Paradigm
Shift in Design for NASA’s New Exploration Initiative
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16.89 Graduate Design Class
Space Systems Engineering
Massachusetts Institute of Technology
May 12, 2004
16.89 Team Members
Students
Sophie Adenot
Julie Arnold
Ryan Boas
David Broniatowski
Sandro Catanzaro
Jessica Edmonds
Alexa Figgess
Rikin Gandhi
Chris Hynes
Dan Kwon
Andrew Long
Jose Lopez-Urdiales
Bill Nadir
Geoffrey Reber
Matt Richards
Matt Silver
Ben Solish
Christine Taylor
Staff
Professor Jeff Hoffman
Professor Ed Crawley
Professor Oli de Weck
2.1 Elements of Sustainability
2.1.2 Budgetary Sustainability
2.1.3 Organizational Sustainability
2.1.4 Technical Sustainability
2.2 Sustainable Exploration Systems
– Dynamics
2.3 Sustainability, Flexibility,
Robustness
2.4 Extensibility – An Enabler of
Sustainability
2.4.1 Reasons for Extensibility
2.4.2 Describing Extensibility
2.4.3 Principles Supporting
Extensibility
2.5 Historical Comparison: Antarctic
Exploration
2.5.1 Technology and Logistics:
2.6 Designing for Sustainability: A
Process
3.
Knowledge Delivery: The Core of
Exploration
3.5 Knowledge Delivery Process Map
3.8
Knowledge Drivers: Apollo Case Study
4.1 Brief Description of Formal
Elements
4.2.3 Requirements and Assumptions
4.2.4 Operational View of Lunar
Baseline Missions
4.2.5 Commonality within Moon
Missions
4.2.6 Discussion of Lunar Baseline
Missions
4.2.7 Scientific and Resource
Knowledge
4.2.8 Knowledge Delivery
Infrastructure
4.3.1 Literature Review – A Brief
History of Mars Mission Designs
4.3.4 Knowledge Delivery
Infrastructure
4.4.2. Summary of Baseline Forms
5.
Commonality Across Missions
6.2 Decision Analysis Using
Multiattribute Utility Theory
6.3.2 Example: Staged vs. Cycler
Transportation System Design
7.2 Reasons for scenario-based
planning
7.3.3 Dawn of the Nuclear Propulsion
Age
7.3.6 Little Green Martian Cells
9.1.3 Elements of the Heavy Cargo
Shuttle Derived Vehicles Study
9.1.5 Solid Rocket Booster derived
launcher considerations
9.1.7 STS derived assembly platform
9.1.8 LabView tool for evaluating
launch capabilities
9.3 Parameters for Calculating Lunar
Mission Mass in LEO
9.4 Mars Initial Mass in LEO
Calculations
9.4.1 Verification of initial mass
in LEO estimates
9.4.2 Example Calculation of Initial
mass in LEO
9.5 Knowledge Transport Calculations
and Architecture
9.5.3 Optical Communication Trades
9.5.4 Mars Science Details
(Knowledge)
9.5.5 Additional Knowledge Materials
(background)
Figure 1: Proposed space systems
design process
Figure 2: NASA budgetary fluctuations in 1996 dollars (courtesy
http://history.nasa.gov)
Figure 3: Interaction of political, organizational, and technical
factors
Figure 5: Boehm's model of spiral
development (picture from Boehm, 1988)
Figure 6: Change in system need
and capability over time
Figure 7: Positive feedback loop for exploration
Figure 8: Space systems design process
Figure 9: Value delivery to
scientists diagram
Figure 10: Value delivery to
technologist/explorers diagram
Figure 11: Knowledge delivery system
OPM (Crawley, 2004)
Figure 12: Five types of knowledge
Figure 13: Example of the quantity
scientific knowledge from Hubble (Beckwith, 2003)
Figure 14: Time and spatial synergy
for robotic and human explorers
Figure 15: Carriers of knowledge
Figure 16: Theoretical news value as
the space exploration system evolves
Figure 17: Knowledge delivery cycle
Figure 18: Knowledge delivery time
examples
Figure 19: Knowledge potential:
maximum exploration coverage per day versus number of crew
Figure 20: Expanding the exploration
potential using a remote base (Hoffman, 1998)
Figure 21: Apollo knowledge drivers
Figure 23: Operational view of Short
Stay Lunar Mission
Figure 24: Operational view of
Medium Stay Lunar Mission
Figure 25: Operational view of
Extended Stay Lunar Mission
Figure 27: Short stay mission to
Mars
Figure 28: Extended stay mission to Mars
Figure 29: Schematic representation
of the Moon and Mars Baseline missions. 84
Figure 30: Mars/Moon Transfer
Vehicle (MTV)
Figure 31: Functional requirements
for a Crew Operations Vehicle
Figure 32: Functional requirements for
a Modern Command Module
Figure 33: Functional requirements
for a Habitation Module
Figure 34: Functional requirements for
a Crew Service Module
Figure 35: Functional requirements
for a Moon/Mars Lander
Figure 36: Flow diagram describing
elements of extensibility in integrated baseline
Figure 37: Decision analysis tree.
Figure 39: Decision tree for L1 capability example
Figure 40: Value of L1 capability
Figure 42: Minimum LEO payload mass penalty for EELV tower escape
Figure 43: Launch escape mass as a function of crew module mass (Source:
Orbital Science Corp.)
Figure 44: Entry vehicle shape
pair-wise option comparison
Figure 45: Comparison scale for
entry vehicle
Figure 46: Parametric comparison
of inflatable versus conventional Earth re-entry technology
Figure 47: EDL pair-wise option
comparison
Figure 48: Mission segmentation
Figure 49: Elements of the MTV, assuming a crew of three for a ten-day
mission
Figure 50: Classification of existing crew transport modules
Figure 51: Configuration masses (10-day to 40-day missions)
Figure 52: Three COV configurations for launch from Earth to LEO
Figure 53: Mars/Moon Transfer Vehicle (MTV)
Figure 54: Historical space habitat pressurized volume (Kennedy, 2002)
Figure 55: Flowchart of scaling analysis
Figure 56: Vehicle mass scaling (broken line: 3-day mission, solid line:
30-day mission)
Figure 57: The reality of
designing an EDL system (Amend, 2004)
Figure 58: Trade space for EDLA
missions (Larson, 1999)
Figure 59: Earth return capsule
design
Figure 60: Lunar Lander design
Figure 61: Martian Lander design
Figure 62: NASA’s missions and
“smart” landing technologies roadmap (Thurman, 2003)
Figure 63: Comparison of Mass in LEO for Different Missions
Figure 64: Mass in LEO for mission to lunar pole with free-return
trajectory requirement
Figure 65: Comparison of a non-reusable and reusable Lunar Lander
Figure 66: Comparison of nuclear
propulsion to chemical propulsion for baseline trajectories
Figure 67: Initial Mass in LEO for Various Mission Architectures
Figure 68: Comparison of Opposition-class mission with and without a
Venus fly-by
Figure 69: Comparison of
Conjunction-class missions
Figure 70: Comparison of Mars trajectories
Figure 71: Interface used for the
Excel CEV model
Figure 72: Linking possibilities
among CEV options and ranking criteria and weights
Figure 73: OASIS CTV Internal Layout
Figure 74: NASA Habitable Volume Standard 8.6.2.1
Figure 75: Habitable volume for
various crew sizes as a function of mission duration
Figure 77: HPM upper section
material
Figure 78: HPM lower section
material
Figure 79: Apollo CM schematic
Figure 80: Shuttle-C elements (Source: NASA)
Figure 83: Ariane V and STS-Derived
Figure 84: STS derived assembly
platform
Figure 85: GUI interface for the
LabView combination tool
Figure 86: Mass margin to ISS for
999 options of launch + CEV configurations
Figure 87: Atmospheric control and
supply (Wieland, 1999)
Figure 88: Water recovery and
management (Wieland, 1999)
Figure 89: Mass and volume of ECLSS
atmosphere and water management systems
Figure 90: Attitude control modes, from
Larson (1999)
Figure 91: Apollo lander mass
breakdown, from Gavin (2003)
Figure 92: Diagram of opposition class mission with a
Venus fly-by (NASA DRM website)
Figure 93: Diagram of conjunction class mission (NASA
DRM website)
Figure 94: Diagram of fast-transfer conjunction class
mission (NASA DRM website)
Figure 95: Communication
Architecture
Figure 97: Apollo landing
sites. Near side of the Moon, center (0,
0).
Figure 99: Far side of the Moon.
Table 1: Knowledge delivery process
Table
2: Apollo mission details (NASA website, 2004)
Table
3: Knowledge drivers model parameters
Table
4: Architectural space transportation forms
Table
5: ΔV requirements assuming parachutes and aerobraking not used
Table
6: ΔV requirements assuming parachutes used
Table 7: Expected utilities from the
Decision Analysis tree for the L1 capability decision
Table 8:
Staged vs. Cycler transportation vehicle design
Table 9:
Staged vs. Cycler design comparison with aerobraking
Table 10:
Staged vs. Cycler design comparison with the pre-positioning of return
fuel
Table 11:
Staged vs. Cycler design comparison with aerobraking and pre-position
return fuel
Table 12:
EDL option ranking and system mass for an Apollo-class Earth re-entry
vehicle
Table 13:
Rover functional requirements
Table 14:
Baseline module masses
Table 15:
Mass benefit using pre-positioning for a Medium Moon mission
Table 16:
Mass benefit using pre-positioning for an Extended Mars mission
Table 17:
Propulsive Δv
requirements for Martian and lunar EDLA
Table 18:
Integrated Lunar and Martian Lander functionality requirements
Table 19:
Three and six-person Lander component mass comparison
Table 20: Suggested landing sites
Table
21: CTV mass estimation (OASIS, 2001)
Table
22: Apollo CM mass breakdown (http://www.astronautix.com/craft/apolocsm.htm)
Table
23: Mass requirements in LEO (ISU SSP Report 99’)
Table
24: Various STS-derived options
Table
25: Various STS-derived options
Table
26: Various combinations
Table
27: Form/Function matrix
Table
28: ECLSS atmosphere management
Table
29: Design process of ADCS
Table
30: Description of actuators, inspired by de Weck (2001) and Larson (1999)
Table
31: ADCS masses for some crew vehicles
Table
32: ADCS mass of communications satellite, from Springmann (2003)
Table 34: DV table for lunar missions using lunar orbit
Table 35: DV table for lunar missions using EM-L1
Table 36: Lunar payload masses
Table 37: Other lunar mission parameters
Table
38 : Mission class overview
Table
39: Comparison of opposition class mass estimates with Walberg
Table
40: Comparison of conjunction class mass estimates with Walberg
Table
41: Comparison of fast-transfer mass
estimates with Walberg
Table
42: Comparison of IMLEO estimates with
Walberg
Table
44: Moon resources - preliminary findings (Taylor, 2001)
Table
45: Methods of creating geophysical networks (LExSWG, 1995)
Table
46: Knowledge levels and instrumentation for a moon mission (Geoscience, 1988)
On January 14, 2004, President George W. Bush presented the nation with a bold new initiative to “explore space and extend a human presence across our solar system…using existing programs and personnel…one mission, one voyage, one landing at a time.” (Bush, 2004) NASA was charged with the task of developing a sustainable and affordable human space exploration program with the initial objective of returning a human presence to the Moon by the year 2020. The directive thus raises two broad engineering questions: First, what is the purpose of an exploration system, and how one evaluates its performance. Second, how does one architect a sustainable space exploration system? The following report makes the case that the primary purpose of an exploration system is the delivery of knowledge to the stakeholders, and that the design should be evaluated with respect to knowledge.
