor.

Neutrons (spontaneous fission) 4.2 x 10
6
n/(s*w)
Energy Output 3 w/g
Neutron Production 1.3 x 10
7
n/(s*g)
Density of Curium 13.5 g/cm
3
Volume of Curium 350 cm
3
Weight of Curium 4,732 g
Neutron Production Total 5.95 x 10
10
n/s
Table 34:
244
Cm Spontaneous Fission Data.

Figure 104: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 1.
0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2
0
500
1000
1500
2000
2500
3000
3500
4000
4500
Energy (MeV)
neutron count (neutrons / second)
Neutron Counts vs Energy With Varying Distance from Asteroid

2 meters
3 meters
4 meters
5 meters
6 meters

140

Figure 105: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 2.

Figure 106: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 3.
0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2
0
1
2
3
4
5
6
7
x 10
4
Energy (MeV)
neutron count (neutrons / second)
Neutron Counts vs Energy With Varying Distance from Asteroid

3 meters
4 meters
5 meters
6 meters
7 meters
0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 2.2 2.4
0
1
2
3
4
5
6
7
8
9
10
x 10
4
Energy (MeV)
neutron count (neutrons / second)
Neutron Counts vs Energy With Varying Distance from Asteroid

4 meters
5 meters
6 meters
7 meters
8 meters

141

Figure 107: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 4.
Figure 108: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 5.
0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2
0
1
2
3
4
5
6
7
8
x 10
4
Energy (MeV)
neutron count (neutrons / second)
Neutron Counts vs Energy With Varying Distance from Asteroid

4 meters
5 meters
6 meters
7 meters
8 meters
9 meters
0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2
0
1
2
3
4
5
6
7
8
x 10
4
Energy (MeV)
neutron count (neutrons / second)
Neutron Counts vs Energy With Varying Distance from Asteroid

4 meters
5 meters
6 meters
7 meters
8 meters
9 meters

142

Figure 109: Neutron Counts v/s Energy with varying distance from the simulated hydrated C-
Class Asteroid for Core Configuration 6.

Figure 110: Circuit Diagram of Pacific Monolithics PM2105 Power Amplifier
The resistor and capacitor values of the power amplifier shown in the above figure are listed in
Table 35 (next page).
0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2
0
1
2
3
4
5
6
7
8
x 10
4
Energy (MeV)
neutron count (neutrons / second)
Neutron Counts vs Energy With Varying Distance from Asteroid

3 meters
4 meters
5 meters
6 meters
7 meters
8 meters
9 meters
10 meters

143
Part Value
C1, C7, C10, C11, C12, C13 and C16 33 pF
C2 6.8 pF
C3 and C4 1.5 pF
C18 6.8

F
C6, C9, C14 and C17 1000 pF

C5, C8 and C15 0.1

F

R1 and R2 51 ohms

R3 3.3 ohms

L1 39 nH

Table 35: The resistor and capacitor values of Pacific Monolithics PM2105 Power Amplifier.

Figure 111: Illustration of the pinger device’s BPSK modulation scheme

Image Size
(m)/pixel
Orbiting
Altitude (km)
Image size on ground (m)
X Y
Image Area
(m
2
)
Pictures
Required
Transmittal
Time (hrs)
80 650 163840 122880 20132659200 1535 1.60
60 487.5 122880 92160 11324620800 2729 2.84
40 325 81920 61440 5033164800 6139 6.39
20 162.5 40960 30720 1258291200 24557 25.58
15 121.875

30720 23040 707788800 43657 45.48
10 81.25

20480 15360 314572800 98228 102.32
6.1 49.5625

12492.8 9369.6 117052538.9 263984 274.98
5 40.625

10240 7680 78643200 392914 409.29
4 32.5

8192 6144 50331648 613928 639.51
3 24.375

6144 4608 28311552 1091427 1136.90
2 16.25

4096 3072 12582912 2455711 2558.03
1 8.125

2048 1536 3145728 9822845 10232.13
0.5 4.0625

1024 768 786432 39291382 40928.52
0.25 2.03125

512 384 196608 157165527 163714.09
Table 36: Tabulation for obtaining the optimal parking orbit after rendezvous with target body.

