# Abstract

The Mars Subsurface Chemical Life Explorer (MSCLE), will take the next
steps toward our nearest terrestrial neighbor, and the next steps in the
NASA exploration vision. Critical for our understanding of life on Earth
and extending it beyond lunar orbit, the search for present Martian life
and habitability will rely largely on chemical analysis of soil from up
to six meters below the surface. Several experiments span a range of
possibilities for discerning biological signatures: molecular
handedness, metabolic process, and isotopic ratio abundances are among
the factors studied.

A rotary percussion auger drill will obtain the necessary depths and
help retrieve the samples to a scientific experimentation environment
contained within the stationary lander. As a PI-class Scout mission,
MSCLE will be ready for the 2011-launch proposal cycle. It is currently
well within budget, and the Public enrichment team has already begun
setting outreach and education gears in motion. With the guidance of Dr.
Herb Frey at Goddard Space Flight Center, MSCLE has grown incredibly
from its inception through the end of this preliminary design phase.

# Table of Contents

[1. Abstract 1](#abstract)

[2. Table of Contents 2](#table-of-contents)

[3. List of Figures 3](#list-of-figures)

[4. List of Tables 4](#list-of-tables)

[5. Executive Summary 6](#executive-summary)

[6. Science Operations and Development
8](#science-operations-and-development)

[6.1. Science Background 8](#science-background)

[6.1.1. History 8](#history)

[6.1.2. Prospects for Martian Life 9](#prospects-for-martian-life)

[6.1.3. Landing Site 11](#landing-site)

[6.2. Primary Objective: Search for Present Life
17](#primary-objective-search-for-present-life)

[6.2.1. Stage 1: Characterizing Chemical Composition of Soil and Rock
Constituents
17](#stage-1-characterizing-chemical-composition-of-soil-and-rock-constituents)

[6.2.2. Stage II: Search for Distinctive Biological Signatures
20](#stage-ii-search-for-distinctive-biological-signatures)

[6.3. Secondary Objective: Measure Atmospheric conditions
22](#secondary-objective-measure-atmospheric-conditions)

[6.3.1. Experiment 22](#experiment-2)

[6.3.2. Instrumentation 22](#instrumentation-2)

[6.3.3. Data Analysis 23](#data-analysis)

[6.4. Tertiary Objective: Obtain Photographic Images of the Landing Site
23](#tertiary-objective-obtain-photographic-images-of-the-landing-site)

[6.4.1. Experiment 23](#experiment-3)

[6.4.2. Instrumentation 24](#instrumentation-3)

[6.4.3. Data Analysis 24](#data-analysis-1)

[7. Engineering Challenges and Solutions
25](#engineering-challenges-and-solutions)

[7.1. Overview and Introduction 25](#overview-and-introduction)

[7.2. Driving Requirements 25](#driving-requirements)

[Drill Depth 26](#drill-depth)

[Soil Material 26](#soil-material)

[Sample Acquisition and Processing
26](#sample-acquisition-and-processing)

[Reliability and Safety 26](#reliability-and-safety)

[7.3. Chief Design Trades 26](#chief-design-trades)

[Drill Selection 26](#drill-selection)

[Solar Panel Arrays 26](#solar-panel-arrays)

[Rover versus Lander 27](#rover-versus-lander)

[Descent Style 27](#descent-style)

[7.4. Present Status 27](#present-status)

[7.5. Subsystem Analyses 27](#subsystem-analyses)

[7.5.1. Structures 27](#structures)

[7.5.2. Payload 32](#payload)

[7.5.3. Power Production and Distribution
35](#power-production-and-distribution)

[7.5.4. Flight Dynamics 41](#flight-dynamics)

[7.5.5. Propulsion 43](#propulsion)

[7.5.6. Communications and Data Handling
47](#communications-and-data-handling)

[7.5.7. Ground Support and Operations
52](#ground-support-and-operations)

[7.6. End of Life Analysis 56](#end-of-life-analysis)

[8. Test and Validation 56](#test-and-validation)

[8.1. Objective 56](#objective)

[8.2. Approach 56](#approach)

[8.3. Tests 58](#tests)

[9. Budget 58](#budget-1)

[9.1. Methods 58](#methods)

[9.2. Cost Estimating Procedure 59](#cost-estimating-procedure)

[9.2.1. First Iteration 59](#first-iteration)

[9.2.2. Second Iteration 61](#second-iteration)

[9.2.3. Third Iteration 62](#third-iteration)

[10. Outreach 62](#outreach-1)

[10.1. Educate the Public 63](#educate-the-public)

[10.2. Demonstrate the Benefits 65](#__RefHeading___Toc81659263)

[10.3. Inspire and Excite 66](#inspire-and-excite)

[11. Conclusions 67](#conclusions-1)

[11.1. Next Steps 67](#__RefHeading___Toc81659266)

[11.2. Lessons Learned 67](#__RefHeading___Toc81659267)

[12. Acknowledgements 68](#acknowledgements)

[13. Appendix A 69](#appendix-a)

[14. Appendix B: Organization 70](#appendix-b-organization)

[14.1. Organization Chart 70](#organization-chart)

[14.2. Engineering Team Overview 71](#engineering-team-overview)

[14.1.1. Chief Engineer/Systems Coordination
71](#chief-engineersystems-coordination)

[14.1.2. Structures 71](#structures-1)

[14.1.3. Payload 71](#payload-1)

[14.1.4. Power 71](#power)

[14.1.5. Flight Dynamics 71](#flight-dynamics-1)

[14.1.6. Propulsion 71](#propulsion-1)

[14.1.7. Communications and Data Handling
71](#communications-and-data-handling-1)

[14.1.8. Ground Operations 72](#ground-operations)

[14.1.9. Subsystem Organizational Changes since Progress Report
72](#subsystem-organizational-changes-since-progress-report)

[15. Appendix C: Glossary 72](#appendix-c-glossary)

[16. Endnotes 75](#endnotes)

# List of Figures

[Figure 1: Possible fossilized microbe in ALH84001
9](#__RefHeading___Toc81459542)

[Figure 2: Clouds of Valles Marineris 11](#__RefHeading___Toc81459543)

[Figure 3: Valles Marineris 12](#__RefHeading___Toc81459544)

[Figure 4: Postscript Identification Information for Melas Chasma
13](#__RefHeading___Toc81459545)

[Figure 5: Gridview Software Visualization on Melas Chasma from MOLA
13](#__RefHeading___Toc81459546)

[Figure 6: Melas Chasma Thermal Imaging Map from THEMIS
14](#__RefHeading___Toc81459547)

[Figure 7: Geomorphology of Landing site, Melas Chasma
15](#__RefHeading___Toc81459548)

[Figure 8: Melas Chasma Landing sites 16](#__RefHeading___Toc81459549)

[Figure 9: MER Descent Configuration 28](#__RefHeading___Toc81459550)

[Figure 10: MER Descent Configuration - Exploded View
28](#__RefHeading___Toc81459551)

[Figure 11: Airbag Landing Sequence 29](#__RefHeading___Toc81459552)

[Figure 12: Mars Pathfinder Lander Configuration
29](#__RefHeading___Toc81459553)

[Figure 13: Honeybee 10 m Drill 33](#__RefHeading___Toc81459554)

[Figure 14: Honeybee Core Sampling System
33](#__RefHeading___Toc81459555)

[Figure 15: Canadrill 34](#__RefHeading___Toc81459556)

[Figure 16: Canadrill Lander Concept 34](#__RefHeading___Toc81459557)

[Figure 17: Honeybee Sample Transport System
35](#__RefHeading___Toc81459558)

[Figure 18: Block Diagram of Electrical Power Subsystem Hierarchy
36](#__RefHeading___Toc81459559)

[Figure 19: Graphic Representation Used to Solve for Average Day Time
37](#__RefHeading___Toc81459560)

[Figure 20: Hohmann Transfer for Cruise 42](#__RefHeading___Toc81459561)

[Figure 21: Launch Profile for Delta II 44](#__RefHeading___Toc81459562)

[Figure 23: EDL Components 49](#__RefHeading___Toc81459563)

[Figure 24: Lander Antenna Array^67^ 49](#__RefHeading___Toc81459564)

[Figure 25: RAD750 3U Compact PCI 51](#__RefHeading___Toc81459565)

[Figure 26: DSN 70-meter antenna 53](#__RefHeading___Toc81459566)

[Figure 27: These are the radiation problems that will be encountered^2^
57](#__RefHeading___Toc81459567)

[Figure 28: Homepage for the MSCLE scout mission 64](#_Ref79743964)

[Figure 29 : A portion of the page that asks visitors to guess which
pictures are from Mars and which are from Earth. Holding the cursor over
\"Answers\" uncovers the mystery. 64](#_Ref79744377)

# List of Tables

[Table 1: Ellipse Landing Site Information (see Figure 8)
15](#__RefHeading___Toc81459596)

[Table 2: E14-01154 (see Figure 1) 16](#__RefHeading___Toc81459597)

[Table 3: E14-01708 (see Figure 8) 16](#__RefHeading___Toc81459598)

[Table 4: E14-02205 (see Figure 8) 16](#__RefHeading___Toc81459599)

[Table 5: Instruments for searching for life on Mars\*
22](#__RefHeading___Toc81459600)

[Table 6: Instrument Specifications from Mars 96 penetrator^21^
23](#__RefHeading___Toc81459601)

[Table 7: Instrument Specifications for the camera of Beagle2 ^22^
24](#__RefHeading___Toc81459602)

[Table 8: Subsystem Masses 30](#__RefHeading___Toc81459603)

[Table 9: Payload Masses 31](#__RefHeading___Toc81459604)

[Table 10: Further Mass Calculations 31](#__RefHeading___Toc81459605)

[Table 11:Modes of Operation for MSCLE 36](#__RefHeading___Toc81459606)

[Table 12: Cruise Mode Requirements 37](#__RefHeading___Toc81459607)

[Table 13: Science Mode Power Requirements
38](#__RefHeading___Toc81459608)

[Table 14: EarthCOM M ode Power Requirements
38](#__RefHeading___Toc81459609)

[Table 15: MarsCOM Mode Power Requirements
38](#__RefHeading___Toc81459610)

[Table 16: Standby Mode Power Requirements
39](#__RefHeading___Toc81459611)

[Table 17: Trade Study of Propulsion Systems
45](#__RefHeading___Toc81459612)

[Table 18: RAD750 3U Compact PCI Summary
51](#__RefHeading___Toc81459613)

[Table 19: Station Support Parameters for Various Mission Phases
54](#__RefHeading___Toc81459614)

[Table 20: Schedule for Payload/Science Operations on Martian Surface
55](#__RefHeading___Toc81459615)

[Table 21: Pros and Cons of each cost estimation method
59](#__RefHeading___Toc81459616)

[Table 22: Mars Mission Analogies 60](#__RefHeading___Toc81459617)

[Table 23: Parametric Cost Model Parameters
61](#__RefHeading___Toc81459618)

[Table 24: Second Iteration Cost Estimates
62](#__RefHeading___Toc81459619)

[Table 25: Third Iteration Cost Estimates
62](#__RefHeading___Toc81459620)

[Table 26: Mass Budget Calculations 69](#__RefHeading___Toc81459621)

# Executive Summary

The result of an advanced concept study, the Mars Subsurface Chemical
Life Explorer (MSCLE) crosses the threshold into the next era of the
human space adventure. Delving up to six meters below the surface in its
search for biosignatures, MSCLE utilizes instrumentation orders of
magnitude beyond its life-searching predecessors. However, the
flight-readiness of most of the engineering components and
instruments---especially the drilling auger---is highly advanced. This
supports the budgetary and scientific feasibility for MSCLE as the
PI-class Scout Mission to Mars.

## Science Development and Operations

In two stages, four different measurement packages will analyze crucial
chemical features of martian rock samples. A gross set of experiments in
the first stage will determine a baseline of evidence of chemical
byproducts and historic habitability. The **gas chromatograph-mass
spectrometer** will decipher the organic molecular composition of a
given sample. ~~Pyrolysis and derivatization extract the gas from the
acquired soil constituents via heating and chemical reactions,
respectively. Once the gas is exposed, the spectrometer pinpoints the
exact molecular species and their molecular abundances via spectral
analysis.~~ Not only have recent observations shown methane in the
martian atmosphere, but terrestrial methanogenesis is almost strictly a
biotic process; this particular organic molecule thus has a high
potential for a tracer of present metabolic processes.

To determine mineral composition, the **alpha/proton/x-ray
spectrometer** will find and discern between low- and high-mass atomic
species via nuclear and photonic scattering. Helium nuclei (alpha
particles) will reveal small atoms upon bombardment, while x-rays will
detect heavier than iron. ~~Alpha particles targeted at a rock sample
will undergo Rutherford scattering over a large energy range for low-Z
atoms. High-mass elemental species can absorb and reemit x-rays
according to their electronic configuration.~~

Direct detection of biological processes will reveal more conclusive
evidence than the baseline chemical analysis. Particularly, amino
acid-generation and microbial respiration are strong signs of active
present life, for which MSCLE will search with two different experiments
in stage two. ~~First, molecular handedness, much like parity, exists in
two types---left and right; a left-handed amino acid will not look like
it's right-handed cousin after any spatial transformation (rotation,
flipping) is performed upon it.~~ Terrestrial biological systems have a
handedness, or chirality, bias, which the **Chirality Recognition and
Analysis Package** intends to exploit in Martian analogy. Once specimens
are obtained from the soil, amino acids are collected through pyrolysis
and electrophoresis, then fed into a small capillary system. Here,
left-handed (that which life prefers) molecules will react more readily
than their cousins when cyclodextrin is injected into the ystem.

Concluding experimentation for the primary objective, **passive gas
exchange** will detect changes in gas concentration for which geological
processes may not account. A cocktail of gases (H~2~, N~2~, O~2~ and
others) to which microbial life might respond is injected into the
samples, and concentrations are monitored for two weeks to watch for
organisms ingesting the new atmosphere. ~~Control environments are
maintained to enable a comparison between the doctored and unaffected
environments. The magnitude of change in abundance of the gases will
reveal the presence or absence of metabolizing life forms.~~

**Atmospheric measurements** and **high-resolution image** acquisition
comprise the second and third objectives, respectively. Temperature,
pressure, and humidity sampling will asses the presence of liquid of
water and provide an additional constraining and discriminative element
for the primary experiments. A wide-angle, multispectral camera will
study mineralogy and cliff stratification; most crucially, it will aid
in the determination of the type of rock that is analyzed. Solar and
Martian lunar observations have potential utility in determining
orientation of the lander.

The location optimal for MSCLE's experiments and studies is a valley
that dwarfs Earth's grand canyon: Valles Marineris lays on the equator
of Mars, sinks approximately 6 km into the surface and spans nearly
4,000 km in length. Particularly, the Melas Chasma region---located
centrally in the valley's 600-km width---has potential for life, due to
extended exposure to the summer sun and depths that suit moisture
accumulation.

## Technical Development

To achieve the desired landing site, debate arose between retrorockets
and airbag cushioning for direct impact---a la Mars Polar Lander and
Mars Exploration Rovers, respectively. In the interest of mass economy
and simplicity, MSCLE will bounce its way into one of four preferred
landing ellipses. Most of the remaining design challenges and solutions
have occurred in the realms of power allocation, mobility and drill
selection. The most efficient energy accumulation comes from stationary
solar array panels; complexity and cost prohibit a roving system; and a
rotary percussion auger was selected from a slew of different types,
including pile drivers, lasers and acoustically driven systems.

Accompanying these design trades are the main design drivers---drill
depth, soil material, and sample acquisition---were dictated by the
science objectives. The latter two were intricately related to the drill
selection: a wide range of soil hardness and composition necessitated
the versatile auger; at a maximum of six meters, the depth required by
science is not dramatic, but optimized for a simple mechanical system,
rather than untested laser or acoustic technology. Safety and integrity
of the sample during processing is essential: to prevent contamination
and ensure efficient experimentation, a redundant system is yet to be
mapped in detail.

## Budget

Alongside the quid pro quo of system resource exchange, budget
constraints shaped the MSCLE lander design. Several iterations and three
main types of cost analysis methods comprised the budget investigation.
With only general information, the first iteration used an analogy-based
method to compare MSCLE with several previous Mars excursions, finding
that it exceeded the cost cap by a little over 10 percent. By the final
evaluation, the most reliable estimation process placed the cost for
MSCLE (including a 20% cost growth margin) nearly 30 % under the limit.

## Outreach

Chief among MSCLE's public enrichment goals is the encouragement of
present and future generations of explorers to work, imagine, and live
outside the limitations. Despite those boundaries imposed by the design
and actualization of space missions, discovery is the heart and
sustenance of exploration. And, more than what we find in the dark outer
reaches, it is that which we encounter in the shadows of ourselves that
feeds us and indeed that which eventually defines the adventure.

The realization of this aspiration takes root in the search for
knowledge and understanding of our surroundings. In this, the
researchers producing MSCLE are dedicated to enhancing the public's
appreciation for science and space exploration, and their myriad
benefits for humanity.

# Science Operations and Development

## Science Background

### History

It has been almost 400 years since Galileo Galilee observed Mars with
the first telescope, and it has been almost 350 years since Christian
Huygens drew the first maps of Mars. Humanity's fascination with the Red
Planet, however, has not wavered in that time. Fortunately, technology
has finally advanced so that spacecraft can be built to visit the
planet. In 1964, Mariner 3 and 4 flew by Mars, taking pictures of Mars
that revealed impact craters and frost on the surface. Only eleven years
later, Viking 1 and 2 landed on the surface of Mars at the western
slopes of Chryse Planitia and Utopia Planitia, respectively. Throughout
all of the investigations of Mars, there has been one common thread: the
innate propensity of humans to wonder whether there is life elsewhere in
the Universe. The Viking landers aimed to answer this question by
testing the soil for metabolic processes and although no definitive
answers were produced, the Viking mission paved the way for future
missions to find the origins of life[^1].

Recently, the National Aeronautics and Space Administration (NASA)
developed a "community-based forum designed to provide input for
planning and prioritizing Mars exploration activities for the next
several decades" called the Mars Exploration Program Analysis Group
(MEPAG). Lead by NASA's Lead Scientist for Mars Exploration, MEPAG
created a list of four goals for the immediate future of Mars
exploration. A document titled "Scientific Goals, Objectives,
Investigations, and Priorities: 2004" presented to the public on July
16, 2004 outlined those four goals. The first of these goals is to
"determine if life ever arose on Mars." The report advises scientists to
obtain this goal by assessing the past and present habitability of the
planet and testing for prebiotic processes that may confirm past or
present life in environments that have habitable potential[^2].

The 2004 NASA Academy at Goddard Space Flight Center presents a design
for a scout mission called MSCLE: Mars Subsurface Chemical Life Explorer
that will help answer the first MEPAG goal. The primary objective of
this mission is to search for present life on Mars in the subsurface
rock layers of Melas Chasma in Valles Marineris. The secondary objective
is to measure the weather conditions at the landing site, and the
tertiary objective is to further analyze the site terrain with visual
data.

The three other goals of MEPAG were not ignored by the Science Team, who
intently discussed and debated its approach in meeting one or more of
the goals. Initially, the team created a very broad objective, planning
to insert three or more penetrators into the surface of Mars that would
measure chemical and physical characteristics of the interior of the
planet and assess the planet's habitability. Upon further investigation
of the engineering requirements and cost of such a mission, the team
narrowed its objective.

As both the mission and vision for NASA state the necessity to search
for and find life beyond our home planet, the 2004 NASA Academy at
Goddard Space Flight Center proposes to do just that.