On January 14, 2004
President George W. Bush presented the nation with a new vision for space. The
National Aeronautics and Space Administration (NASA) will develop a sustainable
human space exploration program taking humans back to the Moon by 2020, and
eventually to Mars and beyond (Bush, 2004). The vision, and plan that goes with
it, calls for the completion of the ISS, the retirement of the Space Shuttle by
2010, and the development of a new Crew Exploration Vehicle (CEV). Bush’s
vision provides a bold push towards mankind’s traversing of the solar system.
The following report, representing the culmination of MIT’s 2004 spring 16.89
graduate design class, presents a design methodology and conceptual tools to
facilitate the achievement of this vision. It addresses two critical questions
facing the space community: What is sustainability in the context of space
systems? How can sustainability be provided for during conceptual design? The
following report addresses these questions. In doing so, it demonstrates that
an exploration program is by definition a knowledge
acquisition and transfer system, and it presents a process by which one may
design for sustainability.
The goal of exploration is knowledge
While the
motivation behind exploration has varied throughout history, the primary
function of any “exploration system” has been to discover the unknown, to gain
knowledge. Some of the more common ways to gain knowledge have been through the
use of visual, electrical, or physical transportation of information. A simple
example of a space knowledge transfer system is the human eye. The human eye
gathers knowledge in the form of light. Several hundred years ago mankind
developed the telescope in a hope to improve upon the amount of knowledge
delivered to the eye through the discovery of magnification. The magnification
of objects resulted in a higher order of knowledge resolution and consequently
more information about space was discovered.
More recently
mankind has sent satellites and drones into the solar system, with sensors that
can gather information unattainable by the human eye alone. Information
gathered by these systems is sent back to Earth through the use of electrical
transmissions where it is turn into knowledge. A number of characteristics
increases the “knowledge resolution” of these satellites and drones compared to
telescopes, including: Shorter distance between optics and target, physical
contact, sample return, in-situ analysis, etc.
It is noteworthy that order to achieve this higher of knowledge resolution,
mankind had move beyond light as the sole transfer-mechanism, to in-situ
measurement and mass transport. Future exploration systems must necessarily
follow this trend, exploiting the duality between mass and knowledge transfer,
with one critical improvement--humans will provide degree of knowledge
resolution previously unimaginable with satellites, drones, and telescopes
alone.
No matter the form
of the space exploration system (human eye, telescope, robotic probe, or human
contact), the end product of the exploration system is knowledge. Currently,
the majority of the work being completed on NASA’s new initiative is directed
towards a new exploration vehicle. The class believes that any new space
vehicle developed by NASA must be designed with an understanding that it will
be but one tool in system whose ultimate function is to gather and transfer
knowledge in space and on Earth.
To say that an
exploration system must deliver knowledge to achieve its goal is to recognize
that while mass transport enables exploration, the ultimate success of an
expedition depends on the acquisition, communication, and synthesis of visual
imagery, scientific data, and human experience to key stakeholders. This
suggests revaluing traditional space system characteristics and trades to
account for the demands of knowledge acquisition and delivery. Further, in
order to make clear decisions about system capabilities and mission goals,
attributes of knowledge must be categorized and valued in accordance with
stakeholder needs. System designers must have a firm grasp of the knowledge
delivery process, and establish how it will occur at each point in the system’s
lifecycle.
Sustainability in the Design Process
Before knowledge
can be incorporated into system valuation and trades, however, there must be a
clear understanding of what is a sustainable space system and how can this can
be addressed during conceptual design? Current space system design methods are
not geared towards enhancing “sustainability.” Traditionally, they have focused
on developing requirements, conducting trades based on assumptions about the
future, and then optimizing the system with regard to some metric. Results are
commonly single point designs optimized for single missions.
While such methods
have proven adequate for low-frequency missions, they rely on assumptions about
an uncertain future. A design that is optimal at one point in time may become
less optimal in the future. Due to the expected duration of the new exploration
initiative, major investments should note be made based on unverified
assumptions. The new exploration system should be designed so that it can
respond to changes in the future. The approach to design described in this
report addresses this problem. Using an iterative process, and emerging system
valuation tools, it creates a rigorous development strategy which is flexible
and robust to environmental changes.
Chapter two
proposes a definition of sustainability. Drawing from recent scholarship and
historical examples, it argues that sustainable exploration programs must first
and foremost have the capability to manage various kinds of uncertainty,
including policy, budgetary, technical, and logistical changes. Conceptual
designs must provide system operators with the ability to anticipate and
capitalize on emerging opportunities and positive feedback loops while
simultaneously adapting to changing value-structures and external
circumstances.
Properties that
enable sustainability have been termed flexibility, extensibility, robustness,
and commonality. Much recent scholarship has addressed the need to rigorously
value these system properties for the purposes of design. Generally, these
properties translate to formal architectural attributes, such as modularity and
platforming, as well as operational attributes such as staged deployment and
spiral development. Chapter three defines these terms in the context of space
systems, and presents methods for their formalization in system architecture.
There are two ways
in which flexibility and extensibility are introduced and evaluated:
mathematical evaluation methods and architecture design considerations. The
mathematical evaluation methods used are based upon decision analysis, real
options theory, and scenario planning. The architectural design considerations
are commonality, scalable systems, and modularity. Both methods evaluate a
given system based on the resulting value of knowledge delivered by the system.
Notice that the system is not evaluated on cost or mass, but on knowledge,
which is the primary purpose of an exploration system.
A major aspect of
this study involves identifying a process to combine these properties and
methods can be systematically incorporated into system design. Part of the
solution involves creating a strategy, rather than a point design, that can
react to change. Chapter 4 presents an example strategy, or “baseline,” which
was conceived through an iterative process of design, needs mapping, and
synthesis of sub-strategies. Sub-strategies consist of small, medium, and large
Moon and Mars expeditions, each designed with principles of extensibility such
as commonality and staged deployment. Individually, these missions are rough
“point-designs.” However, major architectural decisions in each reflect
anticipation of gradually increasing mission scale, and eventual transit to
Mars.
After completing
the sub-strategies, areas of functional commonality and uniqueness can be
anticipated across the system, and architectural forms refined appropriately.
The resulting forms and operations can then be synthesized into an integrated
life-cycle strategy, with options for reacting to uncertainty. The following
schematic illustrates the design process used:
In developing the
integrated baseline, commonality trades at the formal and operational level
become necessary. Chapter 5 details such trade studies and their results.
Once the final
version of the baseline strategy and associated trades has been developed, more
rigorous tools may be applied to determine when, and under which circumstances
different design options become valuable. For example, the decision to transit
through the Earth-Moon Lagrangian Libration Point 1 (EM-L1) while en route to
the Moon may not be optimal for a single mission to the lunar equatorial
region. However, if the frequency of non-equatorial lunar missions is
sufficiently high, the option of utilizing EM-L1 becomes increasingly valuable.
Tools, including modified forms of real options valuation, can inform trade
studies of this sort, resulting in up-front design decisions that drastically
reduce life-cycle cost and increase system flexibility.
Chapter 6
introduces such tools and methods. Scenario planning is applied to the
integrated strategy to examine how the system can react to environmental
changes. Adjustments are then suggested, based upon the baseline’s reaction to
the scenarios. Decision analysis and Real Options analysis techniques are also
used to determine at what point time-critical decisions should be made in the
execution of the baseline strategy, and which investments should be made now to
allow for the option of adapting to future uncertainty.
What exactly is a
sustainable exploration program? In one sense, the answer is rather
straightforward. To “sustain” means literally: to maintain in existence, to
provide for, to support from below (Dictionary.com website). At the
programmatic level, an exploration system will be maintained in existence so
long as it is funded, and it will be funded provided it meets the needs of key
stakeholders, members of Congress, the Administration, and ultimately the
American people. Realistically, however, system designers must recognize that
these needs themselves will change. A multi-year, multi-billion dollar program
in the US Government must expect to face a great deal of uncertainty with
respect to objectives, budget allocations, and technical performance.
In order for an
exploration system to be sustainable, then, it must be able to operate in an
environment of considerable uncertainty throughout its life-cycle. Designing
for sustainability implies identifying sources of uncertainty and managing them
through up-front system attributes. Various terms have been used to describe
such system attributes, including: flexibility, robustness, and extensibility.
While a large
complex system must react to changing environments in order to be sustainable,
technological aspects of systems can themselves impact the environment. Once in
development and operation, a multi-billion-dollar system will mediate political
interests, organizational decisions, and technical alternatives, creating
potential sources of stability and positive feedback-loops, as well as sources
of uncertainty. Early decisions that create high switching costs or large
infrastructure sites, can “lock-in” architectural configurations and influence
the objectives and development path of later systems (Klein, 2000). A
sustainable design will be one in which, to the greatest extent possible, the
dynamics behind political, technical, and financial sources of stability
support, rather than hinder, system development and operations.
The following
chapter identifies three kinds of sustainability, and relates these to formal
system attributes. It reviews current thinking about flexibility and
extensibility, and their relation to architectural form. The chapter concludes
with a historical investigation of Antarctic exploration, drawing lessons for
the sustainability of exploration programs.
It is increasingly
evident that large, complex, technological systems cannot be conceived
independently from the political, economic, and organizational environment in
which they operate. While at a technical level, exploration is dependant on
continuous and reliable logistical support, at a programmatic level, political
and organizational factors greatly affect sustainability. With space activities
in particular, motivations and objectives can change rapidly compared to system
life-cycles, increasing the impact of political and organizational issues on
system development and use. A sustainable space exploration system will
successfully mediate and react to political, organizational, and technical
uncertainty, and also exploit, to the extent possible, sources of “stability”
that arise from the interaction of these factors.
Policy uncertainty
can take the form of changes in objectives or the regulatory environment in
which a system must operate. It stems from the dynamic nature of the
To take one
example, while the decision to build the Space Station Freedom was motivated
largely by Cold War concerns, the fall of the Berlin Wall transformed the
ailing project into a symbol for international peace and cooperation (Wikipedia,
2004). To the extent possible, system designers should consider the
implications of such changes for system operation. If a policy decision to
focus on Mars rather than the Moon is likely in the near term, current designs
should be extensible to both objectives. Similarly, if international
cooperation is based on uncertain agreements, alternatives to international
participation on the critical path of development should be available.
Shifting political
priorities also create changes in funding. During its years of development and
operations, a programs budget may oscillate unpredictably. Figure 1 illustrates how NASA’s budget fluctuates over time.
Figure 1:
NASA budgetary fluctuations in 1996 dollars (courtesy http://history.nasa.gov)
A flexible system
will maintain exploration capability even in the face of budgetary
fluctuations, whether through changes in schedule, scale of operations, or by
other means.
Recent scholarship
has investigated the relationship between organizational structure and
technical design. Charles Perrow (1984) has characterized socio-technical
systems in terms of their dynamics and complexity, drawing conclusions for
system safety and reliability. He defines space systems as highly coupled,
nonlinear, and complex. Organizational structure and technical complexity can
impact system reliability by creating “quite erroneous worlds in [the] minds”
of system operators and managers. (Perrow, 1984)
Diane (1996)
A space system will
be sustainable from an organizational perspective, then, if the technological
system and management structure are designed together to minimize
organizational drift and normalization of deviance.