144

Figure 112: Data downlink transmittal time (hours) v/s picture count
Event Date V
∞

(km/s)

Altitude (km)
Earth Escape 22
nd
November 2021 3.77 200
Venus flyby 14
th
May 2022 8.77 469
Earth flyby 24
th
October 2023 8.05 655
Earth flyby 24
th
October 2025 8.22 1910
Jupiter Orbit Insertion 3
rd
April 2028

- -
ΔV required 3.521
Ganymede flyby 0 3
rd
April 2028 7.9 500
Ganymede flyby 1 18
th
November 2028 3.5 684
Ganymede flyby 2 5
th
February 2029 4.9 252
Ganymede flyby 3 3
rd
April 2029 5.1 1552
Ganymede flyby 4 9
th
May 2029 5.7 643
Callisto flyby 5 17
th
June 2029 5.4 221
Ganymede flyby 6 21
st
July 2029 4.2 117
Ganymede flyby 7 2
nd
September 2029 3.7 469
Ganymede flyby 8 23
rd
September 2029 3.7 3821
Ganymede flyby 9 8
th
October 2029 3.7 170
Callisto flyby 10 23
rd
October 2029 1.7 1148
Ganymede flyby 11 27
th
October 2029 11.7 421
Ganymede flyby 12 24
th
November 2029 2.6 6201
Europa flyby 14 27
th
November 2029 3.8 6681
Europa Orbit Insertion 15 11
th
December 2029 6.7 6800
ΔV required 1.136
Table 37: Earth-Jupiter Cruise Phase Fly-bys & Jovian Moons Fly-bys [Chapter 3 – 13]

Abstract (if available)

Abstract Water can be electrolyzed into LOX and LH₂ or it can directly be used as a propellant in a nuclear propulsion system. There is abundant evidence that water is present in Carbonaceous Chondrites, which forms the C-Class asteroids in the main asteroid belt. There is a strong belief that water is also available in the form of regolith ice and chemically bound sources. Water is vital for manned space programs as it can be used for drinking water, agriculture and radiation shielding. Approximately 90% of this water can be carried from Earth and recycled within a life support system. But for a very long duration human presence in missions beyond the asteroid belt, refueling water from the main asteroid belt for sustainability can be an enabling aspect of such a mission. Hence the value of information for water maps on the main asteroid belt can be a very lucrative market for futuristic space missions. ❧ With NASA’s renewed interest in exploring the gas giants and their moons, cost reduction is on the forefront of concern. Concurrently, the surge in successful CubeSat launches to Near-Earth Orbit, has spawned interest in examining their use for deep space applications. In current CubeSat flight heritage, the communication systems use RF signals on lower bands, milli Watt class of power and compact omnidirectional Antenna system. But in order to close the communication link for distances greater than 10 AU, kW class of power is required to overcome the >250 dB space loss. Subsequently, the antenna diameter size goes into the order of a couple of meters. This results in the communication system to be bulky and spacious, increasing the cost of launch and operations into a billion-dollar tier mission. The study discussed in this chapter targets the reduction of this generic RF communication system’s mass and size, by replacing it with laser communication technology, which can fit in a 6U CubeSat constraint. Achieving this miniaturization can lower the cost of deep-space missions to the order of a million-dollar tier making it more accessible to small budget organizations such as university research labs. ❧ There have been various proposals for a successor to the Huygens lander to further explore the Earth-like nature of Saturn’s moon Titan. Each proposed mission attempts to further understand Titan’s methane cycle, its nitrogen rich atmosphere and hydrocarbon lakes, which are the Solar Systems’ only observed liquid bodies outside of Earth. The ideal science target of Titan is its lake, Kraken Mare. The aforementioned missions each recommend various mid to large mass exploration spacecraft (> 100 kg) to accomplish this task. The following study proposes a low mass and cost effective alternative, the quadcopter. This modified quadcopter concept will be able to survive and operate in the Saturnian moon’s atmospheric and surface environment. The study provides an in-depth model and analysis of the entry, descent, and landing of the mission architecture. It also outlines the scientific objectives and instrumentation that the probe will carry as a payload to Titan and deploy for experimentation and measurement quantification.

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Creator Rajguru, Adarsh (author)

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Publication Date 12/16/2015

Defense Date 12/16/2015

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Tag Americium,asteroid,asteroid mining,Curium,entry decent and landing,Europa Orbiter,ground penetrating radar,heat transfer,laser communication,mission architecture,OAI-PMH Harvest,Plutonium,Power,propulsion,quadcopter,radioisotope,radioisotope thermal rocket,radioisotope thermophotovoltaic power source

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