### Prospects for Martian Life

Although the images Viking returned of the Martian surface revealed a
bleak and seemingly inhabitable planet, recent studies of life on Earth
have made scientists optimistic about the potential of life existing in
certain habitable zones on Mars and other planetary bodies. Bacteria
have been found thriving in some of the most extreme environments on
Earth, for which they are referred to as \"extremophiles,\" suggesting a
strong possibility that life could very well survive in some places on
Mars. For example, a class of bacteria, called nanobacteria, respire
without oxygen. At approximately 20 to 150 nanometers in length, they
are roughly equivalent in size to most viruses. These minute organisms
were discovered just four years ago, three miles below the seabed
surrounding Australia. The existence of these extremophiles reinforce
arguments made proposing the presence of fossilized Martian microbes in
ALH84001, a Martian meteorite uncovered in Antarctica (Figure 1), in
that nano-sized microbial life may very well be able to thrive in
extreme environments like those on Mars.

![](media/image1.jpeg){width="3.196527777777778in"
height="2.1659722222222224in"}

[]{#__RefHeading___Toc81459542 .anchor}Figure 1: Possible fossilized
microbe in ALH84001[^3]

Several factors that make Mars apparently inhospitable for biological
systems; for example, there exist high levels of radiation and extreme
temperatures, factors which terrestrial life forms have demonstrated the
ability to tolerate. A species of "superbacteria," *deinococcus
radiodurans,* was discovered in cow manure after surviving multiple
food-sterilization processes[^4], and it has been found to withstand
intense radiation.

The hyperthermophilic extremophile crenarchaeota thrives in sulfuric
thermal vents deep within the ocean at temperatures between 80 and
100°C.[^5] Other extremophiles have been discovered deep beneath the ice
in Antarctica. Scientists have even found a group of extremophiles
called \"spores\" that have been known to lay dormant for 30 to 250
million years until the right condition revives them.[^6] These
deep-dwelling bacteria have been found as deep as 3 meters in subsurface
terrain, withstanding severe heat and pressure to survive.[^7] Taking
these organisms into consideration, it would appear that life might be
able to survive on Mars. There may actually be many Martian environments
that could be conducive to extremophilic life forms.

Some chemical evidence now exists that suggests such extremophilic life
may exist on Mars currently. The Mars Exploration Rovers have recently
discovered methane in the Martian atmosphere, a gas that is present in a
concentration greatly inconsistent with an oxygenic atmosphere on Earth
because of biological systems. As methane is a byproduct of biological
systems, it serves as a marker of where biological materials are
decaying. Localizing the sources of methane[^8] on Mars may help locate
potential life forms. Continuing along these discoveries, it would be
interesting to search for the presence of ammonia and formaldehyde on
Mars. Ammonia[^9] and formaldehyde[^10], produced terrestrially when
methanol (wood alcohol) is metabolized, also serve as biological markers
coming from decomposition of organic materials. Since ammonia has an
atmospheric half-life of only 2 months and formaldehyde only months,
discovery of such gases on Mars would be most certainly not from natal
cloud formation and star birth, but biological organisms.

Another interesting phenomenon on Mars is the mist over Valles
Marineris, particularly over Melas Chasma and Candor Chasma, that can be
seen drifting toward the west over Noctis Labyrinthus and the Tharsis
region in images from Mars Global Surveyor. Although Mars does not
appear to have obvious global biogeochemical cycling, these clouds may
be an product of local biochemical activity (Figure 2).

![](media/image3.jpeg){width="1.5263888888888888in"
height="3.247916666666667in"}

[]{#__RefHeading___Toc81459543 .anchor}Figure 2: Clouds of Valles
Marineris[^11]

### Landing Site

Valles Marineris is a system of canyons located just south of the
Martian equator (Figure 3). The system is about 4,000 km long, and, if
placed on Earth, would extend all the way across the United States. The
central individual troughs, generally 50 to 100 km wide, merge into a
depression as much as 600 km wide. In places, the canyon floor reaches a
depth of 7 km, 6 to 7 times deeper than the Grand Canyon. The geologic
history of the central canyon system is undoubtedly complex. Most
researchers agree that Valles Marineris is a large tectonic \"crack\" in
the Martian crust, forming as the planet cooled, affected by the rising
crust in the Tharsis region to the west, and subsequently widened by
erosion forces. First the surface collapsed into a few deep depressions
that later became filled with layered material, perhaps as lake
deposits. Then graben-forming faults cut across some of the older
troughs thus widening existing troughs, breaching barriers between
troughs, and forming additional ones. At that time the interior deposits
were locally bent and tilted, and perhaps water, if still present,
spilled out and flowed toward the outflow channels. Huge landslides fell
into the voids created by the new grabens. Wind-drifted material, mostly
dark in color, apparently still moves along the canyon floor and locally
forms conspicuous dunes.

![](media/image4.jpeg){width="3.698611111111111in"
height="2.9569444444444444in"}

[]{#__RefHeading___Toc81459544 .anchor}Figure 3: Valles Marineris[^12]

The Science Team has chosen the Melas Chasma region as its landing site
because of its central location in Valles Mariners and its environment
suitable for possible extremophilic life. This part of the canyon is
very deep, an attribute allowing moisture from clouds to stay near the
surface even after the clouds have dissipated. Since it is located near
the equator, the summer sun remains in the sky longer, providing warmth
and energy for potential life.

Using Gridview's Postscript tool to read Mars Orbital Laser Altimeter
(MOLA) data, it was found that Melas Chasma is approximately 7 to 8 km
deep and has a stable floor of

approximately 140 to 145 kilometers wide inside the total width of Melas
Chasma, approximately 310 to 315 km (Figure 4). Gridview also determined
the geographic location of our landing site: center latitude 9.1861 S to
center longitude 73.413 W (Figure 5). Elevation is determined by the
color of the resolution; black is low, and white is high.

![](media/image5.png){width="4.622916666666667in"
height="2.9381944444444446in"}

[]{#__RefHeading___Toc81459545 .anchor}Figure 4: Postscript
Identification Information for Melas Chasma[^13]

![](media/image6.jpeg){width="2.1083333333333334in"
height="2.7916666666666665in"}

[]{#__RefHeading___Toc81459546 .anchor}Figure 5: Gridview Software
Visualization on Melas Chasma from MOLA[^14]

Thermal data from the MGS Thermal Emission Spectrometer (THEMIS) reveals
that the walls of the canyon in the Melas Chasma region are sloped, but
with minimal regolith, thus making it easier to land and drill. In the
false-color image shown in Figure 6, the blue areas represent loose
regolith and the red areas represent solid rock. This data also tells us
that the daytime temperature of Melas Chasma is between -5ºC to 35ºC,
the warmest temperatures occurring at noon, 12:17am.

![](media/image7.png){width="4.013888888888889in"
height="2.8152777777777778in"}

[]{#__RefHeading___Toc81459547 .anchor}Figure 6: Melas Chasma Thermal
Imaging Map from THEMIS[^15]

The landing ellipse chosen for the landing site in Melas Chasma is one
of the four sites seriously considered for the Mars Exploration Rovers.
It is in the southwestern side of Valles Marineris (Figure 7) in the
vicinity of lava flows from the Tharsis region. From this, we can assume
that most of the geological data taken will reveal the composition of
the flows. The geomorphologic setting of Melas Chasma is shown in Figure
6 as well. Landside deposits overlain by interior deposits and
sheet-sands fill the area, and dunes cover most of the surface. The area
surrounding our landing ellipse is polka-dotted with flat-topped mesas,
often referred to as interior layered deposits. Although their origin is
unknown, it is hypothesized that they are lacustrine deposits, marking
the location of a former lake. These deposits are an ideal location for
the search for life, having once been near and/or under water, yet
landing on top of one takes a great deal of precision and power. In
order to remain in budget and to keep the risk of the mission low, the
Science Team has decided to use a MER landing ellipse near an interior
layered deposit rather than trying to land on top of one.

![](media/image8.jpeg){width="3.738888888888889in"
height="3.6777777777777776in"}

![](media/image9.jpeg){width="2.31875in" height="0.6298611111111111in"}
![](media/image10.jpeg){width="2.3201388888888888in"
height="1.2951388888888888in"}
![](media/image11.jpeg){width="2.3201388888888888in"
height="1.3888888888888888in"}
![](media/image12.jpeg){width="2.3194444444444446in" height="0.7625in"}
![](media/image13.jpeg){width="2.31875in" height="0.7826388888888889in"}

[]{#__RefHeading___Toc81459548 .anchor}Figure 7: Geomorphology of
Landing site, Melas Chasma[^16]

The following are the final stats in the landing report and the actual
dimensions of each ellipse as shown in Figure 8.

  -------------------------------- --------------------------- ---------------------------- -------------------------
  []{#__RefHeading___Toc81459596                                                            
  .anchor}Table 1: Ellipse Landing                                                          
  Site Information (see Figure 8)                                                           

  ***[Latitude]{.underline}***     *[Longitude]{.underline}*   *[Resolution]{.underline}*   *[Ellipse]{.underline}*

  **-8.72**                        76.73                       3.01                         E14-01154

  **-8.84**                        77.40                       3.00                         E14-01708

  **-8.96**                        78.04                       3.00                         E14-02208
  -------------------------------- --------------------------- ---------------------------- -------------------------

  ------------------------------------------ ----------------------------
  []{#__RefHeading___Toc81459597             
  .anchor}Table 2: E14-01154 (see Figure 1)  

  **Altitude:**                              382.67km

  **Slant Range:**                           400.00 km

  **Emission Angle:**                        17.89°

  **Incidence Angle:**                       19.17°

  **Phase Angle:**                           24.41°
  ------------------------------------------ ----------------------------

  ------------------------------------------ ----------------------------
  []{#__RefHeading___Toc81459599             
  .anchor}Table 4: E14-02205 (see Figure 8)  

  **Altitude:**                              381.60 km

  **Slant Range:**                           398.92 km

  **Emission Angle:**                        17.91°

  **Incidence Angle:**                       20.54°

  **Phase Angle:**                           22.97°
  ------------------------------------------ ----------------------------

![](media/image14.jpeg){width="2.821527777777778in"
height="2.7222222222222223in"}

[]{#__RefHeading___Toc81459549 .anchor}Figure 8: Melas Chasma Landing
sites[^17]

## Primary Objective: Search for Present Life

### Stage 1: Characterizing Chemical Composition of Soil and Rock Constituents

#### Experiment

Detecting signatures of present life will require a detailed and
sensitive analysis of the chemical composition of the soil and rock at
the landing site. This will enable detection not only of chemical
evidence of biological systems (such as amino acids, proteins, and other
complex organic molecules), but of axpsthe potential for the
habitability of biological systems in the past and present. The chemical
analysis will be performed on collected samples of crushed soil and
rock, with particular attention to the presence of amino acids,
organics, water, biologically important elements such as nitrogen and
sulfur, and the overall mineral environment. Since the soil and rock
below the surface is potentially warm, wet, and protected from
destructive radiation and atmospheric oxidizers such as peroxides,
taking samples from various depths will increase the chance of finding
possible signatures from past and present life. Additionally, taking
samples at multiple depths will help to establish a concept of the
geological and potential biological history at the landing site. For
these reasons, samples from various depths will be collected and
analyzed.

#### Instrumentation

*Gas Chromatograph/Mass Spectrometer (GCMS)*

###### Rationale {#rationale .unnumbered}

The GCMS will allow for analysis of inorganic and organic compounds in
the Martian soil. The main focus of the instrument will be in searching
for organic compounds. There are normally three possible sources of
organics on a terrestrial planet: cometary/meteoritic infall, chemical
reactions in the atmosphere, and biological processes. On Mars, it is
unlikely that organics could form through abiotic chemical processes in
the strongly oxidizing atmospheric environment. The objective of this
experiment is to search for organics in soil and rock samples and
differentiate between those brought to the surface by other bodies and
those that may have been created by life[^18].

It is important to address the inclusion of this instrument in light of
experiments done by the Viking landers. The initial results from the
Viking GCMS experiment to detect organics in near surface material were
negative, possibly for a variety of reasons. At the surface, there are
probably very small amounts of large organic molecules (if any at all)
because they would be oxidized relatively quickly by the atmosphere (as
ionizing radiation and peroxides are prevalent). In order to find
organic compounds that are not affected by the atmosphere, we will have
to go down to depths of 1-10 meters rather than a few centimeters,
depending on the density and type of the soil[^19]. Another major reason
the Viking experiments were inconclusive is that the instruments were
not sensitive enough. Experiments have been done on Earth in which live
organisms were injected into Martian analog soil, and organic
composition was measured with techniques similar to the Viking
experiments[^20]. After being pyrolyzed at Viking GCMS conditions, the
sample yielded no noticeable organics. The sensitivity required for
detection of these amounts of organics would need to be orders of
magnitude greater. The instrument we propose will be much more sensitive
than the Viking GCMS, and we will use different extraction techniques to
get larger amounts of organics into a gas sample.

###### Instrument Description {#instrument-description .unnumbered}

The gas chromatograph (GC) separates a mixture of gas into its
constituents. The major components of the GC are:

i.  carrier gas reservoir (helium in this case) and pressure controller;

ii. an injection subsystem where the sample is carried into GC columns
    by the carrier gas;

iii. the GC columns, where the separation occurs;

iv. a detector at the end of the column where the gas exits, which will
    indicate when various compounds pass through the instrument.

v.  The mass spectrometer allows identification of each GC peak, but the
    carrier gas must be mostly removed before it enters the mass
    spectrometer[^21].

There are various extraction methods that can be used to get a useable
gas sample out of a soil or rock sample. Pyrolysis, which was used in
the Viking experiments, involves heating a soil sample up to about 500°C
and analyzing the released gases. Pyrolysis allows bulk isotope ratios
of the samples to be measured, but this method is not optimal for
getting isotopic ratios from individual compounds because it destroys
large molecules. Derivatization is another technique that will be
implemented for analyzing non-volatile amino acids. This method involves
application of a chemical derivatization agent that makes the organic
compounds more volatile, and allows them to pass through the GC column
more easily. Since large organic molecules remain intact for analysis by
the mass spectrometer, this enables us to do compound-specific analysis
that provides information about metabolic pathways.

The GCMS will also be coupled to the Passive Gas Exchange Experiment,
described below.

###### Ground Analysis of Data {#ground-analysis-of-data .unnumbered}

The GCMS will provide information about the molecular masses and number
of compounds present in a sample. This information will be used to
determine the composition of chemical species in the soil and rock. From
the database of species compiled, the presence and concentrations of all
species will be analyzed, with certain species analyzed more in depth.

Volatile, low-mass organics are of interest, especially those that could
be a sign of present life. Methane, for example, has a relatively short
lifetime in the Martian atmosphere, so finding substantial amounts of
CH~4~ indicates some steady source (possibly methanogens).

The GCMS will also allow for determination of carbon isotope ratios in
organic compounds on Mars. Life on Earth prefers ^12^C over ^13^C, and
overall fractionations due to biological activity can be measured. If
organics are found on Mars, their origin can be determined by comparing
the δ^13^C values to the expected natural isotopic abundance (^13^C
/^12^C~standard~ = 1.1%). The amount of isotopic fractionation due to
life varies with the type of life that is observed: photosynthetic
organisms fractionate carbon to a lesser amount while methanogens
fractionate to a much greater degree (up to 70 ppm).

Individual compound analysis is useful as well, because determining the
carbon isotope ratio of amino acid pairs indicates the metabolic
pathways that are used. Compound-specific analysis gives information
about general metabolic trends, and allows us to discriminate between
major metabolic processes (such as carbon fixation versus carbon
oxidation). Some pathways are trans-animated, and the δ^13^C value of
the precursor and product amino acids will be equal to the δ^13^C of the
background carbon source. However, some pathways are multi-branched and
enzymes can discriminate between isotopes. Amino acids derived from
these processes reflect both their precursor and pathway.

Other isotopes of interest include ^2^H, ^15^N, and ^34^S. Methanogens
on Earth fractionate hydrogen to a noticeable degree, and we will look
for this same fractionation on Mars. Determination of ^15^N/^14^N and
^34^S/^32^S will also provide information on possible biological
activity, as nitrogen and sulfur are both biologically important
elements in terrestrial biological systems.

*Alpha/Proton/X-Ray Spectrometer (APXS)*

###### Rationale {#rationale-1 .unnumbered}

The APXS provides the ability to obtain a detailed chemical and
elemental analysis of soil and rocks at the landing site by irradiating
a sample and analyzing the induced radiation. Doing so will allow us to
detect elemental compositions heavier than He and determine the mineral
composition of subsurface samples. The instrument allows the study of
the climatological history and past habitability of Mars from a chemical
perspective, which will tell us about the probability of finding extant
life[^22].

###### Instrument Description {#instrument-description-1 .unnumbered}

The APXS irradiates a sample of rock and analyzes the induced radiation.
This allows it to detect elemental compositions heavier than He (Z=2),
which can be used to determine what minerals make up subsurface samples.
The APXS can operate in either \"alpha\" or \"X-ray\" mode to detect
elements of low or high mass, respectively. In alpha mode, sample nuclei
are targeted for collisions with alpha particles, resulting in
Rutherford scattering. The energy of the scattered particles,
characteristic of the scattering nucleus, is then measured. This
approach only works for light elements, however, because at low masses
the maximum energies of the scattered particles are widely varied.

For heavier elements, the instrument\'s X-ray mode can be used. This
method involves exciting atoms in the sample with alpha particles and
analyzing the characteristic X-rays that are emitted. This will provide
us with the ability to detect minerals containing elements heavier than
iron (Z=26)[^23].

###### Ground Analysis of Data {#ground-analysis-of-data-1 .unnumbered}

With the data collected from the spectrometer, identification and
concentrations of chemical species will be deduced and analyzed. The
primary focus is to detect chemical changes in the soil and rocks caused
by water and to determine the current water concentration as a function
of depth. Of particular interest are minerals, such as phosphates,
manganese oxides, and carbonates because biological processes produce
them. Concentrations of nitrates and sulfates will also be analyzed, as
they are major byproducts of biological systems, although not as
strictly attributable to biological systems as the aforementioned
targets. Lastly, as our current understanding of oxidation
processes[^24] in the Martian soil is inadequate, measurements of
subsurface abundances of absolute oxygen will be taken. Aside from
increasing the overall knowledge of geological processes in Valles
Marineris, this will aid in establishing a better understanding of the
\"chemical background\" of the soil so that we can better identify
chemical signatures unexplainable by geology alone (i.e. microbial
contributions to soil chemistry).

### Stage II: Search for Distinctive Biological Signatures

#### Experiment

Although detailed chemical analysis will provide information about the
presence of organics and potential habitability for life at the landing
site, it does not provide direct evidence of past or present biological
systems. For this reason, two experiments designed to collect direct
evidence of biological systems will be conducted---one to ascertain the
presence of biologically-generated amino acids and one to detect
respiration of microbes.

#### Instrumentation

*Chirality Recognition and Analysis Package*

**Rationale**

In terrestrial biological systems, amino acids are all left handed.
Amino acids are produced outside of biological systems, for example in
meteorites, but are usually found in racemic mixtures (both left- and
right-handed species). A bias in the chirality of amino acids found on
Mars almost definitely indicates the presence of biological systems. The
Chirality Recognition and Analysis Package will utilize the Mars Organic
Analyzer system (MOA) developed by the Jet Propulsion Laboratory to look
for both the chemical signature of amino acids and test for their
chirality. Although the GCMS has limited ability to detect amino acids,
the pyrolysis of the GCMS conditions cause the conversion of amino acids
to amines (especially abundant glycine to methyl amine), and these
products buried under the intense carbon dioxide and water peaks in the
mass spectrometer traces making them undetectable.[^25] The MOA system,
however, does not require pyrolysis and can detect amino acids directly.

Although the Viking landers in the 1970s did not conclusively detect
organic molecules on Mars, their sensitivity was such that that they
would have failed to detect life past a threshold of a million bacteria
per gram of soil. Recent discoveries from the Spirit and Opportunity
rovers on Mars have provided strong evidence that standing water once
existed on the surface Mars, suggesting a strong possibility there could
have also been organic molecules.