Technical
sustainability refers to system performance, reliability, and the potential
infusion of new technologies. An
exploration system must support and maintain human and robotic activity at various
fronts of exploration, and incorporate technological advances to continuously
improve system performance without major operational changes. Further, any highly complex system is likely
to fail at some point during its life cycle. A sustainable system will be one
that is robust to failures, both small and large.
An important factor related to technical sustainability is risk
tolerance. Risk tolerance can be divided into three main areas:
By definition,
risk-free exploration does not exist. System designers must balance the risk
associated with architectural form, schedule, and operations, in order to
achieve system objectives. Risk tolerance can change throughout a system life
cycle, and thus change how a given system is operated.
While each of the
three domains above impacts the development and operations of complex systems
in different ways, they are closely interrelated. The dynamic relationship
between the three has important ramifications for sustainability. The relationship between these three broad
domains is shown in Figure
2.
Figure
2:
Interaction of political, organizational, and technical factors
.
The Columbia
Shuttle Accident Report (CAIB) repeatedly stresses the adverse affects that
broader issues such as indecisive national leadership, increasingly stretched
budgets, and continued mischaracterization of Shuttle capabilities have had on
NASA’s organizational and safety culture.
Conversely, Hans
Klein has suggested that the characteristics of a technological system and
development program can facilitate or impede coalition politics, thereby
reducing or exacerbating conflicts between politics and program administration
(Klein, 2000). Technology and politics are linked when program administrators
translate political forces into design requirements. Further, once developed, a
given system architecture together with its supporting facilities can become
“locked-in” and perpetuated through later designs. The space shuttle, for
example, made use of facilities designed partly as the result of short-term
political wrangling conducted during the Apollo era (p. 319[ESS1]).
Annalisa Weigel and
Daniel Hastings have similarly investigated the interrelation between technical
design and political change (2003).
Weigel and Hastings stress that space transportation infrastructures are
affected as much by political considerations as technical problems. It is thus
imperative to understand the coupling of both domains if a system is to operate
successfully in the “politico-technical” arena. Weigel presents a framework to
understand how policy directives couple with technical parameters. Figure 3 is an “influence diagram” used to illustrate such
coupling.
Figure 3: Translating policy parameter affects into
the technical domain: an influence diagram (courtesy, Weigel and Hastings,
2003)
At a different
level, as a later section of this chapter notes, the interplay between news,
politics, and technical development was an important factor in the evolution of
Antarctic exploration. In this respect,
designing for sustainability implies understanding how various design decisions
can lead to organizational and political dynamics that may improve or impeded
the flexibility of the system.
A sustainable
system will have attributes that allow it to cope with, or mediate, various
forms of uncertainty throughout its life-cycle. Many terms have been used to
define characteristics which give systems these properties. They include
flexibility, robustness, and extensibility.
But what are the relationships between these terms?
In many ways this
is simply a question of definition. Flexibility can be defined as the ability
of a system to change or be used differently than intended after it is
initially fielded. Flexibility can be
intentional, but is often unintentional such as in the case of the B-52 or the
use of the LM as a “life boat”. The speed with which a system reacts to change
is a measure of agility.
Extensibility is a particular kind of flexibility. Conversely, robustness is the property of a system
that allows it to be insensitive to change. A system is robust if it continues
to deliver value in changing circumstances.
All of these
“ilities” are enabled by attributes of architectural form. The follow schematic
illustrates how the various concepts relate to each other:
Extensibility is
defined as “the property that new elements can be added to a system in such a
way as to alter the value delivered.” (
Designing systems
for extensibility requires a fundamental shift in the way design decisions are
made, a shift from near optimal fulfillment of immediate requirements at
minimal cost, to minimizing life cycle cost, maximizing life-cycle performance,
etc. In other words, an extensible
design will not be the highest performing design when compared to a point design
optimized for a given set of capabilities- a penalty is placed on ultimate
system performance in order to increase life-cycle value. An extensible design will not be the lowest
cost design under the same conditions either.
The advantages of an extensible design are only realized in the context
of multiple generations of the system.
New metrics must be implemented for valuing the benefits of
extensibility. In addition, a culture
shift must occur from near term to longer-term expectations of success.
The large
investment associated with complex systems dictates the need for an
evolutionary growth path, although not all elements of the system undergo the
same degree of change. Therefore, it is
important to invest “extensibility dollars” only where needed. Investing in extensibility provides an option
for future change. As an example, an
in-space crewed exploration vehicle could be designed for extensibility in
terms of number of crew supported and days of support through decoupling of
living quarters with the command and control portion of the spacecraft. While the initial need may be support of a
four-person crew for two weeks, this need may extend to support of six people
for nine months. Clearly, using the same
vehicle for both missions would unduly penalize the shorter mission while design
of two separate vehicles would result in high costs associated with development
of redundant functions such as the command and control functions. Separating the habitat functions from command
functions through creation of two modules and a common interface, for instance,
would enable the habitation portion of the spacecraft to be easily
modified. If the change is executed, the
implementation of the change is expected to cost less than if the option had
not been put into place. If not
executed, the extensibility feature represents wasted resources in terms of the
expense to implement, reproduce and support the unused feature.
Several concepts
overlap almost directly with extensibility- staged deployment, and
spiral/incremental development. Staged deployment
seeks to match demand and supply through scaled rollout of a system. Expenses are delayed until a later date,
reducing the net present value of the expense and increasing the certainty of
the need, at the time of the expense. De
Weck et al. (2004) describe staged deployment as a potential alternative to
full deployment of the Iridium communications satellite network, with the
potential benefit being lower investment in order to start operations. Additional satellites could have been
deployed as demand increased. While
Iridium was ultimately displaced from most of the expected market due to
widespread cellular coverage, the lost investment could have been significantly
reduced.
Like staged
deployment, spiral development (Figure
4) is also an incremental method of deploying new
systems and their capabilities in a flexible manner. Initial capabilities are selected based on
prioritized goals, enabling quick deployment of high priority capabilities. Additional iterations of the process focus on
deploying lower priority capabilities and addressing newly discovered
needs. The result is quick deployment of
primary capabilities combined with risk reduction through decision delay that
enables incorporation of current technology into new stages and shifts in
strategy as needs become clearer (time advances).
Figure
4: Boehm's model of spiral development (picture
from Boehm, 1988)
Extensibility
reduces overall life-cycle cost and/or increases life-cycle performance through
a number of difference paths. Several
are listed below, along with brief descriptions.
As the lifetime of
a system grows, the rate of change of technology is increasingly mismatched
with the scale of system replacements.
Within a system, different modules have different rates of technological
change. Charles Fine (1998) uses the
term “clockspeed” to describe the rate of change and to highlight the
differences between rates of change.
Extensible systems allow for management of technological change within
the system. As an example, consider a
vehicle, such as a spacecraft. While
structural technology may undergo significant improvements on the timescale of
a decade or more; control system components, especially the electronic elements
such as logic chips, undergo significant change on an annual basis. Designing a system to accommodate varying
clockspeeds enables the design to evolve over time. One method for accommodating technological
change is through grouping components with similar rates of change into
modules, therefore, enabling easy replacement of the module, with minimal
impact to other areas. Ease of change
leads to the ability to keep a system modernized.
Delaying decisions
improves the likelihood of making a correct decision. While delay can cripple a program if not
handled properly, the result of effective use of delay is confidence in
decision-making.
Extensibility is
beneficial in the face of the uncertainties produced by the policy domain, and
the resulting budget fluctuations. The
potential for a change in President occurs once every four years, a timescale
much shorter than that of an exploration program. Given the mismatch in timescales, it is
critical that achievement of intermediate milestones provides lasting value, a
foundation for future work.
Methods are needed
for describing what an extensible system is and how the extensibility is
achieved. Ultimately, the metrics and
descriptions must be quantifiable to enable trades to be made between designs
and design options. Which system is more
extensible? How extensible is the
system?
One view of the
evolution of a system over time is a consideration of the relationship between
available capabilities and required capabilities; in other words, a type of
supply and demand curve. Figure 5 provides a notional view of this concept. The system needs over time are represented as
a continuous curve. While the system
needs curve may in fact be discrete, the aim here is to highlight the high
degree of changing need in relation to the ability of the system to
change. The design points represent the
available capability levels. From a
system performance standpoint, the ideal available capability would be a direct
overlay over the needs curve. While the
ideal curve cannot be reached due to practical considerations such as the cost
of each change (engineering, deployment, etc.), the ideal curve can be
approached through the creation of an extensible architecture. This view is closely related to previous work
in the area of staged deployment. (de
Weck et al., 2004).
Figure
5: Change in system need and
capability over time
The relationship
between the supply and demand “curves” is an important one. As Figure
5 illustrates, a system that is overly capable is
inefficient. More dollars and time have
been spent on unneeded functionality, at the given point in time. The reverse situation means that the system
is not meeting needs, also a problem. As
an example, consider the transition from Design 2 (D2) to Design 3 (D3). Before the transition, needs aren’t met by
capabilities, while after the transition, the system is over-designed, as would
be expected immediately after an improvement.
Also note the transition from D3 to D4.
While this transition was not required to meet new capabilities, since
needs have actually decreased, the change was made in order to maintain design
efficiency.
In order to analyze
the evolution of a system over time, a well-defined method of describing change
is needed. This void can be filled by a
series of operators, such as those defined by Baldwin and Clark (2000):
The above operators
can be used to perform all module-level operations. As was mentioned in the previous section, it
is critical to realize that evolution is synonymous with adaptation or change,
not addition. Continuous adaptation to
changing conditions may mean eliminating functionality that is no longer
needed; an operation accomplished with the exclusion operator. As a simple example of the use of operators,
consider the creation of a launch vehicle.
The augmentation operator is used to add strap-on boosters for heavy
lift capability, while the substitution operator could be used to express the
change of a launch fairing.
Four key principles
support extensibility- modularity, ideality/simplicity, independence, and
integrability. These principles were
originally linked to “flexibility” by Schulz and Fricke (1999) and are briefly
summarized below.
The first principle
supporting extensibility is modularity, defined by Baldwin and Clark (2000) as:
“A module is a unit
whose structural elements are powerfully connected among themselves and
relatively weakly connected to elements in other units. Clearly there are degrees of connection, thus
there are gradations of modularity.” (p.
63)
The principle of
modularity enables complex problems to be broken down through a hierarchical
structure. Changes internal to a module
are isolated at the module boundaries, limiting the cascading impacts of a
required change. Expense is reduced in
development, test, hardware exchange, etc.
Changes made to a modular system can be described in terms of the
modular operators described in the previous section.
Ideality is defined
by Schulz and Fricke (1999) as the ratio between useful and harmful/undesired
effects, a notion of design efficiency (pp. 1.A.2-4, as an additional
reference, see Suh, 2001.) This
principle highlights the importance of the ongoing culling of unneeded
functionality as a system evolves over time.
Failure to do so increases system complexity unnecessarily, eventually
making total replacement of the system a more effective option than change.
The independence
axiom derives from the independence axiom in axiomatic design (Suh, 2001). Each function is satisfied by a different
design parameter. Creating a decoupled
design, in terms of functionality, produces a design that is more easily
managed over time.
Integrability
relates to the degree to which a system’s interfaces are open, or flexible. Compatibility between elements is a critical
enabler of flexibility. As an example,
consider a docking interface on the space station. This interface would ideally be common across
all future spacecraft, ensuring full compatibility. As an additional example outside the
aerospace industry, consider the USB interface standard now used by many
electronic peripheral devices such as keyboards, computer mice, flash memory
cards, etc. The use of dedicated
interfaces for each one of these devices would be highly inefficient,
especially given the fact that only a small subset of the devices is needed at
any one time.
The concept of
extensibility is critical to the creation of a sustainable exploration
system. Extensibility must be an
integral part of the exploration strategy to ensure that forward progress
serves as a continually growing exploration foundation, even in light of policy
direction changes. The concepts of
extensibility are woven into the baseline missions and example conceptual
designs within this report.
The history of
Antarctic exploration provides valuable lessons for space system designers.
From its inception Antarctic exploration and science shared many attributes and
constraints with current space activities. Both, for example, have been highly
dependant upon technological advances, including the need for complex logistics
and cutting-edge life-support capabilities. Months of isolation during
Antarctic expeditions present psychological hardships similar to those
anticipated in extended Moon and Mars missions. More generally, Antarctic
exploration, like space activities, has brought science into close involvement
with politics. The following section
examines how these factors affected some aspects of the development of
Antarctic exploration and science, and draws lessons for space exploration
programs.
“More than any other, Antarctic science is
dependant on logistics, on the ability to place and maintain a scientist and
his equipment in the right place at the right time. Expeditions to Antarctica
up to 1925 depended on techniques of transport, communication, survival, which
remained largely unchanged for 100 years…. after 1925 the development of
mechanized transport, the airplane, radio and technology based on better
understanding of human physiology, were to make access to the Antarctic, travel
within it and survival in its hostile environment, much less difficult.” (Beck
1986, p.131).[ESS2]
The above quote
summarizes well the disjointed nature of Antarctic exploration. Rather than
follow a steady, continuous path of progress, the pace of discovery on the
continent advanced through steps and jumps. Importantly, these advances in
capability often resulted from the congruence multiple technologies, rather
than any single technical development. Each jump offered great advances in
knowledge returned per expedition, a situation that should be anticipated and
exploited in the design of space exploration systems.
Most significant of
these advances involves a shift from what has been termed the “Heroic” age to
the Modern age of Antarctic exploration. The Heroic age is roughly delineated
as the period from 1895 to the dawn of the First World War in 1915 (Walton,
1987). It marked a dramatic shift in capability from the previous era because
of the use of liquid fuel, however, due to the still rather primitive methods
of transport and “life support,” expeditions during this period often brought
extreme hardships. National prestige, sovereignty, and personal fame—not
science—motivated exploration during this period.
The Modern age
begins roughly with the American expedition lead by Richard Evelyn Byrd from
1928-1930. It is characterized by the comprehensive use of airplane travel,
electric communication, mechanized transport, and thus continuous logistical
support (Fogg, 1992). Most of these technologies had existed for some time, and
had been tested and refined through previous expeditions. Byrd’s expedition was
the first to coordinate them systematically, increasing the amount of data
collected by orders of magnitude. The following table summarizes the major
technical advances that enabled this shift, as well as the impact on
exploration capability and knowledge return. Systematic use is defined as use
in everyday operations, as opposed to sporadic use and testing.
Technology |
Introduction for
Exploration |
Systematic Use |
Mission/Logistics Impact |
Initial Knowledge Return
Impact |
Space-based equivalent |
Radio Communication |
1911 |
1929 (Byrd) |
Coordination, safety |
Immediate news of success
increased public interest |
Satellite Communications |
Combustion Engine (land
travel) |
1907 (Shackleton) |
1933 (Byrd) |
Outdoor activity and
travel in harsher conditions |
Distribution of heavy
seismic equipment |
Rover |
Airplane |
1929 |
1928 (Byrd) |
Pre-positioning for
extended expeditions; Aerial photography |
1 field season of
land-based observation per hour (4000 square miles) |
UAV's, Pre-positioning
technology |
Ice Breakers |
--- |
post-WWII |
increased access,
extended access |
More feasible permanent
base |
cyclers |
Implications can be
drawn from these examples for space exploration. Advances fall into rough
classes of technologies with analogues in space systems. Combustion engines,
which enabled the equivalent of surface rovers, had a great impact on the kinds
of fieldwork that could be executed. Their introduction created the possibility
of distributed use of heavy equipment for seismic operations. Their impact on
mission logistics, however, was minimal at first.
The airplane and
the radio had dramatic affects on knowledge return and mission logistics. The Byrd expedition was the first to fly over
the pole. In doing so, he took over 1600 pictures covering 150,000 square miles,
or the equivalent of 37.5 field seasons worth of observations using previous
methods (Walton, 1984). He also discovered two Mountain ranges. The airplane
also allowed for the possibility of pre-positioning and logistical support for
inland bases.
Soon after flying
over the pole, Byrd was able to communicate the accomplishment. His successful
flight was beamed via radio immediately back to the
An interesting
feature of the progression of technological development is the lag between
testing and systematic use. Radio communication and the combustion engine were
tested with little impact in many expeditions before the Byrd expedition.
Interestingly, life
support capabilities advance much more gradually than logistics technology. Man
learned to live the extreme environment gradually, over several hundred years,
with advances coming more through trial and error than scientific or
technological breakthrough. (Fogg, 1992)
In many ways, NASA’s
current task is to transition space activities from a heroic to a modern age.
While national prestige and public attention will continue to play important
roles in space activities, the time has come for more systematic and knowledge
return. The history of Antarctic exploration demonstrates that when this
occurs, as in the case with the first Byrd expedition, public attention and
government funding are likely to increase rather that decrease. The next
section examines this dynamic of science and politics.
Antarctic
exploration requires support at the national level. Thus, as one author notes,
“Antarctic scientists have often been used as political instruments and it
would be unrealistic for them to think that their work can be isolated from the
spheres of interest of economics, law, and [ESS3]politics.”(Klein 2000, p.319) The motivations behind various stages of
Antarctic exploration are extraordinary in their similarity to space
activities. They include included: prestige of geographical discovery,
information and experience for navigation and commerce, and sovereignty. While
science always played an important role during expeditions, and is now the
single most important product of exploration, it is important to note that the
underlying motivation for countries to invest in Antarctic travel has almost
always been the “maximization of influence” rather than knowledge (Lee,
personal communication).
Territorial issues
became increasingly important at the transition from the Heroic to Modern age
of Antarctic exploration. From 1908 until the signing of the Antarctic treaty
in 1961, international tension rose and fell as countries made varied and
conflicting claims to sovereignty. The following events in particular were
important to this dynamic.
1908 and again in 1907
1923 British claim the
Roth Dependency
1924 French claim Adelie
land
1933
~1939
While the
motivations behind these claims were complex and interrelated, the World Wars
and advances in technological capability were certainly central factors. As
with space activities during the Apollo Era, international interest, enabled by
technological advances, fueled funding for exploration.
Byrd’s expeditions
are a particularly interesting example of this kind of feedback loop in the
“The most important thing is to prove (a) that
human beings can permanently occupy a portion of Continent winter and summer;
(b) that it is well worth a small annual appropriation to maintain such
permanent bases because of their growing value for four purposes—national
defense of the Western Hemisphere, radio, meteorology, and minerals. Each of
these is of approximately equal importance as far as we know.” (Fogg, 1984,
p.162)
Following the
Second World War, international interest in
“Because of its position of leadership in the Free
World, it is evident that the United States could not now withdraw from the Antarctic…national
prestige has been committed…. our capacity for sustaining and leading an
international endeavor there that will benefit all mankind is being watched not
only by those nations with us in the Antarctic but also by noncommitted nations
everywhere. Antarctic simply cannot be separated from the global matrix.
Science is the shield behind which these activities are carried out.” (Beck, 1986 p. 64)
While this view is
a product of the geopolitical context, it illustrates how various factors can coalesce
to form a sustainable program from a political perspective. The Byrd
expeditions from before WWII had demonstrated American technical superiority in
exploration and proven that modern technologies could be used to improve access
to the continent. After the war, politicians and diplomats began to view
exploration as an important symbol for global cooperation and competition, and
were committed to continuing operations.
Once implicated, national prestige and technical capability became
intermingled, heightening the perception of value of continuing exploration.
Conclusions – Exploration and Sustainability
An important lesson
that the history of Antarctic exploration provides for space exploration system
designers involves the interplay between news, knowledge, technology, and
funding. While Arctic exploration progressed slowly for decades, it was marked
by distinct stages of increasing capability and increased interest. As the Byrd
expedition illustrates, quite often advances in logistical and knowledge
acquisition and transfer capability translate to increased political interest
and funding. The spread of news creates public interest, while increased
knowledge and logistical capability creates military interest. Both can
generate funding for further expeditions, thus creating a positive feedback
loop of discovery and technological development. Figure
6 illustrates the salient aspects of the
feedback loop, which enabled the Byrd expeditions.
Figure 6: Positive feedback loop for exploration
Of course the real
dynamics behind such a process are complex and varied. Byrd’s expedition
occurred at a time when international interest in
MIT’s 2004 spring
class in Space Systems Design investigated the design of extensible space
system architectures. A central difficulty in this task was the shear complexity
of the problem, and the lack of an established methodology to design system
architectures. An important result of the investigations was thus the methods
developed to approach the problem, and the process by which “sustainability”
could become central to design decisions. The end result was an iterative and
holistic approach to the problem, which will hopefully inform future space
systems architecture projects.
It should be
stressed that not every aspect of the process described was completed rigorously
during the semester. Rather, the process represents a way to integrate the
lessons learned and eventually create a systematic architectural design. Of
course every element of this process did not proceed in clear and neat steps.
Most of the steps were iterative within themselves, and individual elements
were re-worked as
The underlying goal
of the design process was to develop an integrated strategy that could quantify
how the system reacted to changes in the environment. Rather than create a
point design to accomplish a Moon or Mars expedition, the class wanted to
demonstrate that various scenarios could be anticipated and addressed during
conceptual design and, as importantly, that the elements designed to address
these scenarios (which would likely make the system sub-optimal from a
point-design perspective) could be justified quantitatively. A strategy
includes various scales of Moon and Mars missions, robotic scout missions, and
considers the program changes such as budget cuts and regulatory constraints.
Figure
7
illustrates the five step process arrived at to create the strategy. An
important goal was the establishment of common operations and across manned
Moon, Mars and potentially asteroid missions, as well as through stages of
missions at each body. Common elements defined baseline architecture forms and
operations, from which options could be created to address specific missions
and changing scenarios.
Figure 7: Space systems design process
The first three
steps in the process identify common forms and functions needed to explore the
Moon, Mars and other destinations. Two teams conceived of staged Moon and Mars
missions, and created matrices with functional requirements for each stage.
With these functional requirements, a simple Venn diagram captures the
relationship of requirements between the Moon and Mars. An interesting feature
of this part of the process involves the ability to identify how formal
elements can be extracted from functional requirements based on commonality
between Moon and Mars needs at various levels. “Options” can be created to
supplement the core needs, based on requirements outside of the intersection of
the circles.
Functional
Commonality Mapping thus revises the forms created to enable various Missions.
The two teams must then return to the mission storylines and establish how and
whether mission objectives can still be met with the revised forms, and alter
staged missions accordingly. This iterative process can continue until a
satisfactory level of refinement is achieved.
It was found that
this iterative part of the process reveals key trades that need to be made with
respect to commonality and architecture operations. Based on our designs, trades on issues such
lander design, rover design, aerobraking capability, and operational capability
processes such as the use of the Earth-Moon Lagrangian points, could not be
solved by commonality mapping alone. The next step of the process is thus to
evaluate the key trades revealed by the first three steps of the process.