**Instrument Description**[^26]

The first stage of the system is the Mars Organic Detector (MOD) system.
A sample of crushed soil or rock is fed into the system and then heated
under Martian atmospheric pressure. This will vaporize any organic
molecules that may be present. The vapor will then be condensed onto a
cold finger cooled to the ambient nighttime temperature (approximately
-73°C). The cold finger is coated with fluorescamine dye tracers that
bind only to amino acids. Appearance of a fluorescent signal indicates
that amino acids or amines are present. The sensitivity of the system is
one trillionth of a gram of amino acids per gram of soil, that is, a
million times better than Viking.

Upon detection of amino acids or amines, a capillary electrophoresis
system sips the condensed fluid off the cold finger and siphons it into
the second stage of the system, a MEMS lab-on-a-chip for identification
and chiral analysis of the amino acids.

The lab-on-a-chip consists of channels etched by photolithographic
techniques and a microfluidic pumping system sandwiched into a
four-layer disk four inches in diameter, with the layers connected by
drilled channels. The chip utilizes micropumps and microvalves to route
sample fluid past a series of chemicals that will help identify the
amino acids. The different amino acids are first crudely separated by
charge and size as they travel through the system, which helps identify
them partially. The sample is further resolved by chemical analysis of
the amino acids. The amino acids are then routed down a thin capillary
tube lined with cyclodextrin. Since left-handed amino acids interact
with cyclodextrin more than right-handed amino acids, the left-handed
amino acids will take longer to move through and the enantiomers sort
out.

**Ground Analysis of Data**

Information about the identity and chirality of amino acids detected
will be collected and analyzed. If amino acids are detected, this will
suggest possibilities for past or present biological systems. If a
racemic mixture of amino acids is detected, the amino acids are likely
of non-biological origin. If a chiral excess is detected, the amino
acids have to be of biological origin.

*Passive Gas Exchange Experiment*

The Passive Gas Exchange Experiment (PGEEx) aims to detect changes in
concentration of gases (specifically, H~2~, N~2~, O~2~, CO~2~, and Kr)
from a sample of soil over the course of several days---which cannot be
explained by geological processes---to ascertain the presence of active
microbial life. Where the other instruments are more capable of
determining the presence of organic materials and biochemical systems,
they are not expressly capable of detecting present life forms. This
experiment adds that capability, provided positive results from the
other instruments. Unlike the fathering Viking experiment, no water or
\"nutrient broth\" will be added to stimulate microbial activity, as the
construction of Martian biochemical pathways is unknown, and it is
uncertain what types of nutrients to which Martian microbes would
respond.

Experiments done by the Viking landers suggest a cold, arid surface
environment, apparently suffused with oxidants capable of degrading
organic compounds. There appears not to be a dominant global
biogeochemical cycling of major elements on Mars, but this does not
preclude isolated biogeochemical cycles in either local or widespread
environments that are more hospitable. Because of the high probability
of moisture content and protection from radiation at the landing site
and under its surface, it is believed that this experiment may be
valuable.[^27]

The experiment will feature improved sensitivity over the Viking
package, making use of a more sensitive gas chromatograph/mass
spectrometer and should be capable of detecting much more subtle
disequilibria in gaseous components inconsistent with the current Marian
atmosphere.

A sample of Martian soil or rock is divided into four equal portions and
placed into two sealed chambers with a volume of Martian atmosphere. One
of the two chambers is heated to 100°C to sterilize it for a control.
Into one of the sterilized chambers and one of the unsterilized chambers
is injected a specific amount of H~2~, N~2~, O~2~, CO~2~, and Kr gases.
The other two chambers are left alone. This setup allows for detection
of respiration and accounts for the possibility that the addition of
gases to a sample may suffocate or poison microbes in the sample by not
adding gases to one of the chambers. A full day will be devoted to
meteorological measurements at the beginning of the mission. In the
following days, small amount of the atmosphere in each chamber is
extracted and analyzed by the GCMS instrument once each day. [^28]

Concentrations of gases in each of the containers will be monitored over
the course of 14 days. It is believed that a steady increase or decrease
of any one of these gases in a sealed sample of soil or crushed rock may
indicate the presence of microbial respiration. This of course will only
be valid if suitable results are obtained in ground experiments with
sterilized and unsterilized ground samples and if the other experiments
detect the existence of organic molecules in the soil.

+-------------+------------+--------+----------+----------+----------+
| []{#__Re    |            |        |          |          |          |
| fHeading___ |            |        |          |          |          |
| Toc81459600 |            |        |          |          |          |
| .a          |            |        |          |          |          |
| nchor}Table |            |        |          |          |          |
| 5:          |            |        |          |          |          |
| Instruments |            |        |          |          |          |
| for         |            |        |          |          |          |
| searching   |            |        |          |          |          |
| for life on |            |        |          |          |          |
| Mars\*      |            |        |          |          |          |
+-------------+------------+--------+----------+----------+----------+
| ***[Inst    | *[P        | *[Mass | *[Power  | *[Vol    | *[Da     |
| rument]{.un | urpose]{.u | (kg)]{ | (        | (cm^3    | ta]{.und |
| derline}*** | nderline}* | .under | W)]{.und | ^)]{.und | erline}* |
|             |            | line}* | erline}* | erline}* |          |
+-------------+------------+--------+----------+----------+----------+
| *Al         | Elemental  | 0.5    | 0.4 (1.3 | 300      | 16       |
| pha/Proton* | analysis   |        | peak)    |          | k        |
|             | for all    |        |          |          | b/sample |
| *X-ray      | except H   |        |          |          |          |
| Sp          | and He     |        |          |          |          |
| ectrometer* |            |        |          |          |          |
+-------------+------------+--------+----------+----------+----------+
| *Pyrolytic  | Analyze    | 4.0    | 8 (15    | N/A      | 2.5-0.6  |
| GCMS*       | inorgan    |        | peak)    |          | M        |
|             | ic/organic |        |          |          | b/sample |
|             | compounds; |        |          |          |          |
|             | isotope    |        |          |          |          |
|             | ratio and  |        |          |          |          |
|             | chirality  |        |          |          |          |
|             | det        |        |          |          |          |
|             | ermination |        |          |          |          |
+-------------+------------+--------+----------+----------+----------+
| *Chirality* | Identify   | N/A    | N/A      | N/A      | N/A      |
|             | amino      |        |          |          |          |
| *           | acids and  |        |          |          |          |
| Recognition | determine  |        |          |          |          |
| and*        | chirality  |        |          |          |          |
|             |            |        |          |          |          |
| *Analysis*  |            |        |          |          |          |
|             |            |        |          |          |          |
| *Package*   |            |        |          |          |          |
+-------------+------------+--------+----------+----------+----------+
| *Passive    | Detect     | N/A    | N/A      | N/A      | N/A      |
| Gas         | r          |        |          |          |          |
| Exchange    | espiration |        |          |          |          |
| Experiment* | of         |        |          |          |          |
|             | possible   |        |          |          |          |
|             | microbes   |        |          |          |          |
|             | in soil    |        |          |          |          |
|             | and rock   |        |          |          |          |
+-------------+------------+--------+----------+----------+----------+

\*adapted from the National Research Council, 2003

## Secondary Objective: Measure Atmospheric conditions

### Experiment

Because of its geographic situation, its geological history and its
present structure, Valles Marineris proved to be a warm, humid area,
with high pressure compared to the rest of the surface of Mars. This is
due to the facts that the bottom of the valley is 7 km deep and that
Valles Marineris lay in low latitudes. Clouds have even been observed in
the early morning (Figure 8). Thus, it would be interesting to measure
temperature, pressure and humidity of the atmosphere for three reasons.
First, it would enable us to validate the hypothesis of water under
liquid phase in this region. Secondly, it would represent a
discriminative element as far as the possibility of life is concerned
and thus help interpret the results of the other experiments
(spectrometers and biological package). A very low temperature would
tend to indicate that any form of life could not survive in this
environment. Finally, the temperature and pressure of the atmosphere are
some key parameters for the biological experiments, since they involve
some temperature and pressure dependent chemical reactions. We need
these parameters to interpret their results.

### Instrumentation

The meteorological package will allow the in-situ measurements of
weather conditions on the surface of Mars. They will measure
temperature, pressure, wind direction and speed, and the humidity of the
atmosphere, and thus work towards a better understanding of all the data
collected by a wide range of experiments.

  -------------------------------- ----------------------------------------------
  []{#__RefHeading___Toc81459601   
  .anchor}Table 6: Instrument      
  Specifications from Mars 96      
  penetrator^21^                   

  Temperature                      140K to 300K (accuracy 1%)

  Pressure                         3 to 10 mbar (accuracy 1%)

  Wind velocity                    1 to 50 m/s (accuracy 5%)

  Humidity                         10 to 1000 ppm (accuracy 10%)

  Data capacity                    \< 16 kbit/day (waiting mode 1.6 kbit/day)

  Power consumption                \< 1 W (mean)

  Mass                             0.4 kg

  Session time                     first day continuous, then periodically
  -------------------------------- ----------------------------------------------

### Data Analysis

Once the meteorological data has reached the Earth, a scientific team
should interpret them in relation to both the chemical and biological
experiments atmospheric meteorological parameters may be related with,
and to the search for liquid water.

The meteorological parameters measured in Valles Marineris are essential
to validate or invalidate the hypothesis of present life on Mars. Thus
several scientific teams (chemical, biological and physics) should work
together to interpret the results.

A computation of the influence of the meteorological parameters on the
way chemical reactions and biological processes work may be necessary to
interpret the results of the experiments.

## Tertiary Objective: Obtain Photographic Images of the Landing Site

### Experiment

Satellite images of Mars shows that there has been sedimentation
processes in the region of Valles Marineris in the past that were
enabled by flowing water. The gullies observed in Valles Marineris are
evidence for the past presence of water in Valles Marineris. Another
point worth mentioning is that volcanism contributed to the formation of
the valley. Thus, the camera would be the perfect tool to perform visual
analysis of the rocks. It could especially contribute to validating the
assumption that liquid water used to flow on Mars through the
observation of canals and gullies patterns.

The camera should help study the mineralogy of the landing site, as well
as provide information on the structure and the stratification of the
cliff of Valles Marineris, and thus help determine the geological
history and the chemical composition of the area. For instance, pictures
taken by the Viking orbiters have been key elements to determine the
geological history and the formation of Valles Marineris. They have
enabled the observation of new features. They indicate that, although
erosional landforms (such as landslide scars and deposits) and tributary
canyons are common, faulting apparently has been the dominant factor in
canyon development.

Since discovery of Valles Marineris, the method of its formation has
been a nagging puzzle. The canyons do not form a well-integrated
drainage system. Some are completely closed depressions, and lateral
transport by wind or water would be considerably impeded. Now, however,
the new evidence of faulting suggests that most negative relief results
from subsidence. Low, straight scarps, which apparently indicate
downward subsidence of canyon floors along faults, cut across erosion
features on many canyon walls. Similar scale faulting occurs on Earth:
in East Africa the continental crust is in tension across large rift
valleys. Erosion of the Valles Marineris walls apparently continued into
the recent past. The crustal tension causing the faulting within the
canyons may thus also have been a relatively recent phenomenon.

Another exciting discovery resulting from Viking images is the presence
of thick-layered deposits on the floors of several canyons. Layered rock
is also visible in the canyon walls, and thus is part of the Martian
crust predating the canyons. Some materials on the canyon floors are
distinctive for the fine scale and regularity of their layering. Only
climatic modulation of a sedimentary process seems adequate to explain
them. Possibly, regular changes in the Martian climate, governed by
known orbital variations, have controlled the level of dust storm
activity and the rate of deposition of sediment from the atmosphere.
Another theory is that some sections of Valles Marineris were sites of
lakes in which layered sediments were deposited. Before Viking,
regularly layered deposits were known only in the polar regions of Mars,
and their creation may also be associated with cyclical climatic change.

Finally, solar observations could be performed to measure water vapor
and atmospheric dust density. Multi-spectral observations of Phobos and
Deimos relative to the background of stars may enable us to determine
the position and orientation of the lander. Furthermore, it is essential
to know what kind of rock or soil is sampled and analyzed to interpret
the results. We have some knowledge of the rocks that are the most
likely to host life. For example, chert, siliciclastic and evaporite are
the most favorable to life on earth. The camera is perfect for
determining what kind of rock is analyzed.

### Instrumentation

The camera should provide a wide-angle (48° for the camera of Beagle
2^22)^ multi-spectral stereo imaging of the landing site, with high
resolution (0.75 milliradians for the Beagle 2 camera).

  -------------------------------- ------------------------------------------------
  []{#__RefHeading___Toc81459602   
  .anchor}Table 7: Instrument      
  Specifications for the camera of 
  Beagle2 ^22^                     

  Mass                             0.360 kg

  Volume                           747 cm³

  Power                            1.8 W average consumption

  RS422 bus                        10Mbits/s

  Pixels                           1024x1024
  -------------------------------- ------------------------------------------------

### Data Analysis

There are many ways to analyze the image data to get information on
various aspects of the Martian topography. The stereoscopic features of
both the Beagle 2 camera and the Mars 96 camera allow the creation of
three dimensional images, maps, and digital terrain models for studying
the structure of the Martian terrain. Stereoscopic images of the cloud
coverage will provide useful information about the various heights of
the cloud levels in the Valles Marineris region.

The camera has 4 different filters on a color wheel: blue, green, red,
and near-IR. Analyzing these images tells the spectral characteristics
and thus the composition and mineralogy of the rocks. The atmosphere
similarly can be studied in multiple wavelengths to determine the
composition and dust density.

In addition, the camera has the capabilities of multi-phase angle
imaging, which provides information on how the different wavelengths of
light are scattered. This yields the physical properties of the surface
and atmosphere as well as the weathering of the surface materials.
Multi-phase angle imaging of the atmosphere will yield the scattering
properties and the size of the dust particles in the atmosphere.

![](media/image15.png){width="2.3680555555555554in"
height="2.2465277777777777in"}

Figure 8: Mars Global Surveyor image of clouds in Valles Marineris from

Astronomy Picture of the Day

# Engineering Challenges and Solutions

## Overview and Introduction

The purpose of the engineering team is to place the scientific
instruments in a position where they are able to collect the data
required in a safe, reliable, and stable environment and return the data
safely to Earth. Although this goal is simple, it encompasses a wide
variety of challenging engineering problems.

An eight person engineering team has been developed to examine the major
factors involved in sending a scientific probe to Mars. The details of
this team and their results are presented in this section of the report.

## Driving Requirements

During the design's progression, a number of factors have emerged as
drivers. The chief driver is the achievement of the science objectives.
All of the subsystems serve only as support equipment for these
objectives. Several specific aspects have been discovered as being at
the forefront of the discussion in order to achieve the desired science.
These aspects are discussed below.

### Drill Depth {#drill-depth .unnumbered}

One of the most critical aspects of the mission is the drill's depth.
From a scientific perspective, the deeper the drill can dig into the
ground, the more valuable the science, due to the larger range of
samples that will be examined. However, increasing the drill depth
significantly increases the complexity, cost, and mass of the drilling
mechanism. A drilling depth of 3 meters has been selected with a minimum
depth of 0.5 meters and a maximum depth of 6 meters.

### Soil Material {#soil-material .unnumbered}

Understanding the ground composition in which the drill will operate is
essential to determining the drill's performance parameters. The drill's
performance is especially driven by the hardness of the rocks and ground
being mined. It also comes into effect, however, due to the fact that if
the soil is loose then there is a possibility of the hole collapsing. To
prevent this, the hole itself must be reinforced. Knowing the soil
composition at the landing site is not easy and is a continuing question
of scientific inquiry. Thus, the drill must be able to operate in a
range of terrains.

### Sample Acquisition and Processing {#sample-acquisition-and-processing .unnumbered}

Once the hole has been drilled, a sample must be returned to the
instruments on the lander. During this process, it is essential that the
sample not become contaminated. Furthermore, since the drill is creating
heat, it must be ensured that the amount of heat produced by the drill
is not sufficient to harm the integrity of the sample.

### Reliability and Safety {#reliability-and-safety .unnumbered}

For the scientific objectives to be successful, it is very important
that the system be reliable and thoroughly tested. Adherence to these
goals will permeate the design spiral.

## Chief Design Trades

During the design process, a number of design trades were considered in
light of the requirements and criteria that were given. Some of these
design trades are highlighted below.

### Drill Selection {#drill-selection .unnumbered}

The drill is the central piece of technology on the spacecraft, and the
rest of the lander is essentially built around it. Thus, the selection
of the type of drill to use was crucial. A wide variety of drill types
were considered. These included laser drills, ultrasonic drills, pile
drivers, moles, penetrators and rotary percussion augers. Eventually,
the rotary percussion auger was selected as being the most reliable,
robust, and mature technology.

### Solar Panel Arrays {#solar-panel-arrays .unnumbered}

There was some initial discussion about the merits of having solar
panels that tracked the sun verses stationary solar panels. Moving solar
panels would be able to track the sun for a longer period of time and
obtain more favorable angles. Stationary solar panels are mechanically
much simpler. It was determined through analysis that the amount of
power used by the tracking array would be on the order of the amount of
power gained. Furthermore, such a system would expose the system to more
risk. For these reasons, stationary solar panels were selected[^29].

### Rover versus Lander {#rover-versus-lander .unnumbered}

Initially, the decision needed to be made as to the importance of
mobility to achieving the scientific objectives. It was quickly
determined that the limited mobility achieved through the use of a rover
would not justify the increased cost and complexity that significant
access to diverse terrains would entail.

### Descent Style {#descent-style .unnumbered}

Two different methods of landing on the planet were considered. The
first was a retrorocket landing similar to that was used on Viking, was
planned to be used on Mars Polar Lander, and will be used on Mars
Phoenix. The second was an airbag system similar to that used by the
Mars Exploration Rovers and the Mars Pathfinder mission. The two systems
were of varying masses and complexities. Both are flight proven. The
determining factor in this decision was mass. The mission budget was
driven by mass and the retro-rockets proved to be too costly; therefore,
the airbag was chosen for the MSCLE mission.

## Present Status

Currently, a preliminary design for the MSCLE spacecraft has been
developed. Each of the subsystem teams has contributed to the concept of
their specific components and systems. Every subsystem's design has been
conducted in as much detail as possible considering the short time frame
in which to plan out the spacecraft and its mission.

## Subsystem Analyses

Below are the summaries of effort and current technical progress of each
subsystem.

### Structures

The general structural characteristics of the three main phases of the
mission include: (1) the transit satellite, (2) the landing system, and
(3) the platform. These characteristics include mass, dimensions, shape,
and structural materials, which are discussed in more detail below. Many
features are similar to those found on the Mars Exploration Rover (MER)
spacecraft, which will decrease the risk factors of the MSCLE mission.

#### Transit Satellite

The spacecraft will be encased in the payload fairing of the launch
vehicle as it travels to Mars. The payload fairing will contain all
components necessary to power, protect, and guide the spacecraft. The
cruise stage that controlled the transit of missions such as Mars
Pathfinder and Viking missions will not be necessary. All of the
functions of this cruise stage will be embodied in the payload fairing.
The protective fairing will release the spacecraft once it is near Mars.
Figure 9 and Figure 10 provide views of the MER spacecraft that
descended through the Martian atmosphere. The MSCLE craft will be
similar in design.

![](media/image16.png){width="3.3958333333333335in"
height="2.432638888888889in"}

[]{#__RefHeading___Toc81459550 .anchor}Figure 9: MER Descent
Configuration[^30]

![](media/image17.png){width="3.03125in" height="3.060416666666667in"}

[]{#__RefHeading___Toc81459551 .anchor}Figure 10: MER Descent
Configuration - Exploded View[^31]

#### Landing System

The structures subsystem analyzed two landing systems: retro-rockets and
airbags. Originally, the MSCLE mission was going to land softly on Mars
via the retro-rockets. However, this method proved to be too costly in
terms of mass, and the airbag system was chosen for the MSCLE mission.