In order to create
a flexible strategy, however, it was important to evaluate these trades with
consideration for the value of flexibility and robustness, not just optimality.
Tool such as real-options, multi-attribute utility theory, and decision
analysis, can be used to carry out the trades while preserving system
flexibility, thus creating a rigorous development strategy and architecture.
Chapter 6 addresses
how these tools can be used to evaluate strategic and technical options. The
strategy includes staged deployment of Moon and Mars missions, with development
options forming branches from the baseline mission. Ideally the aspects of the
system designed early in the strategy will minimize the need for redesign if
new directions in the strategy are taken.
As noted, the full
strategy was not generated during this design course. Instead, various aspects
of the process were addressed and tools were conceived to facilitate their
design in later studies.
An extensible space exploration infrastructure may be modeled as a mass transportation system, but also as a knowledge delivery system, since mankind is sending robotic and human explorers to space for the purpose of exploring and returning knowledge about the Moon, Mars and Beyond.
To justify knowledge as the deliverable to the stakeholders one must investigate why knowledge is the deliverable and who the stakeholders are. To answer the first question, one must first understand why do humans explore. To summarize, the three main reasons are
Knowledge is the product of the exploration process. The knowledge of our surroundings is closely tied to science. Technological leadership is knowledge delivered to the technologist and explorers. The third point is that inspiration in science and technology is the knowledge delivered to public and commercial enterprises. In other words, the knowledge gained by the space exploration system is the value-added delivery to the beneficiaries or stakeholders. Therefore, to ensure the maximum value delivery, one may model the space infrastructure as a knowledge delivery system. Knowledge returned may be categorized as scientific knowledge, resource related knowledge, technical knowledge, and planning related knowledge. To build up the argument, first one must understand the value delivery to the scientists, which is diagramed in Figure 8. To understand the value identification, the goal of the space infrastructure is to increase the quantity and depth of scientific knowledge of the solar system by sustainably and successfully exploring the solar system, specifically the Earth, Moon, Mars, and Asteroids (EMMA) using an affordable and extensible human and robotic exploration system for the immediate benefit of the scientific community.
Figure 8: Value delivery to scientists diagram
The value delivered to the technologists and explorers is an increase in the quantity and depth of resource and planning related knowledge of the solar system by sustainably and successfully exploring the solar system, specifically the EMMA using an affordable and extensible human and robotic exploration system, and the previously gained resource. The value delivery can also be seen in Figure 9.
Figure 9: Value delivery to technologist/explorers diagram
In addition to the scientists and
technologists/explorers, knowledge may be returned for the benefit of the
Figure 10: Knowledge delivery system OPM (
There are five main types of knowledge: Scientific-, Resource-, Technical-, Operational-, and Experience-related knowledge as seen in Figure 11.
Figure 11: Five types of knowledge
Scientific knowledge can be generalized as the search for the existence of life and Planetary E3 (the characterization of Evolution, Environment, and Existability of a planet or any celestial body). The existence of past or present life drives the search for resources such as water and other biomarkers. Evolution is mainly concerned with understanding the geology of a planet while Environment is the climate characterization. Existability is an assessment of biological potential, or how benign or hostile a planet is to human settlement.
One way to quantify scientific knowledge is through keeping track of the number of scientific publications resulting from the exploration effort. This is “an accepted measure of scientific productivity” and can be easily tracked using databases such as the NASA Astrophysics Data System (ADS) (Green, 2004). An example of this is seen in Figure 12, which captures the number of papers published as a function of the publication year for the Hubble Space Telescope. Using a numeric quantity, such as the number of publications, it is possible to make comparisons between different exploratory missions. It is then possible to understand when a diminishing amount of knowledge is returned and when it may be beneficial to gracefully retire an exploration mission. For example, if as in Figure 12, the number of papers per year were to steadily decrease for several consecutive years, the exploratory phase of the mission would be approaching its end. The final result of the knowledge publication graph might resemble a Gaussian distribution, where a mission is retired after it reaches a certain point in the distribution. Other possibilities for measuring scientific knowledge include news articles, press releases, website hits, educational television programs, PhD dissertations, or proposals.
Figure 12: Example of the quantity scientific knowledge from Hubble (Beckwith, 2003)
Resource knowledge is defined by the existence, location, and amount of planetary resources that can be utilized by human explorers. These indigenous resources are necessary to build and maintain an extensible space infrastructure. Possible indigenous resources include water, Oxygen, Hydrogen, Ores/major metals, Nitrogen, and energy sources such as fusion materials. These resources may be obtained using the following three-step strategy:
1. Existence. The first step to is to determine the existence of the resource, most likely using robotic explorers such as orbiters. First, implied existence of the resource is obtained by knowledge carriers, which transmit passive bits. The next step is to obtain direct proof of the resource’s existence, either by transmitting bits or by transporting atoms.
2. Location and Amount. The second step is to determine the global amount of a resource, possibly using an orbiter or rover. An unmanned rover is beneficial for reconnaissance of biohazardous and toxic regions. As the resolution of resource knowledge about the specific resource locations and amounts increases with exploration, a point is reached when a human mission may begin to extract the resource. This point would occur when the resource location accuracy at least meets the landing accuracy plus the travel distance of a human mission.
3.
In-situ Utilization. The final
step is to begin in-situ resource utilization for exploration needs, such as
propellant, building materials, and energy.
A lander can achieve basic in-situ knowledge, but full exploitation will
likely occur with a human mission. Some
of the issues with in-situ utilization are related to the degree of
manipulation needed. For example
possible water ice on the Moon may need a heating, purification, and extraction
process before it is useable.
Technical knowledge is the assessment of the engineering abilities associated with the space transportation system similar to the NASA Technology Readiness Levels (TRL). The space transportation system will slowly attempt to integrate various new technologies into the existing infrastructure. The level of working ability for each technology is the technical knowledge delivered. An example is the development of in-situ resource technology, where currently designs exist at various conceptual levels. As the system is developed, in-situ resources can be utilized. The degree of success delivered, measured in cycle efficiency, total power consumption, and resource produced by the in-situ technology is the technical knowledge. Technical knowledge gained will affect the evolution of the space transportation system. It will help determine how missions grow, which will be discussed in later sections.
Operational knowledge is the capability of performing activities related to the space transportation system. An example of operational knowledge obtained during the Apollo program is lunar orbit rendezvous, or docking. The technology for docking existed and the procedure for it was known, but not until it was successfully accomplished was there a large amount of operational knowledge gained concerning docking. Other examples include operational knowledge gained from Lagrange point maneuvers, pre-positioning, drilling in low gravity environments, and long duration human factors issues. An interesting point about operational knowledge is that unlike the previous types of knowledge, a good deal can be gained from failures. For example, during Apollo 13, operational knowledge was gained when the Lunar Module was used as a “life boat” and various components were also creatively utilized to ensure crew survivability. It is uses of a system beyond their intended designs, which can lead to operational knowledge. Therefore operational knowledge can be gained by understanding the flexibility of a system.
The human experience can also be a type of knowledge, because there is a unique gain that is achieved outside of data or physical returns. It is may be thought of as a combination of the four other types of knowledge. A human presence can gain knowledge that is different from any robotic explorer or remote sensor due to its rapid cognitive thinking and senses. This idea is very similar to the notion of experience as a knowledge carrier, which is outlined in Section 2.3.3.
Carriers of knowledge are divided into three main categories, bits, atoms, and human experience.
Bits carry knowledge in the form of the data. There are two types of bits, passive bits and active bits. Passive bits are defined by data obtained without interacting with the observed environment, such as taking a picture. Active bits involve interacting with the environment such as by taking a measurement and transmitting the measurement data back.
Atoms are the physical samples that carry knowledge about an exploration excursion. These samples carry two forms of knowledge: implied discoveries and direct proof discoveries. An implied discovery is knowledge that is gained by observation or measurement of a sample, which leads to an implicit discovery; for example, a weathered rock exhibiting the past existence of water by erosion patterns. A direct proof discovery is the knowledge carried by hard evidence of a phenomenon, for example, a Mars rock with a pocket of water carries proof of Martian water by direct observation of the specimen.
The human experience of exploration has the ability to carry the greatest amount of knowledge. While robotic explorers could be the prime means of bringing back bits and atoms, they are most effective for large amounts of systematic returns. In contrast, there are three human physiological traits that provide an optimal combination for returning knowledge:
1. The human brain. Capable of instantaneous programming, the human brain is a “qualitative supercomputer” (Schmitt, personal communication). It can react to field experience and training and adds a high degree of flexibility
2. Eyes. The human eyes have high mobility, dynamic range, and quick three-dimensional integration, especially in the 10 – 15 meter range (Schmitt, personal communication).
3. Hands. Perhaps the most underutilized human tool, but if their dexterity can be used to their full potential they can greatly increase the human exploration ability. For example, hands posses the capability of returning detailed tactile feedback, etc.
An example of the benefit of the human experience can be seen in the NASA Opportunity Rover on Mars. Throughout its mission, it has returned knowledge by observation and interaction with the environment (bits), but it has sent back even more questions about Mars. These questions could have been answered immediately by a human field geologist present on Mars, due to his/her unique ability to analyze the environment with his/her experience, physiological tools, and basic scientific instruments, such as a hammer (Schmitt, personal communication).
Robotic and human explorers have different degrees of time and spatial processing abilities, as seen in Figure 13. Time processing ability is meant by how an explorer is able to take in and understand the value of interesting exploration targets. Spatial processing is defined as the ability to understand the value of exploration targets in a global resolution and also a high-resolution sense. Figure 13 shows three types of robotic explorers, penetrators, orbiters, and rovers. Penetrators are geologic instruments that are embedded in the ground and have no mobility. An example of a penetrator is the Deep Space 2 probes. Penetrators cannot move and only have spatial resolution of their immediate surrounding, and rely on their instruments to record data as it comes to them. They passively gather data and have limited range to interact with the environment and collect additional data. In addition, penetrators are not reprogrammable (yet), once they land, they execute their specified tasks. Therefore they are shown to have low time and spatial processing abilities. Orbiters have a large global resolution, however they are unable to achieve high resolution of specific targets (yet) or look at a target from multiple unique angles. An example of an orbiter’s limitation is that, it would have a difficult time looking inside a cavern. Rovers are shown with greater spatial processing ability because they are able to look at targets from multiple viewpoints and with a high resolution. They cannot achieve the global scale resolution of an orbiter, however with increased mobility and presence rovers can attempt to create a larger global picture with high resolution. Rovers are also shown with higher time processing ability because they can be flexible to their environment. They can take a picture of their surroundings, and then be commanded to move to locations they seem the most interesting. In contrast, an orbiter can only explore targets that are in its orbit’s coverage region. Finally, the human field geologist equipped with a rover and tools such as a microscope has the highest amount of time processing ability due to his training, experience, and brain. Equipping the human explorer with high mobility (rovers) and microscopes, will give him the ability to have global resolution and high specific resolution of targets. Therefore, a human explorer is the optimal combination of time processing ability and spatial processing ability. This forms an argument for exploration by humans in place of robotic systems.
Figure 13: Time and spatial synergy for robotic and human explorers
Figure
14 illustrates a summary of
the knowledge carriers and how they are related by their degree of interaction
with the environment, and the quantity of that specific knowledge carrier that
mankind has accumulated. Passive bits
are represented by pictures of planets and the galaxy and currently carry the most
knowledge. In decreasing amounts of
quantity are active bits represented by graphs of Mars Seismic activity from
Voyager, followed by pictures of Mars rocks from the
Figure 14: Carriers of knowledge
One challenge for the knowledge delivery system is to understand the difference between knowledge and news. To first order, news is the unique knowledge on a generalized subject. For example, the discovery of an extrasolar planet is news; however, discovery of the nth extrasolar planet is not news to the public. News is the knowledge that immediately appeals to the public. A notional graph of news versus exploration milestones can be seen in Figure 15. Shown are theoretical news values for Apollo and future milestones. The diagram shows the notion that a new milestone, such as the first Apollo mission will have a high news value, but there is a decay in news as the Apollo missions progress, shown by the decaying black line. If there is a new unique milestone, such as a human Moon return or a 1st Mars Human Landing, it is possible that there will be a large increase in the news they generate. As with Apollo, these events will be followed by a decrease in news value, since the 2nd and 3rd human Moon return and 2nd and 3rd Mars human landing will not be new milestones. The notional diagram exhibits this high frequency decay. Overall, there is a low frequency decay in news value of the entire exploration system. Thus if the media does follow this trend, it is unfavorable for sustainability.
Figure 15: Theoretical news value as the space exploration system evolves
An important distinction may be drawn between public interest and media interest. When the media loses interest in a subject, the public tends to lose awareness of it. The media interest is also one of the main processes of continuing education that the knowledge delivery system uses. For the system to be most effective, the public interest could be coupled with education, but decoupled from the media. A gradual increase in public interest is necessary to create a knowledge distribution system that is independent of media. A challenge to this solution is that many politicians, who advocate for funding, tend to have some of the strongest associations with the media. A breakthrough in separating public interest from media interest can occur when personal connections with the space transportation architecture are developed. For example, when settlements, be they permanent or semi-permanent, exist outside of the Earth, many people on Earth will have personal connections with those on the Moon or Mars generating interest that is independent of the media. A breakthrough can occur when there is commercial interest in the Moon and Mars.
It is important that the knowledge delivery system does not rely too heavily on the media. The media loves success, the first time, but in general it looks for disaster (Schmitt, personal communication). A good example may be found in the Apollo missions. The media coverage of Apollo 8 and 11 was huge, since these missions achieved historic firsts. Coverage was also large for Apollo 13 because of its challenges, and then Apollo 14 since it was the first after a disaster. However the later Apollo missions did not experience such significant media coverage.
The knowledge delivery process can be summarized by the CDIO phrase/process with an added S at the end. These letters stand for Conceive, Design, Implement, Operate, and Science. During the Conceive stage the mission goals, requirements and trades are identified, then during the Design stage the goals and requirements are used to create an comprehensive design of the mission and all the elements that are required to make it successful. The Implementation stage consists of the building of the elements designed in the previous stage. During the operate stage the mission will collect the data that will eventually be turned into knowledge during the Science stage. Of course there is some overlap in the stages, but for the most part each stage acts as its own step in the knowledge delivery process. Each mission in an extensible exploration system must follow this pattern, which should begin to repeat at about the time that previous mission has reached the O stage. The estimated relative times for each stage for robotic, human Mars and Moon missions are given in Table 1.
Table 1: Knowledge delivery process
Knowledge process Timeline |
||||||||
|
|
|
|
C |
D |
I |
O |
S |
|
C |
D |
I |
O |
S |
|
|
|
|
|
|
|
|
|
|
|
|
Robotic Missions |
1x |
3x |
1x |
5x |
nx |
|
|
|
|
|
|
|
|
|
|
|
|
Human Moon |
1x |
3-4x |
1x |
0.1-0.5x |
mx |
where m<<n |
|
|
|
|
(launches) |
|
|
|
|
|
|
Human Mars |
1x |
4-5x |
1x |
3x (0.1x) |
mx |
|
|
|
In the above Table 1, n stands for a constant amount of time that should be between six months and one year. The amount of time that it takes to evaluate and handle raw science data is less on a human mission than on a robotic one. After the first round of missions, the relative times should change slightly, especially if there is any form of reusability added into the system.
The Knowledge Delivery process is summarized in Figure 16, which shows the overall knowledge delivery cycle and introduces the concept of Knowledge Delivery Time (KDT). There is some time difference between the beginning exploration phase and the knowledge delivery. The primary mission can directly lead to mission delivery shown by the dotted line in Figure 16, or knowledge processing on Earth, after the primary mission on Earth can then lead to knowledge delivery. After knowledge is delivered, the exploration cycle begins again with another conceive, design, Implement, and operate processes.
Figure 16: Knowledge delivery cycle
Knowledge delivery time is defined as the time between collection of knowledge and its delivery. It is usually not instantaneous. Two case scenarios, illustrated in Figure 17, can be used to understand the concept of knowledge delivery time.
Figure 17: Knowledge delivery time examples
For Mars Global Surveyor (MGS), there is a
time difference between orbital insertion, which represents when exploration
and the collection of knowledge begins, to the delivery of a significant amount
of knowledge, in the form of a journal paper discussing the existence of water
on Mars.[a5] Pictures from Mars from MGS, which are
passive bits knowledge carriers, were used for the discovery (Malin,
2000). The KDT from initial exploration
to knowledge return was approximately 35 months. The next exploration of Mars with the
objective of determining recent Mars water was the Spirit and
The following examples of KDT come from robotic missions. Therefore, future robotic explorers could have similar knowledge delivery times. The benefit of human exploration is that it has the ability to decrease and even eliminate the knowledge delivery time. The human exploration experience can process and interact with the environment rapidly and return knowledge with minimal delay. For example, the Apollo explorers could immediately determine that lunar regolith (to first order) was mainly composed of inert dust and rock fragments verifying knowledge from photos taken over many years.
Different aspects of the mission such as crew size, experience, excursion time, exploration time, mobility, range, and instrumentation affect knowledge. All of these with the exception of instrumentation will be modeled. The reason instrumentation is not modeled here is that it varies by mission depending on the specific science objectives of that mission. For example, a mission focused on geology will have very different instruments than a mission focused on climatology. Instrumentation is further discussed in Appendix 9.5. During the later Apollo missions approximately one third of the total time spent on the Moon surface was spent during an excursion (Table 2). The earlier Apollo missions did not have as high of an excursion time for two possible reasons. The first is the lack of experience. Later Apollo missions were able to gain experience with surface operations on both the Earth and Moon from the first Apollo mission. The other reason is because the first few Apollo missions did not have science as their primary mission objective. Only Apollo 15-17 had “extensive scientific investigation” of Moon as a primary mission purpose whereas Apollo 11’s primary mission was a manned lunar landing demonstration (NASA website, 2004). The primary mission for Apollo 12-14 was precision piloted landing and systematic lunar exploration. Thus, experience had an effect on the excursion time for lunar missions, and had a maximum of 30% of the total lunar stay time for a mission.
Table 2: Apollo mission details (NASA website, 2004)
|
kg |
duration (hrs) |
outside LM[min] |
max d from LM (m) |
%outside |
Apollo 11 |
21.6 |
22 |
152 |
61 |
11.5 |
Apollo 12 |
34.3 |
31 |
465 |
411 |
25 |
Apollo 14 |
42.3 |
33 |
563 |
1454 |
28.4 |
Apollo 15 |
77.3 |
67 |
1115 |
5020 |
27.7 |
Apollo 16 |
95.7 |
71 |
1214 |
4600 |
28.5 |
Apollo 17 |
110.5 |
75 |
1324 |
7629 |
29.4 |
Each excursion had a predetermined plan, however there were times when independent exploration was allowed and carried out by the astronauts. An example of independent exploration results was the orange pyroclastic glass discovered in Shorty Crater by Apollo 17. This exploration was not dictated by ground. When the astronauts were exploring in the area, they had 30 minutes of rapid assessment and gathering before the mission controllers were even aware of the events. Independent exploration allows a human to fully utilize his/her training, experience and senses to return knowledge, either as samples, pictures, observations, technology used, or operational procedures.
Using the number of crew, excursion time, exploration time, and mobility, the coverage area during exploration can be determined. Assuming an experienced crew, the excursion time can be maximized as 30% of the mission surface time, which is similar to the Apollo missions. Exploration is defined by examining the surrounding area within 10 meters of the astronaut, since this is the range for which the human eye has an optimal ability to determine unique aspects of the surroundings (Schmitt, personal communications). It is assumed that some portion of an excursion is spent performing this exploration process. For this knowledge model, 30% of the excursion time was spent as this independent exploration time. This could vary a great deal on an excursion by excursion basis, but it was approximated based on personal communications with Jack Schmitt. There are three types of mobility, a walking pace (or gait) while traversing without exploring, a slower exploration pace, and a rover speed. The walking pace is based on the design speed of an Apollo astronaut and determines the maximum traveling distance per day from the starting point, presumably a lunar module . The exploration pace is estimated as four times slower than the gait because the human is more carefully analyzing the environment and perhaps taking measurements or pictures (NASA Headquarters website, 2004). The rover pace is based on the Apollo Lunar Rover (Apollomaniacs, 2004) and is capable of expanding the maximum traveling distance per day. These parameters are summarized in Table 3 and are used to determine how much coverage per day can be accomplished.
Coverage is defined as the area traversed at an exploration pace and can be used to quantify how much knowledge potential is gained. As more area is explored, a greater amount of knowledge is potentially gained, either from science or resource data, or technical and operational procedures. Since exploration pace is slower than the maximum speed by astronauts, either by rover or by walking, there is a certain amount of coverage than can be achieved per day. The number of crew available will directly affect this coverage, which is shown in Figure 18. There is a clear direct relationship between the number of crew and the maximum exploration coverage achievable per day. The coverage achieved while walking can either be completed by increasing the number of days on the surface, or by increasing the number of crew. If walking is the fastest mode of transportation, 100% coverage can be achieved rather easily, after which the knowledge potential is maximized in that particular landing site. Using a rover increases mobility, and in this case, the maximum area that may be traversed in a day dramatically increases, decreasing the percent area covered per day. To increase the exploration coverage requires either a longer stay than the non-rover case or a larger crew.
Table 3: Knowledge drivers model parameters
outside LM
fraction |
0.3 |
exploration
time |
0.3 |
traveling pace
(gait)(m/hr) |
3600 |
exploration
pace(m/hr) |
900 |
rover pace
(m/hr) |
14000 |
Figure 18: Knowledge potential: maximum exploration coverage per day versus number of crew
This model can also incorporate faster speed rovers. Introduction of a pressurized rover that allows astronauts to stay outside of the base can further expand knowledge potential as seen in Figure 19. By creating a remote base, travel can be further expanded from the remote location, thus expanding the knowledge potential from exploration.
Figure 19: Expanding the exploration potential using a remote base (Hoffman, 1998)
Another possible method of increasing knowledge potential is to increase the amount of independent exploration time per excursion. In fact, as missions proceed, a greater amount of independent and flexible time for the astronauts should be encouraged, especially as communication delays increase for farther missions (Schmitt, personal communication). In conclusion, the knowledge potential of a mission can be predicted by the exploration coverage, which is affected by the number of crew, experience, excursion time, exploration time, and mobility.