Before the transit satellite reaches the Martian atmosphere, the payload
fairing will separate from the rest of the vehicle. After entering the
atmosphere, the supersonic parachute, coupled with an aerobraking heat
shield, will deploy and the spacecraft's speed will decrease through the
atmosphere. At a certain altitude (undetermined as of yet), the heat
shield will break off and the lander will separate from the rest of the
craft. Small retro-rockets will decrease the speed even more until huge
airbags are deployed around the entire lander. The lander will bounce on
the Martian surface until it comes to a stop. Figure 11 is a pictorial
demonstration of the landing sequence that the MSCLE craft will undergo.

![](media/image18.jpeg){width="3.375in" height="2.5590277777777777in"}

[]{#__RefHeading___Toc81459552 .anchor}Figure 11: Airbag Landing
Sequence[^32]

The airbag landing sequence is similar to the landing methods used by
the MER and Pathfinder missions. This technology has flown successfully
in the past and is relatively low-risk.

#### Platform

A simple onboard accelerometer will be able to measure the Martian
gravity, which a computer can easily use to calculate at what angle the
rover may be sitting. If the lander comes to stop in an inclined
position, the scientists can use the accelerometer data to determine the
best means of righting the rover. One method may be that the petals will
unfold and flip the lander over, off the rock or inclined surface. If
the lander is unable to lay flat on the ground, then the mission will be
carried out as completely as possible at an angle.

![](media/image19.jpeg){width="3.6875in" height="2.6381944444444443in"}

[]{#__RefHeading___Toc81459553 .anchor}Figure 12: Mars Pathfinder Lander
Configuration[^33]

The machine will drill a hole into the Martian rock by means of a rotary
percussion auger. The Martian rock material will slide up and out of the
hole as the auger spins; in this method, the auger will transport
material from various depths into the lander in order to conduct the
necessary scientific research. A diagram of the Mars Pathfinder Lander
configuration is provided in Figure 12.

The lander will be constructed out of honeycomb aluminum, much like
several past Mars missions. Steel or titanium may be used for structural
elements if more strength or support is needed for the heavier
scientific instruments and tools, such as the drill.

The drilling mechanism will be mounted on top or within the landing
platform, creating a need for the drill to go through the bottom of the
lander before it reaches the ground. A hole through the lander will be
provided for this purpose; during transit and descent, the hole will be
shielded with a flap or door mechanism of sorts that expose the hole
when the lander is ready to drill. One possible problem with the drill
reaching the soil may be caused by the deflated airbags. Once deflated,
the airbags will be pulled out of the way as much as possible, using a
mechanism similar to the MER lander. If need be, mass in the lander may
be placed such that the center of gravity will cause the lander to roll
to a stop on its side rather than right-side-up. Then the air bags
covering the bottom surface of the lander can be easily detracted, and
the petals can flip the lander right side up, leaving the bottom side
completely free of any airbags.

*Mass Budget*

In addition to the design of the spacecraft, the structures subsystem
has set up a mass budget[^34]. The minimum and desired mass ranges of
the individual subsystem masses are located in Error: Reference source
not found.

  ----------------------------------------- -----------------------------
  []{#__RefHeading___Toc81459603            
  .anchor}*Table 8: Subsystem Masses*       

  ***[Subsystems]{.underline}***            *[Desired Mass
                                            (kg)]{.underline}*

  Propulsion                                50

  Attitude and Control                      2

  Command & Data Handling/Communications    5

  Thermal                                   3.72

  Power                                     11

  Atmospheric Entry                         61.29

  Structures & Mechanisms                   37.24

  TOTAL                                     170.25
  ----------------------------------------- -----------------------------

The science team has supplied the structure subsystem with the known
masses for the payload. These values are located in Error: Reference
source not found.

  ------------------------------------------- ---------------------------
  []{#__RefHeading___Toc81459604              
  .anchor}*Table 9: Payload Masses*           

  ***[Device]{.underline}***                  *[Mass (kg)]{.underline}*

  Drill                                       20

  Science Instruments                         20

  TOTAL                                       40
  ------------------------------------------- ---------------------------

In addition, the dry mass was calculated. The dry mass is defined as the
mass of the spacecraft when it lands on the Martian surface; this number
does not take into account the heat shield and other components that are
released during the spacecraft's descent. The spacecraft dry mass is
calculated to be approximately 186.21 kilograms. The spacecraft dry
mass, along with other calculated masses, is presented in Error:
Reference source not found. Effort will be made to reduce the mass of
the spacecraft as much as possible.

  ----------------------------------------- -----------------------------
  []{#__RefHeading___Toc81459605            
  .anchor}*Table 10: Further Mass           
  Calculations*                             

  ***[Component]{.underline}***             *[Desired Mass
                                            (kg)]{.underline}*

  Payload                                   40

  Subsystems                                179.57

  Margin                                    37.24

  Dry Mass                                  186.21

  Propellant                                50

  Loaded Mass                               236.21

  Kick Stage                                105

  Injected Mass                             341.21

  Adapter                                   145.90

  Boosted Mass                              487.11
  ----------------------------------------- -----------------------------

The calculations used in determining the above masses can be found in
Appendix A. Assuming an average spacecraft density of 79 kilograms per
cubic meter yields a spacecraft volume of approximately 3.77 cubic
meters[^35]. Details of the rough volume calculation are located in
Appendix A as well.

*Future Tasks*

In the future, the structure subsystem will need to plan out where all
of the individual subsystem components must be placed within the
spacecraft to locate the center of gravity in the center of the craft.
In addition, a method of protecting the more sensitive equipment from
the sun's radiation is required; blankets will be wrapped around the
components that require radiation and thermal protection. Eventually,
the various forces acting on the spacecraft will need to be analyzed,
and the size of any structural stiffeners located throughout the
spacecraft will need to be determined.

### Payload

The payload delivery system serves as a link between the environment and
the instruments. This includes collecting and transporting sub-surface
samples, as well as atmospheric samples; however the atmospheric
instruments sit on top of the lander and do not require separate
collection or transport systems. The primary goal is to collect and
analyze a 2-m deep sub-surface sample. The ground material is expected
to be a combination of alluvial soils and basalt.

#### Drill

Drilling on Mars poses three key challenges. First, the drilling process
cannot destroy or alter the chemical properties of the sample.
Therefore, continuous lasers and other vaporizing methods cannot be used
unless the sample is taken at least three diameters away from the core.
Next, traditional lubrication methods may not be used for cooling and
transporting slurry due to mass and contamination issues. Finally, the
sample must be transported to the surface for study. Other engineering
constraints include the size and complexity of the device---and related
risk---its power requirements, robustness, and its ability to complete
the mission requirements in ninety days.

*Overview of Drilling Methods*

A LANL report[^36] examined several deep drilling methods (on the order
of 200 m) on Mars with a power budget of 1000 W per day. Of these, the
methods that best matched the science requirements for providing an
uncontaminated core were a hollow-stemmed auger and sonic or ultrasonic
drills. The auger is a more proven technology---Apollo astronauts used
an auger to collect 3-m deep samples on the moon---but requires more
mechanical parts. A JPL program has developed a miniature ultrasonic
drill that runs at 24 W for shallow (on the order of 10 cm) Martian work
and suggests that scaled versions may be available for the current
mission needs[^37]. Variants of the 'mole'---a burrowing sampler sent on
Beagle 2---were considered, but this method works only for regolith, not
rock[^38]. Finally, laser-based drilling methods were excluded due to
their high power requirements.

Commonly used on earth, the auger is the best understood and tested
method. Drawbacks include the risk of bit binding or breaking, and the
mass required to extend the stem at greater depths. However, much
research has been completed on bit behaviors, and the mass required to
drill to 2 m should not be prohibitive. The choice of a
rotary-percussion auger as the best tool to obtain subsurface samples
was supported by JPL research [^39] that tested various drilling
methods---including methods to extend the auger stem---on basalt
samples.

![](media/image20.png){width="1.3541666666666667in"
height="1.9270833333333333in"}

[]{#__RefHeading___Toc81459554 .anchor}Figure 13: Honeybee 10 m
Drill[^40]

*Current State of Auger Research*

Several companies and government agencies are developing 10-m augers for
planetary exploration. Honeybee Robotics[^41] (creators of the rock
abrasion tool used on the Mars Exploration Rovers), is developing a 10-m
drill based on a 1-m drill (see Figure 13). With a current Technology
Readiness Level of 4/5, Honeybee has tested in Arizona and a New York
quarry to depths of 7 m with a 8 mm core sample (Figure 14). A compact
rotary device extends low-friction nickel-Teflon drill sections in
seamless segments. Each section is approximately 0.5 m long; as the
drill digs deeper into the ground more sections are added to extend the
auger's stem between the bit and the motor. Thus the 0.6 m tall drill
can achieve depths of 10 m.

![](media/image21.png){width="2.997916666666667in"
height="1.9458333333333333in"}

[]{#__RefHeading___Toc81459555 .anchor}Figure 14: Honeybee Core Sampling
System[^42]

Swales Aerospace[^43] is also developing a 10-m drill. In December 2002
they demonstrated its ability to 10 m through abrasive red silica
sandstone without bit replacement or repairs. They have optimized their
automation, sample-capture, and mechanical systems to minimize heating
(thus protecting the bit); however, all parts---motors, gearboxes, and
electronics---are all commercially available.

A team from NASA and Baker-Hughes has conducted laboratory tests on
autonomous drilling to a depth of 2 m through sandstone (with a
2-cm-diamter core). Their autonomy software monitors motor currents,
drill rotation and penetration rate, and various temperatures to control
the drill (tested with the Payzone drilling simulator). Other players
include Norcat's Canadrill[^44], shown in Figure 15 and Figure 16. The
Canadrill is a lightweight autonomous drill originally planned for
launch in 2009 (whose team has completed detailed research on autonomy
and drill behavior[^45]) and Guigne, who is developing an advanced bit
that could be used for Martian applications[^46].

![](media/image22.png){width="1.8902777777777777in"
height="2.8208333333333333in"}

[]{#__RefHeading___Toc81459556 .anchor}Figure 15: Canadrill[^47]

![](media/image23.png){width="2.75in" height="1.375in"}

[]{#__RefHeading___Toc81459557 .anchor}Figure 16: Canadrill Lander
Concept[^48]

*MSCLE Drill*

Drawing on publicly available data on the drills surveyed above, the
MSCLE drill should weigh between 15-25 kg and be approximately 15 cm by
40 cm by 60 cm high[^49]. The drill is expected to consume between 80 W
and 120 W, averaging 90 W of electric power for normal operations[^50]
(about one third of that would be converted to mechanical power). The
drill will be capable of drilling between 8 cm and 15 cm an hour[^51].
Data on whether currently produced auger-based drills could survive an
air-bag impact remains to be found.

For the science, no drill temperatures in contact with the samples may
exceed 573 K (or biological molecules will degrade). Companies have
developed several (proprietary) methods to minimize heat production;
here a rough calculation shows that the drill temperature should remain
below our limit. The day surface temperatures in the Vallis Marineris
peak at 308 K. For the calculation, it is assumed that a 2 m sphere
around the bit stays at 308 K, that all heat is dissipated through a 1
cm sphere in the bit, and that all heat is conducted through the basalt
(and not through the nickel stem). Then Fourier's Law of Heat
Conduction, which states that heat transfer is proportional to the
conductivity and the temperature difference, can be used to attain an
upper bound on the temperature as 425 K, safely under the 573 K limit.

Because Earth and Mars are approximately 20 light-minutes apart, any
round-trip data and response will take at least 40 minutes (ignoring
decision-making time on the ground). Drills can bind or break in a much
shorter time span; therefore the drill must be able to operate
autonomously. By monitoring parameters such as the motor currents and
temperatures, the drill will identify abnormalities and halt execution.
If a sequence of escape maneuvers (such as operating the drill in
reverse) does not solve the problem, the drill will go into a safe-mode
and report the problem to Earth for trouble-shooting.

Normal operations will consist of several permission-to-advance points.
First, the predicted drill parameters will be checked before drilling
starts to calibrate and ensure functionality. On the second day, the
drill will proceed to 2-cm and 10-cm checkpoints. Then it will continue
to 1-m and 2-m checkpoints on the following days drilling at a rate of
.5 m per day. Samples will be collected and tested at 10 cm, 1 m, and 2
m; two additional days will be allotted between each checkpoint for
transporting the samples up the drill and running the science
instruments. While the primary mission objective is to drill to 2-m, the
drill will carry enough drill sections to go to a maximum depth of 8 m.

#### Sample Transport

Honeybee robotics also builds transport systems to deliver core samples
to the instruments. Their system can handle samples of .07 cc, which can
be hermetically sealed to the instrument. Each sample can be delivered
with three degrees of freedom to .0001-inch precision (see Figure 17).

![](media/image24.jpeg){width="1.3520833333333333in"
height="1.9270833333333333in"}

[]{#__RefHeading___Toc81459558 .anchor}Figure 17: Honeybee Sample
Transport System[^52]

### Power Production and Distribution

The power production and distribution subsystem is primarily concerned
with the production, storage, regulation, and distribution of electrical
power to all of the necessary components onboard the MSCLE Lander. The
pages to follow describe the major divisions and design of the Lander's
power system, and therefore only loosely present the actual design of
the system's components that will be necessary for flight readiness. A
majority of the MSCLE Lander's power system is based on that of the
current Mars Exploration Rovers (MERs), and because of their success,
this design should provide the necessary flight worthiness rating.

#### Additional Research

This subsystem report mainly concentrates on the design of the Lander's
solar panels, battery system, and thermal protection of the necessary
electronics. Additional research on the power distribution and control
bus remains. Many of the spacecraft control buses used today are
somewhat similar and only their load handling and battery charging
capabilities change from system to system.

#### System Composition

As mentioned earlier, the electrical power subsystem is divided into
four fundamental sections. The following block diagram (Figure 18) shows
a pictorial representation of the system.

![](media/image25.jpeg){width="4.441666666666666in"
height="1.4583333333333333in"}

[]{#__RefHeading___Toc81459559 .anchor}Figure 18: Block Diagram of
Electrical Power Subsystem Hierarchy

The power source pictured in the block diagram will constitute the power
production source for the Lander which is solar power, and the energy
source refers to the lithium-ion batteries that will be used onboard.
The power regulation and control refers to the dc-dc converter that will
convert the solar generated voltage to a standard, tightly regulated
voltage that will be carried to the necessary locations on the Lander by
the power distribution block components. These will include voltage
buses, wiring harnesses, and connectors.

#### Modes of Operation

During the course of the entire mission, the Lander will undergo five
different modes of operation from a power standpoint. They are Cruise
(Error: Reference source not found), Science (Error: Reference source
not found), MarsCOM (Error: Reference source not found), EarthCOM
(Error: Reference source not found), and Standby (Error: Reference
source not found)[^53]. Each of the modes corresponds to a different
power consumption model and only provides the necessary electricity to
power the instruments and/or components needed during that period. The
following table (Error: Reference source not found) describes the modes,
their time spans, and their purposes. Following Error: Reference source
not found is a figure (Figure 19) and equations that show how the
average day for science in the Valles Marineris was computed. Next, the
times needed for communication were provided by the communications
subsystem and the rest of the day was portioned for standby mode. The
cruise time is based on information received from the flight dynamics
subsystem.

+-----------+------------+---------------------------------------------+
| []{       |            |                                             |
| #__RefHea |            |                                             |
| ding___To |            |                                             |
| c81459606 |            |                                             |
| .anc      |            |                                             |
| hor}Table |            |                                             |
| 11:Modes  |            |                                             |
| of        |            |                                             |
| Operation |            |                                             |
| for       |            |                                             |
| M         |            |                                             |
| SCLE[^54] |            |                                             |
+-----------+------------+---------------------------------------------+
| ***[Mode  | *[Total    | *[Purpose]{.underline}*                     |
| of        | Time       |                                             |
| Operati   | Span]{.u   |                                             |
| on]{.unde | nderline}* |                                             |
| rline}*** |            |                                             |
+-----------+------------+---------------------------------------------+
| *         | ≈ 8 months | Power during cruise stage to Mars after     |
| *Cruise** |            | leaving earth's orbit                       |
+-----------+------------+---------------------------------------------+
| **        | ≈ 8        | Power while on Mars to perform science      |
| Science** | hr./day    | objectives                                  |
|           |            |                                             |
|           | X 90 days  |                                             |
+-----------+------------+---------------------------------------------+
| **        | ≈ 8        | Power used to transmit collected data to    |
| MarsCOM** | min./day X | Mars orbiting satellite when possible       |
|           | 90 days    |                                             |
+-----------+------------+---------------------------------------------+
| **E       | ≈ 10       | Power used to transmit collected data to    |
| arthCOM** | min./day X | Earth when visible and Mars satellite isn't |
|           | 90 days    | visible                                     |
+-----------+------------+---------------------------------------------+
| **        | ≈ 16.65    | Power required to perform the necessary     |
| Standby** | hr./day    | life functions on the rover when batteries  |
|           |            | are the primary energy source because of    |
|           | X 90 days  | little or no sunlight                       |
+-----------+------------+---------------------------------------------+

$$

![](media/image26.png){width="4.759722222222222in"
height="1.5833333333333333in"}

[]{#__RefHeading___Toc81459560 .anchor}Figure 19: Graphic Representation
Used to Solve for Average Day Time

The angle α was estimated to be 45° to allow for contingency so that the
maximum amount of sunlight that the Lander sees on the worst day is
expected everyday. If the weather is good then more than 6 hours of
energy time can be used, but the safest bet is to consider the worst
case scenario.

The following tables detail each of the operating modes of the Lander
and the breakdown for each mode.

+------------------------+----------------------------+---------------+
| []{#__Re               |                            |               |
| fHeading___Toc81459607 |                            |               |
| .anchor}Table 12:      |                            |               |
| Cruise Mode            |                            |               |
| Requirements           |                            |               |
+------------------------+----------------------------+---------------+
| *[Instruments          | *[Power (W)]{.underline}*  | *[% Used      |
| Used]{.underline}*     |                            | During]       |
|                        |                            | {.underline}* |
|                        |                            |               |
|                        |                            | *[Mode]       |
|                        |                            | {.underline}* |
+------------------------+----------------------------+---------------+
| Power Reg/Cond/Charger | 10.75                      | 100           |
+------------------------+----------------------------+---------------+
| Inertial Ref. Unit     | 20                         | 100           |
+------------------------+----------------------------+---------------+
| Star Tracker           | 5                          | 100           |
+------------------------+----------------------------+---------------+
| Sun Sensor             | 1                          | 100           |
+------------------------+----------------------------+---------------+
| CDH Computer           | 7.5                        | 100           |
+------------------------+----------------------------+---------------+
| Propulsion             | 280                        | .1            |
+------------------------+----------------------------+---------------+
| Thermal                | 5                          | 100           |
+------------------------+----------------------------+---------------+
| *Time in Mode*         | *8 months*                 |               |
+------------------------+----------------------------+---------------+
| *Time w/out Light*     | *100 min. max each         |               |
|                        | occurrence*                |               |
+------------------------+----------------------------+---------------+
| *Avg. Power Used*      | *48.16 W*                  |               |
+------------------------+----------------------------+---------------+

+---------------------------------+---------------+--------------------+
| []{#__RefHeading___Toc81459608  |               |                    |
| .anchor}Table 13: Science Mode  |               |                    |
| Power Requirements              |               |                    |
+---------------------------------+---------------+--------------------+
| *[Instruments                   | *[Power       | *[%Used            |
| Used]{.underline}*              | (W)]          | Du                 |
|                                 | {.underline}* | ring]{.underline}* |
|                                 |               |                    |
|                                 |               | *[                 |
|                                 |               | Mode]{.underline}* |
+---------------------------------+---------------+--------------------+
| Power Reg/Cond/Charger          | 10.75         | 100                |
+---------------------------------+---------------+--------------------+
| CDH Computer                    | 7.5           | 100                |
+---------------------------------+---------------+--------------------+
| Auger                           | 90            | 100                |
+---------------------------------+---------------+--------------------+
| Mass Spec/Gaschromatic          | 20            | 50                 |
+---------------------------------+---------------+--------------------+
| APX Spec                        | 0.34          | 50                 |
+---------------------------------+---------------+--------------------+
| Biological Package              | 15            | 50                 |
+---------------------------------+---------------+--------------------+
| Meteorological Package          | 1             | 1                  |
+---------------------------------+---------------+--------------------+
| Camera                          | 1.8           | 1                  |
+---------------------------------+---------------+--------------------+
| Thermal                         | 5             | 0.25               |
+---------------------------------+---------------+--------------------+
| *Time in Mode*                  | *6.16 hrs.*   |                    |
+---------------------------------+---------------+--------------------+
| *Avg Power Used*                | *130 W*       |                    |
+---------------------------------+---------------+--------------------+

+---------------------------------+---------------+--------------------+
| []{#__RefHeading___Toc81459609  |               |                    |
| .anchor}Table 14: EarthCOM M    |               |                    |
| ode Power Requirements          |               |                    |
+---------------------------------+---------------+--------------------+
| *[Instruments                   | *[Power       | *[%Used            |
| Used]{.underline}*              | (W)]          | Du                 |
|                                 | {.underline}* | ring]{.underline}* |
|                                 |               |                    |
|                                 |               | *[                 |
|                                 |               | Mode]{.underline}* |
+---------------------------------+---------------+--------------------+
| Power Reg/Cond/Charger          | 4.25          | 100                |
+---------------------------------+---------------+--------------------+
| CDH Computer                    | 7.5           | 100                |
+---------------------------------+---------------+--------------------+
| High Gain Antenna               | 20            | 100                |
+---------------------------------+---------------+--------------------+
| Thermal                         | 5             | 100                |
+---------------------------------+---------------+--------------------+
| *Time in Mode*                  | *3 hrs.*      |                    |
+---------------------------------+---------------+--------------------+
| *Avg. Power Used*               | *46.75 W*     |                    |
+---------------------------------+---------------+--------------------+

+---------------------------------+---------------+--------------------+
| []{#__RefHeading___Toc81459610  |               |                    |
| .anchor}Table 15: MarsCOM Mode  |               |                    |
| Power Requirements              |               |                    |
+---------------------------------+---------------+--------------------+
| [Instruments Used]{.underline}  | [Power        | [%Used             |
|                                 | (W)           | D                  |
|                                 | ]{.underline} | uring]{.underline} |
|                                 |               |                    |
|                                 |               | [Mode]{.underline} |
+---------------------------------+---------------+--------------------+
| Power Reg/Cond/Charger          | 5.75          | 100                |
+---------------------------------+---------------+--------------------+
| CDH Computer                    | 7.5           | 100                |
+---------------------------------+---------------+--------------------+
| Mid Gain Antenna                | 20            | 100                |
+---------------------------------+---------------+--------------------+
| Low Gain Antenna                | 15            | 100                |
+---------------------------------+---------------+--------------------+
| UHF Antenna                     | 10            | 100                |
+---------------------------------+---------------+--------------------+
| Thermal                         | 5             | 100                |
+---------------------------------+---------------+--------------------+
| *Time in Mode*                  | *8 min.*      |                    |
+---------------------------------+---------------+--------------------+
| *Avg. Power Used*               | *63.25 W*     |                    |
+---------------------------------+---------------+--------------------+

+---------------------------------+---------------+--------------------+
| []{#__RefHeading___Toc81459611  |               |                    |
| .anchor}Table 16: Standby Mode  |               |                    |
| Power Requirements              |               |                    |
+---------------------------------+---------------+--------------------+
| *[Instruments                   | *[Power       | *[%Used            |
| Used]{.underline}*              | (W)]          | Du                 |
|                                 | {.underline}* | ring]{.underline}* |
|                                 |               |                    |
|                                 |               | *[                 |
|                                 |               | Mode]{.underline}* |
+---------------------------------+---------------+--------------------+
| Power Reg/Cond/Charger          | 1.63          | 100                |
+---------------------------------+---------------+--------------------+
| CDH Computer                    | 7.5           | 100                |
+---------------------------------+---------------+--------------------+
| Low Gain Amp                    | 15            | 100                |
+---------------------------------+---------------+--------------------+
| Thermal                         | 5             | 50                 |
+---------------------------------+---------------+--------------------+
| *Time in Mode*                  | *15.35 hrs.*  |                    |
+---------------------------------+---------------+--------------------+
| *Avg. Power Used*               | *17.88 W*     |                    |
+---------------------------------+---------------+--------------------+

As seen from the above tables, the power requirements are
straightforward. Notice the power regulator/conditioner/charger in each
mode. Each time it is needed the entire time to regulate the bus voltage
and distribution. However, its power consumption rate changes because of
an inherent dc-dc conversion loss and the changes in the loads. The
dc-dc conversion loss was estimated to be around 10% based on the design
used by Luther Richardson and Derek Inaba for a satellite mission
concept done for JPL[^55]. Another item worth noting in the tables is
the "%Used During Mode" column. This column is a percentage of the total
time the instrument will be used while the Lander is in that particular
mode.

#### Solar Panel Design

There are two major aspects to consider when designing a solar panel.
The first is to determine if a flat-panel array or a movable array will
be used and the second is to determine the type of solar cell that will
be used. For this mission a flat panel array has been chosen because of
the general consensus among many NASA engineers and scientists that
because of all the dust in the Martian air the sunlight will be
scattered at all angles and the little benefit received from movable
and/or focusing arrays isn't worth the weight and cost increase[^56].
The second design consideration is to decide the type of solar cell that
will be used on the array. Today, there are two main elements used for
solar cell construction - they are silicon and germanium. Numerous
different hybrids and combinations of these two exist. In this situation
however, the germanium cells are the best choice for the Lander's solar
panels because they are more efficient than silicon and more radiation
resistant. The one drawback to germanium cells is the price. They are 3
times more expensive than the silicon cells[^57], but are well worth it
for this particular type of mission.

The next part to the solar panel design is to determine the size
necessary to produce the power required. The first step is to figure out
how much power must be produced by the solar arrays. The following
equation shows how to accomplish this:

![](media/image27.wmf)

P~d~ is the power required during the day, P~e~ is the power during an
eclipse (no sunlight), T~d~ is the length of time where power will be
needed during the day and T~e~ is the length of the eclipse and X~d~ and
X~e~ represent the efficiency of the paths from the solar arrays through
the batteries to the individual loads and directly from the solar array
to the loads or equipment, respectively[^58]. When using the PPT (Peak
Power Tracker) power control technique (discussed in power distribution
section) X~d~ = 0.6 and X~e~ = 0.8. Using the highest average power from
the above tables (Table 3) and the average cruise power needed, the two
solar array powers can be computed. The Lander's array needs to produce
162.5 W and the cruise vehicle's array needs to produce 60.2 W raw
output. The next step is to determine the amount of power output that
the cells will produce on the Martian surface[^59]. The equation for
this is as follows:

![](media/image28.wmf)

In this equation P~o~ is the output power, C~eff~ is the cell efficiency
rating, and M~solin~ is the solar intensity on the surface of Mars[^60].
M~solin~ = 589.2 W/m^2^ and the power output from the cells produced is
166 W/m^2^. The efficiency of the cells chosen for this application are
28.3% and made by Spectrolab Photovoltaic Products[^61]. These cells are
the new generation of the current cells used on the Mars rovers. Next,
the beginning of life (BOL) and ending of life (EOL) power production
capability of the cells is determined and then applied to final
necessary area of the solar array. The array is sized according to the
EOL power so that the mission won't be compromised after the cells have
degraded. After making all of these calculations, the size of the array
for Lander will be approximately 2.3 square meters and the array for the
cruise vehicle will need to be 0.86 square meters [^62].

#### Battery System

The batteries on the vehicle need to be sized according to the power
that will be needed during the times when no sunlight is available for
the Lander. A primary battery will be used for the cruise stage and a
secondary/rechargeable battery will be used for the Lander while it is
on Mars. The battery for the cruise vehicle only needs to be large
enough to provide the cruise vehicle with power for approximately 100
minutes. This is the estimate time that the vehicle will be facing away
from the sun. Therefore the battery for the cruise vehicle will be a
10.5 Ah 14.4V lithium ion cell made by Lithion Inc. The battery for the
Lander will be sized according to the Standby Mode requirements and will
be a 35 Ah 14.4V lithium ion cell made by the same company. Both of the
battery requirements were sized according to the guidelines set forth in
[Space Mission Design and Analysis]{.underline}[^63]. Two batteries
connected in series will be required for both vehicles to create the 28V
system bus. Lithium-ion cells offer a significant energy density
advantage and operate over a wider temperature range than previous space
rated battery cells. Their energy density is in the 70-110 W\*hr/kg
range as opposed to the 25-57 W\*hr/kg density range of nickel-cadmium
and nickel-hydrogen cells[^64]. Lithion, Inc. produced the batteries for
the current Mars rovers (MER), and thus, their space-rated battery
performance has been tested and verified.

*Power Distribution and Control*

As mentioned earlier, the power regulator/conditioner/charger was
modeled after a JPL concept project developed by Luther Richardson and
Derek Inaba. It will have 90% dc-dc conversion efficiency and will
provide the necessary control and distribution for all 5 modes of
operation. The system voltage will operate at 28V DC and will implement
the Peak Power Tracker (PPT) control system. The PPT control system is
well suited for missions under 5 years in length. It is a
non-dissipative system and only extracts the exact amount of power that
the spacecraft requires up to the peak power capability of the arrays,
as opposed to the Direct Energy Transfer (DET) which dissipates the
additional power produced.

*Thermal Considerations*

All of the power system components will housed inside of a warm
electronics box (WEB) to keep them in the correct operating environment
of -40°C to 40°C. This type of box was also flown on the Mars rovers
(MER) and contains 8 radioisotope heating units that provide 1 watt of
heat each. In addition an electric heater is also included to provide
extra heating if necessary. The box is lined with aerogel for insulation
and painted gold to prevent heat from escaping[^65].

### Flight Dynamics

Transporting the lander from the surface of Earth to the surface of Mars
requires a great amount of time and energy input. The mission consists
of three primary sequential flight phases: (I) launch from Earth, (II)
cruise to the orbit of Mars, and (III) arrival at Mars. Each of these
phases include different flight and trajectory considerations, but all
are subject to the common laws of two-body orbital mechanics. The most
flight time is spent during the cruise (*phase II*) , where any
interplanetary transfer planning begins.

#### Phase II Mechanics

Cost remains a major consideration for any major exploration venture.
For interplanetary travel, fuel consumption accounts for a large amount
of the mission cost. A Hohmann transfer between the Earth and Mars
requires the least amount of energy input, so employing that transfer
instead of a direct transfer reduces the cost of flight during the
cruise phase. Direct transfers offer a sooner arrival time for the
spacecraft, but because the MSCLE spacecraft will not carry any
biological or short-lifetime components, a fast transfer does not
justify the high cost of a higher energy input.

Figure 20 below shows the Hohmann transfer (green line) for the
spacecraft from Earth's orbit to Mars' orbit.

![](media/image29.jpeg){width="3.28125in" height="2.957638888888889in"}

[]{#__RefHeading___Toc81459561 .anchor}Figure 20: Hohmann Transfer for
Cruise

The transfer requires that Earth and Mars be at exactly opposite sides
of the Sun in a particular way. The Earth must be in the location shown
above when the spacecraft leaves Earth's gravitational influence, and
likewise for Mars when the spacecraft enters Mars' gravitational
influence. It is important to note that the location of the planets as
shown does not occur at the same instant of time. Knowing the distance
between the Earth and Mars in the configuration shown allows one to
calculate the necessary heliocentric (Sun-centered) velocity of the
spacecraft on both the Earth side (periapsis) and Mars side (apoapsis)
of the transfer orbit. The spacecraft need not generate all of the
required periapsis velocity to begin its transfer. Since the Earth has
some upward velocity as shown in the diagram and the spacecraft begins
its flight within the Earth's atmosphere, the spacecraft need only
generate a velocity equal to the difference between the required
periapsis velocity and Earth's orbital velocity. This velocity is called
V~∞Earth~ and the required value for this excess velocity is 2.940
kilometers per second. The square of V~∞Earth~ is called the C3 value
for the launch and it helps determine what launch vehicle is required to
deliver the spacecraft to the transfer orbit with the proper excess
velocity. This notion is discussed with the launch phase later in this
section of the report. At the end of the transfer orbit, Mars moves
along the direction of the spacecraft's apoapsis velocity. The excess
Mars-centric velocity of the spacecraft, V~∞Mars~ is 2.670 kilometers
per second for the Hohmann transfer. Kinetic energy from this velocity
must be completely dissipated during entry to Mars to avoid a
catastrophic collision on the Martian surface. Energy dissipation is
discussed in the landing phase section of this report.

*Other Phase II Activities*

A number of activities occur during the free flight cruise phase. Star
trackers and sun sensors aboard the spacecraft will determine the
position of the spacecraft during cruise, and the Deep Space Network
will provide tracking and ranging for the spacecraft. The Deep Space
Network is a ground based satellite tracking system that communicates
with spacecraft beyond low Earth orbit. Since it is possible and likely
that the spacecraft will not follow the transfer trajectory exactly as
planned, the onboard propulsion system will execute trajectory
correction maneuvers two or three times during the cruise phase as
needed. A set of three-axis onboard momentum wheels will control the
spacecrafts attitude to counteract solar radiation pressure forces and
keep the communication antennae aimed appropriately during the cruise
phase.

#### Phase I

Determining the launch window for the MSCLE spacecraft depends on design
characteristics of the cruise and Mars entry phases. Since the excess
velocity at Mars must be completely dissipated to facilitate a soft
landing on the surface, it is important to reduce the excess velocity in
order to reduce the amount of fuel required to slow the spacecraft. With
the help of specially designed software, one can examine a variety of
launch windows to determine how to minimize V~∞Mars~ based on the
required C3 at launch. The required C3 for this mission is 8.6, but
performing the analysis within a stated C3 of 10 allows one to choose a
launch window and vehicle that will have enough power (and a little
extra) to deliver the spacecraft to orbit as desired. Using a launch
vehicle with such a C3 value places the launch window during the
intended mission timeframe to the three weeks between October 28 and
November 18, 2011. Taking a transfer orbit that minimizes V~∞Mars~
specifies an arrival to Mars on September 9, 2012. The time of flight
for a mission launching in the middle of the window is about 305 days.
After the first stage of the expendable launch vehicle disconnects, the
second stage will put the spacecraft into a circular low Earth orbit for
a short time. The spacecraft will circle the Earth until it reaches the
proper departure point for V~∞Earth~ to be collinear with the Earth
orbit velocity. A similar circularization will occur during the Mars
encounter phase.

#### Phase III

Upon approach to Mars, the spacecraft will enter an orbit around Mars.
This will allow the spacecraft to perform its final insertion burn at
the appropriate time to send it through the Martian atmosphere to its
designed landing site. The modified Hohmann transfer chosen to minimize
V~∞Mars~ reduces it to approximately 2.152 km/s. A combination of
atmospheric drag, parachutes, and reverse thrusting just prior to
touchdown dissipates the energy of the spacecraft to facilitate a soft
landing in the Martian surface.

### Propulsion

The MSCLE propulsion system was selected by using previous Mars mission
data to determine a launch vehicle and by completing trade studies to
determine the most efficient onboard system. The following sections
detail each system and the steps used in the selection process.

#### Launch Vehicle[^66] 

The launch vehicle will provide the energy required to escape Earth's
gravity and reach Mars. The Delta 7925, produced by Boeing, was chosen
for this mission due to its reputable record and performance
capabilities. The Delta 7925 is a three-stage vehicle with nine strap-on
solid rocket boosters. Spacecraft fairing options are 2.9 m and 3 m in
diameter. In the 2.9-m fairing, the available dimensions are 2.48 m by
4.68 m. In the 3-m fairing, the available dimensions are 2.74 m by 5.03
m. To perform calculations, it was assumed in this report that the
smaller fairing will be used.

A summary of the launch profile is shown in Figure 21. The first stage
is powered by a RS-27A main engine and six solid boosters, referred to
as graphite epoxy motors (GEMs). As the six GEM's reach burnout, the
remaining three boosters are fired. All solid boosters are jettisoned
after they burnout. After Main Engine Cut-Off (MECO) and a period of
coasting, first stage separation occurs. The second stage is a
bipropellant system powering an Aerojet AJ10-118K, and has the
capability to be restarted. Both the first and second stage are gimbaled
for pitch and yaw control. The third stage is spin stabilized using
Thiokol Corporation's Star-48B solid rocket motors (SRMs). A de-spin
before third stage separation is available if it is determined that a
lower spin rate is necessary. The figure below details the mission
profile for a Delta II rocket.

![](media/image30.png){width="4.277083333333334in"
height="3.095833333333333in"}

[]{#__RefHeading___Toc81459562 .anchor}Figure 21: Launch Profile for
Delta II [^67]

#### Onboard Propulsion System

Several iterations were completed in developing the onboard propulsion
system, which is responsible for trajectory corrections, entry into the
Martian atmosphere, and descent to the surface. Methods and equations
used in this process are found in the next section. Originally, the
spacecraft was first placed in an orbit around Mars, rather than
directly entering the atmosphere. A dual mode liquid system was chosen
for this purpose. In a dual mode system, hydrazine (N~2~H~4~) is fed
from a common tank to be used as the monopropellant for catalytic
thrusters and as the fuel for a high-performance bipropellant
engine.[^68] However, idea was rejected because the system required an
extremely large amount of propellant.

The decision was made to convert to a direct insertion profile, similar
to the *2003 Mars Exploration Rovers (MER)*. MER required only a
blow-down hydrazine monopropellant system and a solid rocket system
(Rocket Assisted Descent, or RAD) for terminal descent.[^69] By removing
the bipropellant system, the amount of propellant required was decreased
from 877 kg to 112 kg, and the total spacecraft mass was decreased from
over 1270 kg to 680.5 kg. This will also enable the use of a smaller
launch vehicle.

The monopropellant system is intended to provide final trajectory
corrections after the last stage of the launch vehicle is released (to
prevent the stage from impacting Mars).[^70] The RAD system is fired
about 160 seconds before landing to bring the spacecraft to a halt, at
which point the airbag-enclosed lander is released from the parachute.

#### Methodology 

This section demonstrates the equations used to define the propulsion
subsystem. Values are shown for the first iteration using a dual mode
system for orbit insertion. The launch mass is taken to be the maximum
lift capacity of the Delta II 7925-9.5, which is 1270 kg. The mass of
the adapter, which is the connection between the launch vehicle and the
spacecraft, is scaled using Equation 1. This mass is then subtracted
from the launch mass to obtain the initial mass of the spacecraft in the
cruise stage:

$$m_{\text{adapter}}\  = \ 0\text{.}\text{0755} \cdot m_{\text{launch}}\  + \ \text{50}\  = \ 0\text{.}\text{0755} \cdot \text{1270}\  + \ \text{50}\  = \ \text{145}\text{.}9$$

$m_{\text{initial}\ } = \ m_{\text{launch}}\  - \ m_{\text{adapter}}\  = \ \text{1270}\  - \ \text{145}\text{.}9\  = \ \text{1124}\text{.}1\ \text{kg}\ $.

To solve for the useable propellant mass:

$m_{p}\  = \ m_{i} \cdot \left\lbrack 1 - \text{exp}\left( - \Delta vg_{o} \cdot I_{\text{sp}} \right) \right\rbrack$,

where m~i~ is the initial mass taking into account the propellant mass
lost in previous burns. Table 17 shows a comparison between each
possible system. The specific impulses for monopropellant and
bipropellant systems are estimated to be 220 seconds and 325 seconds,
respectively. Trajectory corrections are also negated in the
calculation. It is clear that a dual mode system will be the most
efficient because it requires the least amount of propellant.