The Apollo missions offer a good case study for how knowledge return is affected by knowledge drivers. In this case, knowledge returned is quantified by the mass of samples returned from the Moon. If Moon rocks are carefully chosen, increasing amounts of samples gathered should return increasing amounts of knowledge (to a first order), most likely scientific or resource related. The amount of knowledge can also be driven by the exploration time and the distance traveled. Figure 20 illustrates the returned mass as a function of the maximum distance from the Lunar Module and as a function of the time spent outside of the Lunar Module (data from NASA website, 2004). Clearly there is a direct relationship, however there are other inherent drivers that are not as easily captured. They are human experience, timing, and new technology. Apollo 11 lacked experience and thus had a limited exploration time and did not traverse much distance, as shown by the point closest to the origin in each plot. As the missions progressed in time, experience increased, resulting in longer and further exploration. Each of the graphs illustrate a large jump in the mass of samples, and therefore the knowledge, returned, coinciding with Apollo 14 to Apollo 15 because Apollo 15 was the first mission to include a rover. Figure 21 illustrates a graph of the cost of the knowledge returned and the table showing the percent increase of cost and mass returned of one Apollo mission relative to the previous. It is important to note that the cost shown here does not include development costs. The largest percent increase in knowledge coincides with the largest percent increase in cost, which occurred in Apollo 15 because of the introduction of the rover. This is a clear example of how infusing new technology into an existing architecture can result in an increase of knowledge returned. Apollo 16 and 17 continued using the rover and experienced a much higher sample return than the non-rover missions. With added human experience from previous rover-enable missions, they were also able to traverse further and explore for longer periods of time.
Figure 20: Apollo knowledge drivers
Figure 21: Apollo cost trends
The design of a sustainable space infrastructure will use knowledge as the deliverable as a metric for its design. The next sections will discuss how different missions gain various types and levels of knowledge. Knowledge will affect mission characteristics such as landing locations, surface mobility, and mission lifetime. The knowledge model developed in this section was not directly used in the mass transportation architecture, but it would be beneficial for future research to incorporate the knowledge drivers into the mass architecture in a more integrated fashion.
Having identified knowledge as the key value-added deliverable, and examined how kind and quality of knowledge vary with mission profile and technology, the problem remains how to integrate this understanding into a space system conceptual design that takes into account sustainability. This chapter describes the first step in the process described in chapter two to create sustainable exploration systems: designing staged missions to the Moon and Mars. Staged missions are defined as Short, Medium, and Extended Stay, and were designed with the goal of maximizing commonality through each stage. Later steps in the design process, described in chapters five and six, help determine functional and formal commonality across Moon and Mars missions.
It is important to note that once commonality across Moon and Mars missions has been identified and exploited through changing mission forms and operations, the individual staged Moon and Mars mission must be revisited and altered according to new capabilities and requirements. Steps one through three of the design process are thus iterative. This chapter represents the culmination of one iteration of this process, and thus presents considerable formal and operational commonality between staged Moon and staged Mars missions. For this reason, it begins with a discussion of the forms that will be used for both the Moon and Mars staged missions.
The Crew
Operations Vehicle (COV) is functionally similar to the Apollo Command
Module, capable of transporting a crew of three and supporting the crew for a
short duration mission. The Habitation Module (HM) is an extensible
habitable volume, made up of multiple modular sections. The Habitation Module can sustain life for
long duration missions. When COV and HM
modules dock, they form the Crew Exploration System (CES). The Service Module (SM) is capable of
providing propulsion for transiting the crew from Earth to destination or
destination to Earth. Service Module #1
is the engine for the trip to the destination while Service Module #2 is the
engine for the return trip. In
combination with the COV and HM, this module is defined as the Moon/Mars
Transfer Vehicle (MTV). The Mars
Landers (ML) or the Lunar Landers (LL) are functionally similar to
the Apollo type Lander, although have slightly different forms for Moon and
Mars missions, and capable of transporting three crewmembers from orbit to the
surface and back into orbit. Capable of
providing accommodations for three crew members for launch into LEO and descent
back to Earth, the Modern Command Module (MCM) is
functionally similar to the COV. Two
MCMs are needed for missions with crew sizes of six. These modules are summarized in Table 4.
Table 4: Architectural space transportation forms
The
lunar baseline mission includes Short, Medium, and Extended missions. The
objective of Short Stay Lunar Missions (SSLM) is to demonstrate the basic
technology for lunar missions. The SSLM
is equivalent to the first “return to the Moon” mission described in President
Bush’s New Vision for Space Exploration Program. Approximately two SSLMs are suggested, with a
frequency of two per year, as dictated by Earth launch considerations aimed at
maximizing launch cost efficiency. Refer
to Section 6.4.2 for more information about launch considerations.
The objectives of the Medium Stay Lunar
Missions (MSLM) are to acquire scientific knowledge and assess the value of
potential locations for Extended Stay Lunar Missions (ESLM). Approximately five
MSLMs are suggested with at least one landing on the far side of the Moon or a
lunar pole. A frequency of two MSLMs a year is suggested based on Earth launch
considerations. The MSLM is based on the
“stepping stone” approach mentioned in President Bush’s New Vision for Space
Exploration Program to steadily increase our mission complexity while expanding
our reach out into the solar system.
Finally, the Extended Stay Lunar Missions
(ESLM) will involve a six-month surface stay and will include a semi-permanent
base. This is a continuation of this
“stepping stone” approach. These
missions are to serve as a testbed for future Mars missions and provide a
platform for long-term lunar-based science investigations. Approximately two
ESLMs are suggested on the far side of the Moon and/or a lunar pole to carried
out at a frequency of approximately 1.5 missions per year.
Although the baseline mission architectures
presented in this paper are not optimized, previous lunar mission architecture
studies were reviewed to inform architectural decisions for the lunar baseline
mission presented in Section 4.2.4.
Houbolt (1961) extensively studies many
combinations of mission architectures, concluding that Lunar Orbit Rendezvous
(LOR) is the fastest and most reliable architecture. Houbolt also shows that
this method of going to the Moon greatly reduces the amount of mass required to
be launched from Earth. Houbolt outlines
requirements for power, instrumentation, life support, and navigation, as well
as launch masses for different mission architectures. These include a direct to the Moon concept,
Earth orbit rendezvous, and the use of different fuels and launch
vehicles. He considers small, medium,
and large landers, and also suggests the possibility of using two small
landers, because “this combination has a rescue capability not possessed by
direct or other forms of lunar landing missions” (Houbolt, 1961, p.13). Trade-off studies and calculations include
trajectory options, errors in guidance, abort options at different stages of
the mission, possible fuels, size of landers and return vehicles, and the mass
required in Earth and lunar orbit for several architecture options. Notably,
Houbolt’s report includes sections on safety and reliability, development program
scheduling, and facility needs.
Eckart (1999) qualitatively describes the
advantages and disadvantages of six lunar mission architectures. The first
architecture, a Direct One-Way Mission, is particularly beneficial for cargo
missions with expendable transfer and landing stages. A slight variation of
this mission, the One Way Mission with Lunar Orbit Staging is beneficial for
both cargo missions and crewed missions, assuming a return vehicle for the crew
is positioned on the lunar surface. The third architecture, an Apollo-type
Condon and Wilson (2004) describe mission
architectures similar to those put forth by Eckart in quantitative detail.
Condon and Wilson analyze ten different mission profiles grouped into three
architecture types: lunar surface rendezvous (LSR), lunar orbit rendezvous
(LOR), and libration point rendezvous (LPR). They compare the mission profiles
in terms of DV
with the constraint that the crew should not be required to wait longer than
three times the period of the lunar phasing and rendezvous orbit to initiate a
lunar departure. Condon and Wilson conclude that for a sustained, ambitious
program of lunar exploration requiring global access to the lunar surface, a
stay time greater than twenty-eight days, and a capability to abort at any
time, LPR costs less DV
than LOR missions. However, LSR missions require the lowest overall DV,
because they do not require plane changes in lunar proximity or the additional DV
associated with a stopover in EM-L1.
Joosent (2001) analyzes different space
system architectures in the Earth’s Neighborhood, defined as its gravitational
sphere of influence with a radius of 1.5 million km. Joosent specifically
discusses the benefits of using the EM-L1 point as staging area for reaching
high latitude lunar landing sites. The
author also explores different physical architecture designs and elements,
using current space transportation and infrastructure elements (in particular
the Space Shuttle and the International Space Station). For instance, he analyzes the advantages of
using the ISS as a LEO staging facility, to decouple in time that the complex
launch choreography that a long stay Moon mission will require. Joosent suggests that a single “gateway”
located in the EM-L1, will centralize all human deep space operations by
providing accessibility to the Moon, the Earth-Sun Lagrange points, and to Mars
transfer orbits. He mentions that an “Omega” Space Station at the EM-L1 is a
likely complement of the “Alpha” ISS already in place.
Engineers since before Apollo have theorized
and analyzed different architectural possibilities for human Moon
missions. Houbolt did not consider
pre-positioning, because propulsion technologies such as electric propulsion
were not feasible at that time. Current
authors have the option of using existing infrastructure not available prior to
Apollo as well as advancements in launch technologies, navigation and
communication tools, and other technologies.
As the nation’s space program has matured, the technologies available to
space architects have increased substantially.
The number of combinations of possible elements has increased to
thousands, and has made the process of architecture selection a question of
rigorous mathematics, as well as space mission “common sense.” This literature review summarizes necessary
considerations for choosing a lunar baseline mission architecture and relevant
trade studies.
The
requirements for the Moon missions are threefold. First, these missions must demonstrate the
capability to transport humans safely from the Earth to the Moon. This includes launch, transfer, rendezvous,
and landing. Second, these missions must demonstrate the capability to support
humans in terms of life support, communication, and in-space and ground
operations. Third, these missions must
serve as a technology testbed for future Mars Missions and expand our knowledge
of the Moon.
The purpose of the Short Stay Lunar Mission
is to demonstrate the mission capability of going to the Moon. The purpose of the medium stay lunar mission
is acquisition of scientific data and knowledge, and the purpose of the Extended
Stay Lunar Mission is to demonstrate technologies for a mission to Mars
including long duration habitation technologies.
In addition, it should be mentioned that a
crew size of three is used on the Short and Medium missions to the Moon. This crew size was chosen because it is a
smaller scale than the Mars missions, which involve crews of six. The rationale for using crew sizes of six is
discussed later in Section 4.3.2.1. The
Extended+ lunar mission has a crew of six to gain experience with such a large
crew size in preparation for missions to Mars.
The baseline mission assumptions include:
-
Pre-positioned
modules will transit to the staging location using electric propulsion.
-
Manned
mission segments will use cryogenic chemical propulsion.
-
Technology
will be developed to store cryogenic chemical fuel for long durations without
significant boil off.
-
Radiation
and low-gravity countermeasures will be developed by the time Extended Stay
Lunar Missions are performed.
-
Advanced
EVA spacesuits are developed for medium and long duration missions.
-
The
ability to land humans and cargo on the far side of the Moon is developed in
time for medium stay lunar missions.
-
The
ability to separately land humans and cargo within walking distance on the
lunar surface will be developed prior to the Extended Stay Lunar Missions.
Figure
22: Operational view of Short
Stay Lunar Mission
Short Stay Lunar Missions have a crew of
three astronauts and two of these astronauts spend approximately two days on
the lunar surface.
A Crew Operations Vehicle (COV) containing
three astronauts is launched into low Earth orbit on
a man-rated launch vehicle such as a man-rated heavy EELV. A Lunar Lander (LL)
is launched into LEO separately using an STS-derived launch vehicle. Launch vehicles are chosen based on mass estimates
documented in Appendix 9.3. The COV and
LL dock in LEO and transit to lunar orbit together using cryogenic chemical
propellant. Once in lunar orbit, two crewmembers transfer to the LL, undock
from the COV and descend to an equatorial landing site on the near side of the
Moon using cryogenic chemical propellant. One crewmember remains in the COV in
lunar orbit.