  -------------------------------- ---------------------- ------------- --------- ------------- --------- ------------- --------
  []{#__RefHeading___Toc81459612                                                                                        
  .anchor}*Table 17: Trade Study                                                                                        
  of Propulsion Systems*                                                                                                

                                                          **MONO PROP**           **BI PROP**             **DUAL MODE** 

  *[Burn Type]{.underline}*        *[Delta                *m~p~*        *m~f~*    *m~p~*        *m~f~*    *m~p~*        *m~f~*
                                   V]{.underline}*[^71]                                                                 

  ** **                            *m/s*                  *kg*          *kg*      *kg*          *kg*      *kg*          *kg*

  Orbit Insertion                  2532.0                 776.3         347.8     616.1         508.1     616.1         508.0

  1st Burn (descent)               150.0                  23.4          324.4     23.3          484.7     23.4          484.7

  Retro Rockets                    720.0                  92.0          232.4     98.0          386.7     92.0          392.7

  ***TOTALS***                     ***3402.0***           ***891.7***   *** ***   ***737.4***   *** ***   ***731.4***   * *
  -------------------------------- ---------------------- ------------- --------- ------------- --------- ------------- --------

To account for unusable fuel, reserve requirements, and to ensure a
flexible and reliable system, twenty percent of the useable fuel mass
was added. Thus, the loaded (total) mass of the dual mode system is:

$$m_{p}\  = \ \text{877}\text{.}3\ \text{kg}$$

The on-orbit dry mass (maximum spacecraft payload and bus mass) is:

$$m_{\text{dry}}\  = \ m_{\text{initial}}\  - \ m_{p}\  = \ \text{1124}\text{.}1\  - \ \text{877}\text{.}3\  = \ \text{246}\text{.}4\ \text{kg}$$

The propulsion subsystem inert mass (dry mass of the propulsion system)
is calculated using the propellant mass fraction. This value ranges
between 0.70 and 0.90 for liquid bipropellant systems. A value of 0.90
is assumed for this case.

$$m_{\text{inert}\ } = \ \left( \frac{1}{\lambda} - 1 \right) \cdot m_{p}\  = \ \left( \frac{1}{0\text{.}\text{90}} - 1 \right) \cdot \text{877}\text{.}3\  = \ \text{97}\text{.}5\ \text{kg}$$

A mass margin of 20% (or 29.8 kg) was applied to the dry mass to account
for uncertainties. This value was selected based on the assumption that
the mission is in the preliminary design phase and requirements are
likely to change.

$$m_{\text{allocate}}\  = \ \frac{m_{\text{dry}}\  - \ m_{\text{inert}}}{1 + \ \text{margin}}\  = \ \frac{\text{246}\text{.}4 - \text{97}\text{.}5}{1\text{.}\text{20}}\  = \ \text{124}\text{.}1\ \text{kg}$$

Based on Mars Pathfinder and Polar Lander, the atmospheric entry system
consists of approximately 45% of the dry mass without the propulsion
inert mass.

$$m_{\text{atms}}\  = \ \left( m_{\text{dry}}\  - \ m_{\text{inert}} \right) \cdot 0\text{.}\text{45}\  = \ (\text{246}\text{.}4\  - \ \text{97}\text{.}5) \cdot 0\text{.}\text{45}\  = \ \text{67}\text{.}0\ \text{kg}$$

The mass available for the remaining subsystems is:

$m_{\text{dry}}\  - \ m_{\text{inert}}\  - \ m_{\text{atms}}\  - \ m_{m\text{arg}\text{in}}\  = \ \text{246}\text{.}4\  - \ \text{97}\text{.}5\  - \ \text{67}\text{.}0\  - \text{24}\text{.}8\  = \ \text{57}\text{.}1\ \text{kg}$

This mass clearly cannot satisfy the mission requirements. This is due
largely to the large delta V and consequent large propellant mass
required for orbit insertion.

#### Conclusions

While several options were investigated, the monopropellant hydrazine
and solid rocket systems were selected for their overall efficiency and
simplicity. The next step will be to size the propellant tanks and
determine the overall configuration of the propulsion system.

*Recommendations for further study include:*

-   Smaller launch vehicle: The launch vehicle is currently sized for a
    much larger lander, and should be reevaluated for the significant
    decrease in spacecraft mass.

-   Electric Propulsion (EP) Systems: The use of EP could save mass
    because it has a high I~sp~ and accelerates continuously. The system
    would replace a portion of the thrust requirements placed on the
    launch vehicle, enabling it to lift more mass. This would make
    possible the inclusion of a bipropellant system for Mars orbit
    insertion, which is safer than direct descent. However, such a
    system would necessitate large solar arrays and a significant
    increase in cost.[^72]

### Communications and Data Handling

Communications and data handling are the instruments that converse with
earth or orbiters and the instruments that provide the topic of
conversation. The former is related to the communications antennas and
transmission bands; the latter is related to the electrical systems that
control data flow.

#### Communications Hardware

The proposed communications hardware and process is identical to the
system used on MER. To review and elaborate, the communications systems
from launch to surface operations will be discussed briefly. The process
will be discussed in four parts: tracking techniques, cruise stage, EDL
(Entry, Descent, and Landing) stage, and surface operations.

#### Tracking Techniques

Throughout the life of the mission the craft is tracked and communicated
with using the Deep Space Network (DSN) antennas. The DSN listens for
signals from outer space and uses properties of those signals to compute
position information. This navigation service is called \"tracking
coverage\" and it includes Doppler, ranging, and delta differential
one-way ranging, or \"Delta DOR.\"

*Doppler*

Doppler data is used to construct a plot of the spacecraft velocity
along the line of sight between Earth and its current position.
Electromagnetic radiation experiences a frequency shift relative to the
observer as the craft moves. The size of the frequency shift, or
\"Doppler shift,\" depends on how fast the light source is moving
relative to the observer. Light from an object coming towards us is
shifted to the blue end of the spectrum (higher frequencies), and light
from an object moving away is shifted towards the red (lower
frequencies).

Ground controllers know the frequency of the signal that is emitted from
the spacecraft. By comparing the received frequency with the emitted
frequency, the magnitude of the Doppler shift can be calculated. Based
on that information, velocity is straightforward to calculate. The
longitudinal Doppler Effect considers the simpler case of a source
moving directly towards you or away from you along a straight line:

> $$f = f^{'}\sqrt{\frac{1 + \frac{v}{c}}{1 - \frac{v}{c}}}$$

$c\left( \frac{\delta^{2} - 1}{1 + \delta^{2}} \right) = v$ , where
$\delta = \left( \frac{f}{f'} \right)^{2}$.

*Ranging*

Ranging is accomplished by pinging the spacecraft (sending a code to the
spacecraft, having the spacecraft receive that code and immediately send
it back to Earth). The time between sending the code and receiving the
code, minus the delay in turning the signal around on the spacecraft, is
twice the light time to the spacecraft. This distance is accurate to
about five to ten meters:

> $d = c\left( \frac{t - t_{\text{delay}}}{2} \right)$.

*Delta DOR*

Delta DOR is similar to ranging, but uses a third signal from a
naturally occurring radio source in space to reduce location measurement
error. Quasars are used as extremely well known positions in the sky to
provide system calibration. Being able to do quasar and spacecraft
ranging near the same time and subtracting the answers cancels a
majority of the common errors in atmospheric environment and equipment.

*Cruise*

Communication during the cruise state relies on a cruise low-gain
antenna (CLGA) and a cruise medium-gain antenna (CMGA) both pictured in
Figure 9: MER Descent Configuration. The cruise low-gain antenna is
mounted inside the inner ring and the cruise medium-gain antenna is
mounted in the outer ring. Periodic orientation updates will make sure
the antenna stays pointed toward Earth and that the solar panels stay
pointed toward the sun. During the early cruise phase, a CLGA provides
the communications link with Earth. A low-gain antenna does not need to
be pointed as precisely as a higher-gain antenna, making it ideal for
early communications. Later in the cruise, the angle between the Sun and
Earth shrink, or the sun enters a larger portion of the area between
Earth and the craft, reducing the amount of the energy that reaches
Earth from the CLGA. The spacecraft switches to a CMGA, which can direct
the same amount of energy into a tighter beam. This reduces the amount
of area taken up by the sun in the transmission line between Earth and
the spacecraft, enabling communications.

The CMGA has a waveguide (a pipe) connection to the rover's antennas.
The LGA on the cruise stage connects directly to the LGA on the
backshell (Figure 9: MER Descent Configuration), which in turn connects
to the LGA on the lander (Figure 23: Lander Antenna Array). Power of
signal transmissions is not an issue for the cruise stage as the system
is already in existence and has been tested during MER.

![](media/image31.jpeg){width="5.554861111111111in"
height="1.9791666666666667in"}

[]{#__RefHeading___Toc81459563 .anchor}Figure 22: EDL Components[^73]

*Entry, Descent and Landing (EDL)*

From cruise stage separation until the lander (Figure 9: MER Descent
Configuration) is separated from the backshell (seen in Figure 22: EDL
Components), communications are passed directly to Earth across an
X-band (8.4 GHz) link. Communications are conducted as a series of 128
basic, distinct radio tones that will ring during different phases of
the entry, descent, and landing process. The backshell low gain antenna
(BLGA) provides the necessary capability. After landing the BLGA can no
longer be used. Until the rover lands, two addition communications
systems are used: a DTE (Direct to Earth) link over the rover low-gain
antenna (RLGA) and an UHF relay link. The UHF link transmits data to any
of the Mars orbiters that are present or will be present at time of
arrival to the planet. Data is relayed from the orbiters top Earth using
a standard phase-coherent X-band link. UHF must be provided along side
the DTE X-band signal because the DTE link is unreliable during EDL. UHF
will be used as the primary data link until landing, but will not be
used after that time. The X-band signals throughout this period are
tracked by one of three 70-meter Deep Space Network (DSN) antennas.

One of the Mars orbiters will need to be adjusted in preceding weeks to
place it over the landing vicinity during those crucial minutes to
receive the transmissions. The orbiter will later transmit the data to
Earth. This first relay link is designed to provide information for
later reconstruction of the events during the spacecraft\'s descent and
landing, not to provide immediate play-by-play notice as they happen.

![](media/image32.jpeg){width="4.124305555555556in"
height="2.082638888888889in"}

[]{#__RefHeading___Toc81459564 .anchor}Figure 23: Lander Antenna
Array^Error:\ Reference\ source\ not\ found^

*Surface Operations*

Communications through the duration of the mission will be conducted
with three different X-band systems: a low-gain antenna (LGA), an UHF
antenna, and a high-gain antenna (HGA). DSN antennas communicate with
the lander directly through a HGA link or using the orbiters as a go
between (antennas can be seen in Figure 23: Lander Antenna Array). DTE
communications suffer from power loss in transit from Mars to Earth,
making the signal power so weak that even DSN communications become a
difficult task. The benefits of using the orbiting spacecraft are that
the orbiters are much closer to the rovers than the DSN antennas and
that the orbiters can communicate with Earth longer than a body rotating
on the Martian surface.

When the rovers communicate directly to Earth, they send messages via
both the low-gain antenna (LGA) and the high-gain antenna (HGA). Both
antennas are capable of receiving the 7.2 GHz uplink signal and
transmitting the 8.4 GHz downlink signal. The LGA is a choked circular
waveguide design having about a 70° beam width pattern with about 6 dBic
boresite gain at 7.2 and 8.4 GHz. Although showing some directivity, the
LGA is considered a quasi omni directional antenna. The omni directional
UHF antenna communicates through orbiters passing overhead.

The HGA is mounted on a gimble motor (HGAG). The dish is able to orient
itself towards Earth, providing the lander with a DTE link with the DSN
without moving the lander. About half of all communications will go
through the HGA, a dish 28 cm in diameter that beams data directly to
DSN over the X-band (8 to 12 GHz). The downlink data rate to the orbiter
is established with the HGA at 264 bytes/sec (1.85 kilobits/sec) while
the uplink data rate is 125 bytes/sec (0.875 kilobits/sec) utilizing the
LGA. The power required for transmissions is approximately 10 W for the
LGA and 15 W for the HGA. The power consumption is reasonable, never
placing the combined subsystem over a 25 W consumption. Further
information can be found on the Mars Exploration Rover description
hosted by Jet Propulsion LaboratoryError: Reference source not found.

*Data Handling Electrical System*

The communications backbone is controlled by the electronics subsystem
of the lander. Traditionally, the RAD6000 microprocessor and interface
board (U6 Compact PCI) handled the task. RAD6000 microprocessors are
radiation-hardened versions of the PowerPC chips that powered Macintosh
computers in the early 1990s, with 128 megabytes of random access memory
(RAM) and capable of carrying out about 20 million instructions per
second (MIPS). The Mars Exploration Rovers, Mars Odyssey, and Mars
Pathfinder relied on the RAD6000 to provide all computer functionality,
including data handling.

BAE Systems, manufacturer of the RAD6000 placed into production the
RAD750 processor during 2001. BAE Systems have designed the RAD750 as a
replacement to the RAD6000, providing processing capabilities out
reaching other space hardware currently on the market. The company has
taken special care to develop code migration software to transform
RAD6000 software into RAD750 software. Because the RAD6000 will be
replaced by the RAD750 by the proposed flight date, the RAD750 will be
used to provide all data handling, payload interface, and equipment
control. BAE Systems provides the RAD750 on a single board, RAD750 3U
Compact PCI, with standard interfaces and memory.

![](media/image33.png){width="3.1243055555555554in"
height="2.2847222222222223in"}

[]{#__RefHeading___Toc81459565 .anchor}Figure 24: RAD750 3U Compact
PCI[^74]

**RAD750 3U Compact PCI**

The RAD750 is functionally identical to the G3 processors used by Apple
Computers. The RAD750 is expected to run at 166 MHz, providing a
throughput of 300 MIPS. Power consumption can be reduced by attaching
slower oscillators to the 3U interface board (Figure 24: RAD750 3U
Compact PCI). Due to power requirements, the project plans to outfit the
board with a 33 MHz oscillator to reduce power consumption from 11.5 W
to 7.5 W at 3.3 V. Computational power will be reduced; however, the
system will still be computationally superior to the RAD6000 systems of
previous spacecraft. A 10-20% performance increase is possible via the
addition of an external L2 cache chip. The board already has facilities
setup to allow inclusion of the L2 cache. If the system becomes
computationally crippled by reducing the oscillator speed from 133 MHz
to 33 MHz, the L2 cache will be included to help mitigate the effect.

+----------------------------+-----------------------------------------+
| []{#                       |                                         |
| __RefHeading___Toc81459613 |                                         |
| .anchor}*Table 18: RAD750  |                                         |
| 3U Compact PCI             |                                         |
| Summary*[^75]              |                                         |
+----------------------------+-----------------------------------------+
| *[Board                    |                                         |
| Sp                         |                                         |
| ecifications]{.underline}* |                                         |
+----------------------------+-----------------------------------------+
| **Form Factor**            | Dimensions: 100mm X 160 mm              |
|                            |                                         |
|                            | Weight: 549 grams                       |
+----------------------------+-----------------------------------------+
| **Processor**              | RAD750 Dhrystone 2.1 MIPS @ 133 MHz     |
+----------------------------+-----------------------------------------+
| **Radiation Hardness**     | Total Dose \> 100 Krad (SI)             |
+----------------------------+-----------------------------------------+
| **PCI Backplane Bus**      | 32 bit, 33 MHz PCI version 2.2          |
+----------------------------+-----------------------------------------+
| **Power Supply**           | 3.3 V +/- 10%                           |
+----------------------------+-----------------------------------------+
| **Power Dissipation**      | less than 10.2 W                        |
+----------------------------+-----------------------------------------+
| **Rail Temperature Range** | -55 to 80 C                             |
+----------------------------+-----------------------------------------+
| **Mean Time between        | greater than 390 K hours                |
| Failures**                 |                                         |
+----------------------------+-----------------------------------------+

The Power PCI is manufactured on a 0.5 micron radiation hardened process
line with 0.5 micron Leff and up to 5 levels of metal with planarized
passivation. The technology will achieve Total Ionizing Dose hardness
levels of 1 Mrad(Si). This single board computer consists of 128 MB of
synchronous DRAM packaged in 4 high stacks and 256 KB of EEPROM, both
with error correction. An on-board regulator generates the 2.5 volts
required for the RAD750 processor core, so the board only requires a 3.3
volt supply. The compact size of the 3U board () provides greater
flexibility in spacecraft size and shape, and is well suited to modular
design of systems based on standard products. The issues related to the
RAD6000 and temporary loss of communications with the MER Rovers has
been fixed with the 750 hardware revision. The problem was related to a
filling of the additional external memory used on MER. Once the memory
was filled, the computer continually rebooted attempting to clear the
state of the memory. The solution found was to place the rover in
cripple mode. If the computer is placed in cripple mode, the interface
to external memory and devices are disabled, allowing the computer to
boot up normally. This mission should never enter into the same
situation because there is no plan for an additional memory expansion.
However, the computer will sent in cripple mode to avoid any other
problems that were simultaneously solved by placing the machine in
cripple mode to clear the external memory. A summary of the RAD750 3U
Compact PCI can be found in . More detailed information can be found in
the hardware specification for the RAD750[^76].

### Ground Support and Operations

The Ground Operations Subsystem has made design decisions for a
communications system between the earth and the MSCLE spacecraft. The
design explored the use of the Deep Space Network (DSN), based out of
the Jet Propulsion Lab, as the telecommunications system. After
examining the services offered by the DSN, those deemed necessary for
this mission were selected. The need for human involvement in spacecraft
operations was also considered, and a framework was developed for ground
crew scheduling.

#### Overview of Deep Space Network

*Facilities*

The Deep Space Network is a set of three ground stations positioned
approximately 120 degrees apart in longitude. They are located in
Madrid, Spain; Canberra, Australia; and Goldstone, California. The
geography is designed such that communication is never disrupted by the
earth's rotation. There are three types of antennas at each DSN station.
The three used for deep space missions are the 34-meter High Efficiency
(HEF), the 34-meter Beam Waveguide (BWG), and the 70-meter antenna. The
two types of 34-meter antennas have similar performance, and are both
far surpassed by the 70-meter antenna (see Figure 25[^77]), which can
track a spacecraft up to 16 billion km away[^78].

![](media/image34.jpeg){width="2.2958333333333334in"
height="2.3194444444444446in"}

[]{#__RefHeading___Toc81459566 .anchor}Figure 25: DSN 70-meter antenna

*Interaction with Customers*

The Deep Space Network offers the Deep Space Mission System Services
Catalog[^79], which provides detailed explanation of the options
available to a potential user. The services are divided among the
following categories.

i.  Command Services

ii. Telemetry Services

iii. Mission Data Management

iv. Tracking and Navigation

v.  Experiment Data Product Services

vi. Critical Event Communication

The services catalog is designed to be used at the proposal stage. The
catalog is organized according to what the DSN can provide, not
according to what equipment or manpower will be used. A partner
document, called Mission Operations and Communications Services
(MOCS)[^80], provides fee estimates, detail about the actual equipment
within the DSN, and the minimum mission parameters needed for a proposal
requesting the use of the DSN. Both documents were used in developing
the following communication system design. The services discussed below
are based on the needs of a Discovery Class mission, such as Pathfinder.
As the cost cap for Discovery class missions is \$350 million, the price
and capability range for ground operations based on these documents is a
reasonable starting point for our \$450 million Scout Mission[^81].

#### Deeps Space Network Services

*Command*

There are two options for sending commands from the earth to the
spacecraft. The first is Command Radiation; it guarantees timely
delivery of command files, but not error free delivery. Command
Delivery, which does guarantee error-free delivery, was chosen.

*Telemetry*

This service has several options about the format of the data that is
sent from the spacecraft to earth regarding its state of health,
position and the state of the instruments. For this design, information
will be received in the form of standard packets and delivered the
project Mission Operations System. The packets will be ordered by the
time received on earth. Once the data packets are received the DSN also
provides a service to assemble the data into files and store them in a
database. Any delivery failures are reported to the users and to the
spacecraft, allowing for the resending of the data.

*Mission* *Data* *Management*

This service provides the buffering, staging and safe-keeping of mission
data and ensures safe delivery to the scientist. The data can be stored
for up to 30 days.

*Tracking* *and Navigation*

The tracking needs of the transit portion of the mission are being met
with a Validated Radio Metric Data Service. This is recommended over raw
radio metric data, allowing for mistakes in Doppler, ranging or angle
data. This service allows for validation of the above data.

*Experiment Data Product*

In order to facilitate processing of the scientific data, the Level 1
Processing service will be applied. This allows for preliminary
calibration of data and data quality check.

*Critical* *Event Communication*

During descent into the Martian atmosphere, communication is difficult
due to the extreme environment. As was done for the Mars Exploration
Rovers[^82], communication during the so-called "six minutes of terror"
will be through a beacon tone service. This service provides minimal
information about the state of health of the spacecraft by sending a
four tone Beacon Monitoring signal to any of the DSN antennas. This is a
simple way to confirm spacecraft survival.

#### Station Support Requirements

The section on the Flight Dynamics subsystem describes the sequence of
mission phases before arriving on Mars. Table 19 below summarizes this
sequence and specifies which antennas will be used during the various
phases. Columns three and four show the frequency with which the
spacecraft will be contacted during transit. The most frequent
communication is during the launch and planetary insertion phases
because of their critical nature. Not shown are the various trajectory
correction maneuvers (TCM's) that would be planned in real time
depending on the progress of the spacecraft in order to keep it on the
optimal path.

  -------------------------------- ------------------------- --------------------------- ---------------------
  []{#__RefHeading___Toc81459614                                                         
  .anchor}*Table 19: Station                                                             
  Support Parameters for Various                                                         
  Mission Phases*                                                                        

  ***[Mission                      *[Antenna]{.underline}*   *[\# of                     *[\#
  Phase]{.underline}***                                      Tracks/week]{.underline}*   Weeks]{.underline}*

  **Launch**                       34 m BWG                  21                          3

  **Cruise**                       34 m BWG                  2                           4

  **Planetary Insertion**          34 m HEF                  21                          6

  **Surface Operations**           70 m                      14                          2+
  -------------------------------- ------------------------- --------------------------- ---------------------

#### Overview of Ground Crew Schedule

The primary surface operations for this mission are drilling into the
Martian surface and testing the soil and rock samples. As discussed in
the section on the Payload Subsystem, the following drilling schedule
will be implanted (see Error: Reference source not found).

  -------------------------------- --------------------------------------------------------
  []{#__RefHeading___Toc81459615   
  .anchor}*Table 20: Schedule for  
  Payload/Science Operations on    
  Martian Surface*                 

  ***[Day]{.underline}***          *[Activity]{.underline}*

  **1**                            Test all motors

  **2**                            Camera Pictures

  **3**                            Meteorological Measurements

  **4**                            Drill down 10 cm

  **5**                            Test sample

  **6**                            Bring drill up

  **7**                            Test sample

  **8**                            Drill down .5 meter

  **9**                            Drill to 1 meter

  **10**                           Bring drill up

  **11**                           Drill down to 1 m

  **12**                           Drill down to 2 m

  **13**                           Bring drill up

  **14**                           Test sample
  -------------------------------- --------------------------------------------------------

Constraints on the drill schedule are based on power availability and
drill speed. There are 6 hours/sol when enough power will be available
to drill and estimated drill speed is 10 cm/hour. As a result, the
ground crew on earth working approximately eight hour shifts will be
sufficient for surface operations. Those shifts will be scheduled
according to power availability and the position of Mars relative to the
earth. Drilling and science will start when the spacecraft is facing the
earth and is able to use its high gain antenna to communicate. Enough
power is available for a brief interaction with the ground control
through the high gain antenna before surface operations begin on any
given day. The earth crew would check that the environment and equipment
were safe for surface operations and give the command to proceed. The
orbiter passes overhead for eight minutes each day. Whenever it passes,
all data would be sent to the orbiter. The rest of the data would be
sent during the low power session for three hours at the end of each
sol.

#### Other Issues

*Signal Time Delay*

The time delay to send a signal between Mars and the earth one way can
vary from a minimum of 5 to 20 minutes[^83]. Design was made assuming a
20 minute delay.

*Sharing* the DSN

It is well understood by the designer that the DSN is the primary ground
system for all Deep Space missions, particularly those that go beyond
two million kilometers from the earth. The use of the DSN would thus
require very specific scheduling and efforts would be made to minimize
DSN usage through methods such as the Multiple Spacecraft Per Aperture
(MSPA) capability, which allows up to two spacecraft to communicate
through the network at one time[^84].

## End of Life Analysis

After achieving the scientific goal of having the lander drill to a
depth of two meters, the MSCLE team hopes to obtain more science. If the
various subsystems are still adequately working, the lander will drill
down to eight meters and examine the Martian samples at various depths.
Once the maximum depth is reached, the lander may be used as a testbed
for other drilling applications, such as software. In addition, the
meteorological equipment will continue to monitor the atmosphere
conditions as long as possible. Members of the engineering team have
contributed other ideas that could take place aboard the lander; however
feasible or infeasible, these ideas are listed below:

i.  Analyze the samples over time and determine if the samples are
    affected by season

ii. Catch some dust in the sample bay and analyze the particles

iii. Friction heat the environment by having the drill spin in place and
     analyze the results

iv. Determine if the extended drill can behave as an antenna and analyze
    any frequencies picked up under the surface

After all systems have died, the lander will be stranded on Mars;
retrieving the lander in any fashion is economically infeasible. When
the lander cannot no longer be used, it will remain as a landmark to
science and exploration.

# Test and Validation

## Objective

The objective of this subsystem is to make sure that the craft that is
sent up into space is of high caliber quality. A part of verifying that
the platform will work once in the Martian and vacuum space environment
is to make sure that the configuration of the system lies within the
standards set forth in the MIL Q 9858A, which describes the quality
assurance for military programs. Testing and validation also encompasses
making sure that the equipment and the instruments being supplied from
the various vendors are compatible with our programs and sufficiently
tested within the constraint limits set forth for the mission.

## Approach

Tests will be organized in the following order:

i.  Functional test

ii. vibration

iii. functional test (for verification and redundancy)

iv. shock

v.  Functional test

vi. Thermal vacuum functional test during exposure

vii. Flash X-ray Functional test

As a part of qualifying the craft each spacecraft function and all of
the environmental effects the spacecraft may encounter must be simulated
and tested:

*Step 1*: Identify a spacecraft and payload functions

i.  The payload has been identified as being a drilling device, and the
    scientific instruments; mass spectrometer, neutron spectrometer,
    gamma ray spectrometer, x-ray spectrometer and chemical and
    biological analysis packages. The platform in order to accomplish
    the mission objectives consists of a lander with an auger adorned
    drill.

ii. Testing of the spacecraft and payload function for assurance on
    proper operation need to be performed.

*Step 2*: Identify the environment

i.  Identify environment for transportation, storage, launch, and orbit
    vibration, shock, temperature, vacuum and radiation amounts

ii. These values all need to be calculated ^3^

Launch conditions will be tested for the vibrations and the heat
incursions for the Delta 7925 rocket. En route conditions will be tested
for temperature variation between -123 and 120 degrees Celsius.[^85]
Figure 26 portrays the radiation that the craft will encounter while in
transient to Mars from Earth.

![](media/image35.png){width="5.247222222222222in"
height="3.171527777777778in"}

[]{#__RefHeading___Toc81459567 .anchor}Figure 26: These are the
radiation problems that will be encountered^2^

i.  Surface pressure: 6.36 mb at mean radius (variable from 4.0 to 8.7
    mb depending on season) \[6.9 mb to 9 mb (Viking 1 Lander site)\]

ii. Surface density: \~0.020 kg/m3

iii. Scale height: 11.1 km

iv. Total mass of atmosphere: \~2.5 x 10^16^ kg

v.  Average temperature: \~210 K (-63 C)

vi. Diurnal temperature range: 184 K to 242 K (-89 to -31 C) (Viking 1
    Lander site)

vii. Wind speeds: 2-7 m/s (summer), 5-10 m/s (fall), 17-30 m/s (dust
     storm) (Viking Lander sites)

viii. Mean molecular weight: 43.34 g/mole

ix. Atmospheric composition (by volume):

    a.  Major: Carbon Dioxide (CO~2~) - 95.32% ; Nitrogen (N~2~) - 2.7%

    b.  Argon (Ar) - 1.6%; Oxygen (O~2~) - 0.13%; Carbon Monoxide (CO) -
        0.08% Minor (ppm): Water (H~2~O) - 210; Nitrogen Oxide (NO) -
        100; Neon (Ne) - 2.5;

    c.  Hydrogen-Deuterium-Oxygen (HDO) - 0.85; Krypton (Kr) - 0.3;

    d.  Xenon (Xe) - 0.08^4^

*Step 3*: Correlate Functions and environments

i.  Test all of the operating equipment, that will be powered during
    launch, during the spacecraft vibration and check all modes of en
    route operation.

ii. The parameters for the instrumentation limits are unknown at the
    current time.

*Step 4*: Identify Main configurations:

i.  This includes boost configuration and orbital configurations and
    flight paths

ii. The lander will be launched from the top of a Delta 7925 rocket

iii. The Spacecraft will launch between October to November 7th 2011,
     circle around the Earth and depart at point required for a proper
     heading to Mars, cruise en route is approximately 305 days in a
     Hohmann transfer, the craft will then circle around Mars upon
     encounter and descend at a point required to hit within the landing
     ellipse, arrival will be September 9th 2012.

*Step 5*: Devise Functional Tests for each Major configuration:

-   Test all software and equipment in configuration to assure
    compatibility

*Step 6*: Layout a sequence of functional tests and environmental
testing:

-   This has already been done, see above for test organization.

*Step* 7: Identify Span Times and Special Facility requirements

-   The life span of this mission has been set for 90 days + en-route
    time to Mars

## Tests

There are a variety of tests that need to be issued upon the structure
before the craft can be deemed flight ready. Structural tests include;
static loading tests, and modal surveys. Deployment tests involve; solar
array testing, antenna testing, and appendages testing. Vibration
testing consists of acoustic testing and vibration by low frequency sine
waves, and the use of shaker tables. Pyro-shock tests are a firing of
the ordinance in order to test the mechanical deployments on the
platform. Payload testing includes testing of all of the instruments on
board. While on launch the lander will experience vibrations and heat of
the magnitude corresponding to that of a Delta 7925 rocket. Tests will
need to be administered in order to verify that the instruments that are
to be placed on the lander will be able to withstand the launch
temperature variations and vibrations. Tests will also need to be done
on the configuration in order to determine how the craft will behave in
the en route vacuum of space as well as test for the conditions on the
surface of Mars.

Testing the Mars Scout mission platform will involve a lot of
integration of subsystems as well as management of configurations of
equipment. This is possibly the most important portion of a mission due
to the fact that if testing is not accomplished than the final product
may not be worth flying. It is essential. ^3^

# Budget

## Methods

In general, all methods of cost estimation for space missions can be
grouped into three categories---detailed bottom-up estimation,
analogy-based estimation, and parametric estimation.[^86]

**Bottom-up estimation** is the process of calculating the cost of a
mission by adding each of its individual components. It would begin with
a detailed list of all scientific instruments and spacecraft hardware,
development timelines and personnel costs, ground facility requirements,
and other costs. Adding the costs of each component of a mission will
give the cost of the total mission. This method is very detailed and
time-intensive.

**Analogy-based estimation** uses knowledge of the costs of previous
missions to estimate the cost of a future mission. Previous missions
have been flown with various combinations of scientific instruments and
bus hardware, and these data points can be combined to estimate the cost
of the MSCLE mission.

**Parametric estimation** involves large formulae that mathematically
account for various parameters such as mass, development time, and
mission difficulty. These formulae are called Cost Estimating
Relationships, or CERs. The formulae are already developed, and to
estimate the cost of the MSCLE mission, the mission specific parameters
can just be plugged in.

  ----------------------------------- ------------------------ --------------------------
  []{#__RefHeading___Toc81459616                               
  .anchor}*Table 21: Pros and Cons of                          
  each cost estimation method*                                 

                                      *[Pros]{.underline}*     *[Cons]{.underline}*

  ***[Bottom-up]{.underline}***       Estimate is specific to  Very time-consuming, and
                                      each mission, and is     not accurate at the early
                                      generally accurate once  stages of a project
                                      all components are known 

  ***[Analogy-based]{.underline}***   Many missions similar to Relies completely on past
                                      MSCLE                    information, does not take
                                                               into account newer
                                                               practices

  ***[Parametric]{.underline}***      Simple, top-down         Inaccurate
                                      approach                 
  ----------------------------------- ------------------------ --------------------------

To do the cost estimation for the MSCLE mission, the team chose to use
both the analogy method and the parametric method. The bottom-up method
was not feasible for two reasons -- there was not sufficient time, and
accurate information about the content of the spacecraft was not
available. The objectives of the mission were changing rapidly, and
trying to do useful bottom-up cost estimation would not have been
fruitful. The other two methods are both fast and are useful in the
early stages of a mission.

The benefits of using two methods instead of one are obvious. The
parametric method is easy but inaccurate. It is useful for generating
rough estimates in short order. When the specifications of the mission
stop changing so fast, the analogy-based method generates a different
cost estimate.

## Cost Estimating Procedure

The cost estimating team performed three iterations of cost analysis
during the course of the mission design.

### First Iteration

During the first iteration, an analogy-based method was used based on
six previous unmanned Mars missions:

i.  *Mars Pathfinder* addressed key engineering issues involved with
    landing and operating a rover on Mars at low cost. Its scientific
    instrumentation included an imager, meteorology equipment, and an
    alpha proton x-ray spectrometer.

ii. *Mars Odyssey* was a medium-cost orbiter mission launched with a
    substantial scientific package that included a thermal emission
    imaging system, a gamma ray spectrometer, and a Martian radiation
    environment experiment.

iii. *Mars Express* was developed by the European Space Agency to
     perform valuable Mars science at low production cost. Its science
     instrumentation included a high-resolution stereo camera, an
     energetic neutral atoms analyzer, a Fourier spectrometer, visible
     and infrared mineralogical mapping spectrometers, a subsurface
     sounding radar altimeter, a radio science experiment, an
     ultraviolet and infrared atmospheric spectrometer, and the Beagle 2
     probe.

iv. *Mars Polar Lander* was an unsuccessful mission developed to study
    the planetary features of the polar regions of the Martian surface.
    Its instrumentation included a microphone, meteorology equipment, a
    thermal gas analyzer, a mass spectrometer, a stereo camera, and a
    light detection and ranging experiment.

v.  *Mars Global Surveyor* was developed at very low-cost with a modest
    scientific package that included a camera, altimeter, magnetometer,
    and celestial sensor.

vi. *Viking*, by far the most costly unmanned mission, was developed at
    a cost equivalent to approximately three billion 1997 dollars. The
    Viking mission included multiple craft units, and its scientific
    payload consisted of two seismometers, a gas chromatograph, x-ray
    spectrometer, mass spectrometer, meteorology equipment, cameras,
    thermal imagers, and orbital water detectors. The multiple
    spacecraft units, as well as the extensive mission timeline and high
    reliability of spacecraft operation contributed to this large cost.

The proposed Academy scout mission was compared to these examples, and
the cost team concluded that the scientific and engineering requirements
of MSCLE fell far short of the demanding Viking mission. At the same
time, MSCLE's challenging landing requirements and drilling method
exceeded the requirements of the Global Surveyor, Pathfinder, and
Odyssey missions. The platform was very similar to the Polar Lander, but
the sample preparation required added additional engineering complexity.
For these reasons, the cost of the Academy mission was estimated to be
approximately \$500 million.

  -------------------------------- -------------------------- ---------------------- --------------------------------
  []{#__RefHeading___Toc81459617                                                     
  .anchor}*Table 22: Mars Mission                                                    
  Analogies*                                                                         

  *[Mission]{.underline}*          *[Cost                     *[Type]{.underline}*   *[Payload]{.underline}*
                                   (millions)]{.underline}*                          

  Viking                           \$3,000                    Series                 Seismometer, gas chromatograph,
                                                                                     x-ray spectrometer, mass
                                                                                     spectrometer, meteorology,
                                                                                     cameras, thermal imagers,
                                                                                     orbital water detectors

  Polar Lander                     \$356                      Lander                 Microphone, meteorology, thermal
                                                                                     gas analyzer, mass spectrometer,
                                                                                     stereo camera, light detection
                                                                                     and ranging

  Express                          \$325                      Orbiter and Lander     Stereo camera, neutral atoms
                                                                                     analyzer, Fourier spectrometer,
                                                                                     visual and infrared mapping
                                                                                     spectrometers

  Odyssey                          \$297                      Orbiter                Thermal emission imaging system,
                                                                                     gamma ray spectrometer, Martian
                                                                                     radiation environment experiment

  Pathfinder                       \$250                      Orbiter and Rover      Imager, meteorology,
                                                                                     alpha-proton x-ray spectrometer

  Global Surveyor                  \$157                      Orbiter                Camera, altimeter, magnetometer,
                                                                                     celestial sensor
  -------------------------------- -------------------------- ---------------------- --------------------------------

### Second Iteration

In the second iteration, a preliminary mass estimate was used as a
parameter in two cost-estimating relationships:

The *Spacecraft/Vehicle Level Cost Model (SVLCM)* is a rough order of
magnitude cost-estimating relationship available online that uses
parameters of mission type, dry mass, and spacecraft quantity to produce
a total cost estimate.[^87] The chief advantage of the model is its
simplicity and availability, but it does not reflect the technological
sophistication of the mission or the interaction of the mission
subsystems. This model was compared to other, more robust cost estimates
to confirm that these additional estimates were reasonable.

The *Advanced Missions Cost Model (AMCM)* is a more thorough order of
magnitude cost-estimating relationship, also available online, that uses
more detailed mission type parameters, dry mass, spacecraft quantity,
Initial Operating Capability (IOC) year, block number (the number of
similar design iterations performed earlier), and difficulty (on a Very
Low to Very High scale) to produce a total cost estimate.[^88] The model
is also simple to use and readily available, but unlike SVLCM, the block
number and difficulty parameters adjust the cost to reflect the
complexity and technological readiness of the mission design.
Additionally, the IOC year parameter reflects the increase in mission
cost as the time scale of development increases. For these reasons, AMCM
is favorable to SVLCM. Table 23 summarizes the parameters in each cost
model.

During the second iteration, mass estimates for each subsystem were not
yet available, so the cost team assumed a dry mass equal to the
comparable Mars Polar Lander mission: 1200 lbs. Estimating each of the
other parameters in the cost-estimating relationships, and comparing
these estimates to the analogy-based method developed under the first
iteration, the budget team concluded that the mission design still
exceeded budget requirements (Table 244: Second Iteration Cost
Estimates).

  -------------------------------- ----------------- ----------------- ----------------
  []{#__RefHeading___Toc81459619                                       
  .anchor}*Table 244: Second                                           
  Iteration Cost Estimates*                                            

  **Model**                        Analogy           SVLCM             AMCM

  **Estimate**                     \$500m            \$516m            \$773m
  -------------------------------- ----------------- ----------------- ----------------

### Third Iteration

In the third cost-estimating iteration, each subsystem of the
engineering design team provided reliable mass estimates. Since the dry
mass parameter of the AMCM cost model specifically referred to the
unmanned planetary lander, the engineering team identified the system
dry mass as the total mass of the lander when it will first contact the
planetary surface. This assessment of the system mass resulted in a mass
estimate much lower than the Mars Polar Lander estimate (1200 lbs.).
Since this dry mass was a key parameter driving the cost in both the
AMCM and SVLCM models, cost estimates in the third iteration were
considerably lower than in the second (Table 25).

  -------------------------------- ----------------------- ----------------------
  []{#__RefHeading___Toc81459620                           
  .anchor}*Table 255: Third                                
  Iteration Cost Estimates*                                

  **Model**                        SVLCM                   AMCM

  **Estimate**                     \$272m                  \$369m
  -------------------------------- ----------------------- ----------------------

Although SVLCM suggested that the mission cost was so low that
additional subsystems should be added, the AMCM estimate was favored
again for its ability to model the technical complexity of mission
designs. With a final predicted cost of \$369 million, a 20% margin for
cost growth produces a maximum mission budget of \$442.8 million.

There are several ways this cost-estimation method could be improved.
The parametric cost models used above are designed for quick, pre-phase
A cost estimation. In a Phase A project design, a more accurate
parametric cost model should be used. A variety of models are
commercially available, but one of the most common spacecraft
cost-estimating models is the NASA / Air Force Cost Model (NAFCOM).
NAFCOM considers a much larger set of overall system and subsystem
parameters, producing cost estimates with much higher fidelity than AMCM
or SVLCM.[^89] The next step in the MSCLE cost estimation process is to
acquire this model and use it to identify a more accurate cost estimate.

# Outreach

"....To explore the universe, and search for life. To inspire the next
generation of explorers, as only NASA can."

It is vital that the present generation inspire and motivate a new
vanguard, the next wave of explorers. Future space endeavors require a
team effort from individuals with great expertise and well-roundedness;
it does not only require those whom have been inspired toward math and
the sciences, but those whose hearts question and wonder. This is the
propelling vision behind the Mars Subsurface Chemical Life Explorer
(MSCLE) mission.

The purpose of the outreach team is to make the mission accessible to
the public to engender curiosity and excitement about the search for
life. The goals of the outreach team are to (1) educate the public on
the mission's goals, science, and purpose, (2) demonstrate to the public
the benefits of scientific and technological programs like MSCLE, and
(3) to inspire and excite the public. The message that the outreach team
wants to convey is that searching for life on Mars is both an exciting
endeavor and an investment in the future. The investment is in the
expected future benefits to the public that stem directly from the
scientific and technological advances needed to make the mission
succeed.

## Educate the Public

One approach of the education and outreach team is the development of a
website explaining many aspects of the group project. The audience for
the majority of the website is the public including students in high
school. There is one page in the website designed to target younger
students in late elementary school through middle school.

A number of the pages in the website give information specific to the
MSCLE mission. The homepage provides a brief description of the project
and the motivation behind the website (). An "About Us" page summarizes
the academy and lists the structure of the group project teams. Another
page provides a description of the responsibilities of each group
project team. On the science page is a description of the scientific
instruments used on the mission, while the lander page describes the
platform that houses the instruments and how the drill that brings the
samples to the instruments.

The other pages in the website contain information about Mars in
general. One of these is a "Fast Facts" page that summarizes basic facts
about Mars and includes some details on Valles Marineris since it is
directly relevant to the group project. The "Future Investments" page,
which echoes the themes of the video clip (described in section 3),
lists some past missions with their spin-offs and describes the benefits
of the journey. It explains how searching for life on Mars will probably
have long-term benefits that are not immediately recognized or
understood.

The "For Kids" page includes general information about searching for
life on Mars, including some questions to help get kids thinking about
science. There is a page showing pictures from both Mars and Earth that
asks students to identify which planet each picture is from (). Included
is a brief explanation of the geological features seen in the pictures
and how they relate to the search for life on Mars. There are also links
to the NASA Kids website and a search-a-word puzzle containing words
associated with Mars.

The web page will be online soon at the following address:

<http://academy.gsfc.nasa.gov/2004/research/group_project/module/index.html>

![](media/image36.png){width="5.004861111111111in"
height="3.5395833333333333in"}

[]{#_Ref79743964 .anchor}

**Figure 27: Homepage for the MSCLE scout mission.**

![](media/image37.png){width="4.996527777777778in" height="3.3125in"}

[]{#_Ref79744377 .anchor}

Figure 28 : A portion of the page that asks visitors to guess which
pictures are from Mars and which are from Earth. Holding the cursor over
\"Answers\" uncovers the mystery.

## Demonstrate the Benefits

As a result of trying to accomplish a mission, new technologies have to
be developed along the way in order to obtain everything necessary to be
successful. Going to Mars in search of life is revolutionary and opens
up a new frontier of opportunity. By making the public aware of the
possibilities that exist and exciting them about the mysterious unknown
and uncharted, people can become motivated to think outside the box and
want to get involved. A short video clip has been made that outlines a
few previous missions and their spin-offs and poses the concept that
there are so many things that are not yet conceived that could perspire
from the technology advances of this mission.

The first spin-off that is used for demonstration is rubber in tennis
shoes. The technology used in the bottoms of certain sneakers was
originally developed for the boots of the Apollo astronaut space suits.

![](media/image38.jpeg){width="2.1069444444444443in"
height="1.6930555555555555in"}

**Figure 3: Apollo boot footprint on the moon**[^90]

The next spin-off is a new technology for better detecting breast cancer
that came about from the Hubble Space telescope project. Space Telescope
Imaging Spectrograph (STIS), a system used to detect black holes has
been modified and developed as a very sophisticated method for detailed
imaging and is used as a painless detection method.

![](media/image39.png){width="3.125in" height="2.0833333333333335in"}

**Figure 4: A depiction of a STIS**[^91]

After the two of these spin-offs have been presented then the film shows
pictures of people and poses the question, what if we were to go to Mars
to detect life, just imagine what the out come has the potential to be.

## Inspire and Excite

A poster/flyer type advertisement has been created in order to excite
the imagination. The mission will be portrayed in the form of a runner,
the starting point being at Earth and the finishing point being on Mars.
The picture contains a sphere that is on the left side a depiction of
the Earth and on the right side a depiction of Mars. The runner will be
seen running from the left side of the sphere to the right, running from
Earth to Mars. The statement, the prize lies in the journey not the
destination, will appear on the bottom of the picture. The MSCLE logo
will be displayed on the poster as well as a few words about the
program. See below for a copy of this advertisement.

![](media/image40.jpeg){width="4.850694444444445in"
height="5.915972222222222in"}

# Conclusions

For each team, myriad lessons lay in the aftermath of the preliminary
design phase of MSCLE. In a useful light, they may be cast as future
plans and the next steps to honing, toning, and further improving MSCLE.

## Lessons and Future Development

*Science*

Specification and further research into the instrumentation includes
physical (geometric) complementarity, continuing searches for
state-of-the-art materials and methods, and the integration with data
analysis software. Closely linked with those aspects are the sample the
experimentation schedule and the sample acquisition method. A top-down
explanation of the entire process from the mission, to the 14-day
sequence of the PGEEx, to each sample is necessary; a detailed set of
steps for the sample analysis sequence is a main feature of this overall
schedule. Finally, a set of contingencies for additional and limited
experimentation will be formulated. In the event that the objectives are
fulfilled with time and power remaining, there will be plans to exhaust
the resources; in the case that the mission is altered to a state
impairing the fulfillment of the objectives, we must have in place a
process to get the most out of the mission.

*Technical Development*

Aside from specification of the physical payload design and of the exact
equipment required by each subsystem, the next greatest task is a
process for sample acquisition. For the two former aspects, normal
progress through another design phase will suffice; for the latter,
intimate corroboration with science team may produce the most efficient
results. Finally, a new subsystem lead will be charged with seeking
state-of-the-art flight-ready technology to further pull down the
mission cost and increase the mission efficiency.

*Budget*

Monetary concerns will persist throughout design. Effective cost
controls will come from good communication with the technical and
scientific teams, as well as more specific estimation tools, and a more
rigid schedule of analysis. NAFCOM is a much more accurate and
detail-oriented software estimating package. At timely specified
junctures, budget estimations will occur to ensure design remains on
course. Also, the budget team will provide backup in checking costs for
engineering subsystem leads; this will provide redundancy and a safety
net.

*Outreach*

Along with spreading the word about MSCLE and space exploration, the
team must well-define the targeted audiences and more specific plans for
reaching them. Possibilities for the latter include finalizing and
constantly updating the web site, airing commercials and educational
segments on public television and radio.

This report will be made available to all parties seeking assistance or
information for their mission development. As an example or stepping
stone for other exploration teams, MSCLE is a source of technical
information, imaginative science and engineering, and novel public
outreach aspiration. Moreover, the project has inspired many of its
researchers to further utilize MSCLE as a tool of education beyond the
scope of mission preparation.

*For Proposal*

The next steps bifurcate into two branches: (1) increased readiness for
proposal and (2) enhancement as teaching tool for public or other
research groups. For the 2011 launch cycle, MSCLE will require further
preparation and endorsement by 2006. Prior to seeking support, MSCLE's
technical sophistication and adherence to Scout Mission guidelines
require improvement. University astrobiological programs (i.e.,
University of Washington) are growing nationwide, and they will be the
first point of research for possible support. Other programs focusing on
geological and planetary sciences are the next targets.

# Acknowledgements

First and foremost, thanks to Guinness, the Space Ferret for his
inspiration and wily ways throughout this challenging and rewarding
process. As our muse, he added tons of fun, in a wholesome, family
channel way; as our friend and confidant, he drew our attention to
life's priorities---family and our homies.

Dr. Herbert Frey has been instrumental as a mentor and guide, much like
Obi Wan Kenobi. The dark side, however, will not receive our hearts,
except perhaps for the already-tainted Canadian. We gained from Dr.
Frey's experience and wisdom not only technical knowledge, but an
exploratory attitude mediated by caution: "You've chosen a noble goal,
\[Luke, but you may have to use the force to even get you close to under
budget\]."

Thanks for the support of all the NASA Academy staff. We hate the hard
deadlines, but otherwise we'd be sipping tea in Shangri La instead of
having completed a fun team project. Nonetheless, you've got it coming.

Further, thanks to the entire NASA community without whose worm and
meatball, we would not fully appreciate the beauty of a logo.

Finally, a growing exhaustion actively corrected concepts of position
and portion until a final product manifested.  The viability of that
product is a testament to the stamina of the constructing party.

From the bottoms of all our hearts, bliddy.

# Appendix A

*Structure Subsystem Calculations*

*Table 26* lists the steps and equations used to estimate the various
masses of the MSCLE spacecraft. The percentages are located in the
comments column.

  -------------------------------- ------------------------ --------------------------
  []{#__RefHeading___Toc81459621                            
  .anchor}*Table 266 - Mass Budget                          
  Calculations*[^92]                                        

  *[Element]{.underline}*          *[Weight]{.underline}*   *[Comments]{.underline}*

  **Payload**                      M~P/L~                   Provided by science group

                                                            

  **Spacecraft Subsystems**        M~SS~                    Equal to subsystem total

  **Propulsion**                   M~propulsion~            Provided by subsystem

  **Attitude Control**             M~gnc~                   Provided by subsystem

  **Command & Data                 M~cadh~                  Provided by subsystem
  Handling/Communication**                                  

  **Thermal**                      M~th~                    2% to 5% of M~dry~, 2%
                                                            used in calculation

  **Power**                        M~ep~                    Provided by subsystem

  **Atmospheric Entry**            M~atm~                   45% of (M~dry~ -
                                                            M~propulsion~)

  **Structure & Mechanisms**       M~sam~                   15% to 25% of M~dry~, 20%
                                                            used in calculation

                                                            

  **Margin**                       M~mar~                   5% to 25% of M~dry~, 20%
                                                            used in calculation

                                                            

  **Spacecraft Dry Weight**        M~dry~ = M~P/L~ +        
                                   M~SS~ + M~mar~ - M~atm~  

                                                            

  **Propellant**                   M~prop~                  

  ***Loaded Weight***              M~loaded~ = M~dry~ +     
                                   M~prop~+ M~atm~          

  ***Kick Stage***                 M~kick~                  

  ***Injected Weight***            M~inj~ = M~loaded~ +     
                                   M~kick~                  

  ***Adapter***                    M~adapter~               

  ***Booster Weight***             M~boosted~ = M~in~+      
                                   M~adapter~               
  -------------------------------- ------------------------ --------------------------

The steps listed describe how the spacecraft's volume was
approximated[^93]:

i.  Dry Mass: 181.21 kg

ii. Estimate Propellant Mass: 50 kg

iii. Estimate Atmospheric Mass: 61.29 kg

iv. Find Spacecraft Loaded Mass: 181.21 + 50 + 61.29 = 297.50 kg

v.  Estimate Spacecraft Volume:

    a.  Divide Loaded Mass by Estimated Density

    b.  Density Range: 20-172 kg/m^3^ = 14.88-1.73 kg/m^3^

    c.  Average Density: 79 kg/m^3^ = 3.77 kg/m^3^

# Appendix B: Organization

## Organization Chart

## ![](media/image41.jpeg){width="5.365972222222222in" height="7.598611111111111in"} {#section .unnumbered}

## Engineering Team Overview

The engineering team is comprised of eight different subsystems, each
corresponding to a specific aspect of the project. Each group was
responsible for both the transit of the spacecraft to Mars and the
descent of the lander to the surface, as well as the operation of the
lander on the planet surface. The Chief Engineer (CE) was responsible
for the overall functioning of the engineering team and reported to the
project manager.

### Chief Engineer/Systems Coordination

The CE was responsible for the operation of the engineering group. This
includes developing agendas for scheduling and running meetings of the
engineering team. He was also responsible for establishing deadlines,
and ensuring the productivity and timeline of the engineering team.

The CE was also responsible for the overall systems integration of the
spacecraft and ensuring that all of the systems interact in a manner
that is most beneficial to the mission.

### Structures

The structures subsystem was responsible for the structural design of
the spacecraft, including the physical design of the spacecraft.
Structures was primarily concerned with the development of the overall
mass budget of the spacecraft.

### Payload

This subsystem addresses the question of how the surface operations will
be accomplished to reach the scientific goal. The payload subsystem was
responsible for not only the drill but also for incorporating the
scientific instruments into the spacecraft. This includes working to
ensure that the samples are not contaminated or compromised before they
are analyzed.

### Power

The Power subsystem group addresses the generation, distribution, and
budgeting of power throughout the spacecraft.

### Flight Dynamics

The concern of the Flight Dynamics group was the determination of the
proper orbital trajectory for transit from Earth to Mars. This included
developing the burn strategy and the delta-V requirements for the
mission.

### Propulsion

Propulsion was responsible for the development of the propulsion
strategy throughout the entire mission. This includes thruster
selection, fuel management, and other propulsion needs. The propulsion
group was also responsible for the selection of a suitable launch
vehicle.

### Communications and Data Handling

The Communications and Data Handling (C&DH) group was responsible for
the processing, storage and transmission of all data and communications
from the spacecraft. This includes both scientific data and housekeeping
information throughout the flight.

### Ground Operations

Ground Operations will develop an operating strategy for the spacecraft
during both transit and science operations. This includes labor needs,
personnel scheduling and the determination of the needed ground assets
including the use of the DSN.

### Subsystem Organizational Changes since Progress Report

A series of changes have been made to the organization of the
engineering team. These are detailed below.

*Guidance Navigation and Control*

The name of the Guidance Navigation and Control subsystem has been
changed to Flight Dynamics subsystem. This better reflects the work that
was being done in trajectory planning and reflects the decreased role of
attitude control in this initial analysis.

*Structures*

The structures subsystem assumed control of the mass budget from the
chief engineer. This work tied in closely with the work that was being
done in structures.

*Thermal Environment*

The thermal / environmental team which was responsible for ensuring that
the spacecraft was able to operate in the thermal and radiation
environments that it would experience was disbanded due to manpower
issues. The responsibilities of this subsystem were spread among all of
the remaining subsystems. Each subsystem was thus responsible for
ensuring that thermal issues incorporated into the design.

# Appendix C: Glossary

*Angle data*: a report on antenna-pointing that provides the azimuth and
elevation of the antenna [^94]

*BLGA* : Backshell Low Gain Antenna

*Chiral* (Chirality): The state in which an object cannot be mapped onto
its mirror image by rotations or translations; an object has chirality
if its structure and 3-dimentional configuration is always chiral
regardless of how the molecule is conformed; often referred to as
"handedness" [^95]

*CLGA*: Cruise Low Gain Antenna

*CMGA*: Cruise Medium Gain Antenna

*Doppler* *data*: a count of the number of cycles of a spacecraft's
carrier frequency during a certain time period; its precision depends on
the signal strength and can be as small as a fraction of a cycle [^96]

*DOR*: differential one-way ranging

*DRAM*: Dynamic Random Access Memory

*DSN*: Deep Space Network

*DTE*: Direct to Earth transmission

*EEPROM*: Electrically Erasable Programmable Read Only Memory

*EDL*: Entry, Decent, Landing

*Graben*: A geologic, structural feature in which a block of crustal
rocks is downthrown and bordered on both sides by normal faults; very
common in rift valleys[^97]

*HGA*: High Gain Antenna

*HGAG*: High Gain Antenna Gimble

*L~eff~* : Effective Gate Length

*LGA*: Low Gain Antenna

*PCI*: Peripheral Component Interface

PGEEx: Passive Gas Exchange Experiment

*Pyrolysis*: Chemical decomposition occurring as a result of high
temperature; a form of incineration that chemically decomposes organic
material, typically under pressure and at high temperatures [^98]

*Ranging* data: a measure of the time required for the round trip of a
signal going to the spacecraft and returning to the earth station [^99]

*Scarp*: A general term for a steep slope or cliff found at the margin
of a gently sloping area; often formed by faulting and subsequent
erosion [^100]

*UHF*: ultra high frequency

*Volatile*: A substance that easily passes into a gaseous state with the
application of heat and evaporates readily at normal temperatures and
pressures [^101]

# Endnotes

[^1]: "Exploring Mars" http://www.exploringmars.com/history/1600.html.

[^2]: "Mars Exploration Program Analysis Group"
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[^3]: http://www.astronomie.de/vds/astronomietag-2003/alh84001.jpg.

[^4]: Lottman, Tim. [Deinococcus radiodurans]{.underline}, NOV 2002:
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[^5]: Drs. Cano, Raul, Lambert, Chip, and Orkand, Paula. [Nature's
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[^6]: ^5^ Long, Jeff, Tribune Environmental Writer, OCT 2000,
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[^7]: ^6^ Microbial Wonders of the Deep, [Thermophiles]{.underline},
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[^8]: Lorenson, Tom. And Kvenvolden, Keith, JUN 2001. [USGC Workshop on
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[^9]: Facts about Ammonia ... ,
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[^10]: Environment Health Center, [Formaldehyde]{.underline},
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[^11]: http://www.windows.ucar.edu/mars/images/MOC_clouds_Vallis_M.jpg.

[^12]: http://pds.jl.nasa.gov/planetsimages/browse/mars/marscany.jpg.

[^13]: http://marsoweb.nas.nasa.gov/landingsites/mer2003/topsites/final/Melas/maps/Melas_small.jpg.

[^14]: http://antwrp.gsfc.nasa.gov/apod/ap021216.html.

[^15]: http://www.flatoday.com/galleries/viking/072001viking4.jpg.

[^16]: http://marsoweb.nas.nasa.gov/landingsites/mer2003/doc/pasadena_02/talks/parker/Melas_geomorph/jpg.

[^17]: http://marsoweb.nas.nasa.gov/landingsites/mer2003/topsites/final/Melas/maps/Melas_small.jpg.

[^18]: Ch 2: The Search for Life on Mars, Report from the ESA Exobiology
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[^19]: National Research Council, Assessment of Mars Science and Mission
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[^20]: Levin, G., Kuznetz, L., and Lafleur, A.; \"Approaches to
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[^21]: ESA, 1997-1998.

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[^23]: ESA, 1997-1998.

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[^25]: NASA Astrobiology Magazine. \"Fleshing Out Martian Proteins:
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[^27]: Exobiology Program Office, NASA Headquarters. \"An Exobiological
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[^28]: National Aeronautics and Space Administration. *SP-4212 On Mars:
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    ^21^ http://
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