The astronauts on the lunar surface will
live in the LL for approximately two days and explore the landing site on foot;
EVA will have minimal science capabilities since the purpose of this mission is
to be a basic technology demonstration.
Upon the conclusion of the surface stay, the
two astronauts ascend to lunar orbit in the LL using cryogenic chemical
propellant and dock with the COV. One person is left in the COV as a safety
measure for the basic technology demonstration; in case the LL fails to dock
with the COV, the astronaut in the COV can manually maneuver to dock with the
LL. Then, the astronauts transfer to the COV, undock with the LL, and initiate
the return trip using cryogenic chemical propellant. The COV performs a
ballistic re-entry, returning the astronauts to Earth.
Animation of short-stay lunar mission
Figure 23: Operational view of
Medium Stay Lunar Mission
Medium Stay Lunar Missions have a crew of
three astronauts and use the same spacecraft forms as the Short Stay Lunar
Missions. However, there are three differences between the Short and Medium
Stay Missions: the LL is pre-positioned in lunar orbit using electric
propulsion, all astronauts transfer to the LL to descend to the lunar surface, and
the astronauts stay on the lunar surface for one week.
The justification for pre-positioning the
Lunar Lander in lunar orbit before the arrival of the crew is to test the
technology of pre-positioning essential mission cargo utilizing electric
propulsion technology. While not
providing a major mass savings for missions to the Moon, the capability of
pre-positioning will allow for dramatic mass savings for missions to Mars. This is one of the ways the Moon can be used
as a testbed for future missions to Mars.
First, a LL is launched into LEO alone using an STS-derived launch vehicle.
Electric propulsion is then used to pre-position the LL in lunar orbit. Later,
a COV containing the three astronauts is launched into low Earth orbit using an EELV (Delta IV Heavy). The COV transits to lunar
orbit together using cryogenic chemical propellant. Once in lunar orbit, the
COV docks with the pre-positioned LL, the three crew members transfer to the
LL, undock from the COV and descend to non-equatorial landing sites on the near
side of the Moon using cryogenic chemical propellant. No crewmembers remain in
the COV in lunar orbit; it is assumed LL ascent was proven to be reliable
during the Short Stay Lunar Missions.
The astronauts on the lunar surface will
live in the LL for approximately one week and explore the landing site using an
“open-air” rover to aid mobility within walking distance from the LL; EVA will
have high science capabilities including research in some of the areas outlined
in Section 3.2.1: Scientific Knowledge.
Upon the conclusion of the surface stay, the
three astronauts ascend to lunar orbit in the LL using cryogenic chemical
propellant and dock with the COV. The astronauts transfer the COV, undock from
the LL, and initiate the return trip using cryogenic chemical propellant. The
COV performs a ballistic re-entry, returning the astronauts to Earth.
Animation of medium-stay lunar mission
Figure 24: Operational view of
Extended Stay Lunar Mission
Extended Stay Lunar Missions have a crew of
six astronauts and require a pre-positioned surface habitat for the crew to
live in for up to six months. A crew of
six is used for this mission since future missions to Mars will have crew sizes
of six.
It is important to mention that the
increased complexity of the Extended Stay Lunar mission is essential for two
major reasons. First, the increased
complexity results from the multitude of technologies being tested in preparation
for missions to Mars. Second, the
increased complexity is used to return even greater amounts of knowledge back
to Earth. This is made possible from a
crew living on the Moon for an extended period of time in a location that is
highly valuable from a knowledge point of view.
First, a surface habitation module (SHM) is
launched into LEO possibly using two STS-derived launch vehicles. The SHM is
then pre-positioned on the lunar surface, using electric propulsion for the
transit to lunar orbit, and cryogenic chemical propulsion for the descent.
Second, two Lunar Landers are launched into LEO separately using STS-derived
launch vehicles. Electric propulsion is then used to pre-position the Lunar
Landers in lunar orbit. One module of the habitation module (HM) is launched
into LEO along with one of the Lunar Landers. Third, a COV containing three
astronauts is launched into LEO using a man-rated launch vehicle such as a
man-rated EELV. The same launch vehicle
also contains a Modern Command Module (MCM).
Each of these vehicles contains three astronauts. The COV docks with the HM and MCM, the crew
of the MCM transfers to the HM, and the MCM undocks.
The docked COV and HM then transit to lunar
orbit using cryogenic chemical propellant. In lunar orbit, the pre-positioned
LL1 docks with the COV and HM, the three crewmembers transfer to the LL1, the
LL1 undocks from the COV and descends to the pre-positioned SHM on the far side
or pole of the Moon using cryogenic chemical propellant. Likewise, the second
pre-positioned Lander, LL2 then docks with the COV and HM and transfers the
crew to the SHM. No crewmembers remain in the COV in lunar orbit.
The astronauts on the lunar surface will
transfer from the Lunar Landers to the SHM for a surface stay of approximately
six months. The semi-permanent base allows for extensive science capabilities,
possibly including but not limited to Moon-based observatories, greenhouse
technology demonstrations for closed-loop life support, and nuclear power
production. A habitable, pressurized rover for overnight field trips will aid
surface mobility.
At the end of the surface stay the six
astronauts ascend to lunar orbit in the two Lunar Landers. Each Lander individually docks with the COV
and HM. After each docking, three astronauts transfer to the COV and then the
COV and Landers undock. Finally, the
astronauts initiate the return trip to Earth in the COV using cryogenic
chemical propellant. The COV and HM use aerobraking to establish Earth orbit
and then dock with the MCM. Three astronauts transfer to the MCM to return to
Earth. The other three astronauts remain in the COV, undock from the HM, and
return to Earth.
Animation of long-stay lunar mission
The lunar baseline missions were designed
with a stepping-stone approach; complexity is added to missions in stages. For
example, the SSLM requires one COV and one LL. While the MSLM requires the same
forms as the SSLM, complexity is added in two ways: first, the LL is
pre-positioned. Second the LL remains
unmanned in lunar orbit. The ESLM requires another increase in complexity to
prepare for future Mars missions. For these missions, three astronauts travel
to LEO in the COV as in previous missions, however three other astronauts
travel into LEO in a new form, the MCM. The transit to the Moon is carried out
with the COV and another new form, the HM. Finally, for this mission, two LL
are pre-positioned rather than one.
This section describes Mars technology
demonstrations carried out on Extended Stay Lunar Missions. The short and
medium stay missions are not intended for technology demonstration, although by
necessity certain technologies will be demonstrated on these missions.
Lunar Lander (LL)
The Lander
technologies that must be demonstrated on the Moon for the future Mars missions
are:
·
Slow
descent engines. The
Mars mission will initially use aerobraking and a parachute to decelerate to
the surface. Once near the surface,
descent engines are used to touch down; descent engines will be demonstrated on
the Moon.
·
Ascent
stages. The Mars Lander will use a staged ascent, and
the Lunar Lander will not; nevertheless, the “launch pad” technology used for
ascent is similar for both Moon and Mars missions.
·
Reduced
gravity. The gravity constraints are different for the
two planets; the Moon has 1/6-G and Mars has 3/8-G. The Moon will be used to demonstrate reduced
gravity landings, but due to the gravity disparity, the Moon can only serve as
a partial testbed.
·
Life
support. Both Landers will use the same life support
systems.
·
Ability
to land unmanned. For
rescue capabilities, an unmanned orbiting Lander should be able to land,
unmanned, and rescue crewmembers on the lunar or Martian surface. This is not a
direct requirement for the Martian missions, but may be used in future Mars
missions. The capability is deemed
important enough for the lunar missions alone, that it will be demonstrated for
Mars, even if the unmanned Lander is ultimately not used for the Mars
missions.
Surface Habitat Module (SHM)
The SHM technologies to be demonstrated on
the Moon for the future Mars missions are:
Rovers
A pressurized version of the rover will be
used for Extended Stay missions to the Moon as well as Extended and Extended+
Mars missions. The rover technologies
that must be demonstrated on the Moon for future Mars missions are:
Other technology demonstrations:
Pre-positioning
Successful pre-positioning must be
accomplished during lunar missions before a
Mars mission is performed.
The components that must be pre-positioned
in LEO are:
The components that must be pre-positioned
in lunar or Martian orbit are:
The component that must be pre-positioned on
the lunar or Martian surface is:
Docking
Successful docking must be accomplished
during lunar missions before a mission to Mars.
Components that must dock for both lunar and martial missions are:
These components will need to be designed
considering the requirements needed for docking in orbit. Maneuvering propulsion systems as well as
guidance, navigation, and control capabilities may be required on board these
modules to facilitate on-orbit docking.
Unmanned Orbiter
Both the lunar and Martian missions will
involve an unmanned orbiting craft around the destination. For example, in the Extended lunar mission,
the docked HM and COV will dock with the Lunar Lander, to transfer the crew,
then the HM an COV must be able to orbit, unmanned, for six months or longer
before the crew ascends and docks to transfer for the return trip. The capabilities of orbiting unmanned, and
maintaining life support capabilities for the return trip, must be
demonstrated.
The Moon is a natural laboratory for
studying planetary and geologic processes, described more fully in Appendix
9.6. Among many others, we wish to explore and expand our knowledge in the
areas of:
In addition to the scientific and resource
knowledge potential of the Moon, it can also be used as testbed for future
technologies such as in-situ propellant production (ISPP) or as the location of
an astronomical observatory.
The knowledge delivery infrastructure will consist
of two parts: the delivery of data in the form of bits and the delivery of
samples from the planets surface. This
section is specifically about the delivery of knowledge in the form of bits. This delivery of bits is referred to as the
communication delivery system.
The same communication radio frequency has
been selected for all lunar and Martian missions in order to provide an easily
extensible system. The radio frequency
that each of these missions will use is Ka-Band or 32 gigahertz. This frequency was selected because it can
support a high data rate with comparably lower power than all lower frequency
bands, and because the Deep Space Network (DSN) ground infrastructure will
support it by the year 2007, while other higher frequency bands are not
supported by the DSN. There is some
concern about weather interference especially when communicating with Mars,
however a Martian sand storm would prevent an X-band communication as it would
a Ka-band communication, the differences would mainly lie in the moderate
weather such as a cloudy day, or light dust storm in which case the Ka-bands
data rate would be decreased.
For the Short Stay Lunar Missions, a direct
link can be set up between the Lunar Lander and one DSN station. This would allow constant communication
between Earth and the Moon throughout the mission. The data rate required for this mission would
be approximately 0.07 megabits/sec and would require 0.01 Watts of constant
power per transmission forty minutes.
After the mission is completed the communication equipment that placed
on the Moon will be left there for two reasons. First, if a future mission
decides to return to the same location, placing new equipment at the
destination will not be necessary. Second, in the unlikely case that mission
communication equipment fails, the crew may have the option of traveling in a
rover to a previous landing site to use the equipment left at that site.
For the Medium stay missions, the
infrastructure is essentially the same as the Short mission. The main difference, however, is that the
medium mission requires higher data rates and transmitting power. The
transmission data rate will be 0.7 megabits/sec with a required constant power
of 0.1W over the forty minutes of transmission time
The long stay missions require the ability to communicate between the far side of the Moon and Earth. The astronauts will communicate through one of four possible ways. For the first option, a communications relay satellite placed at the L4 point in the Earth Moon system. This infrastructure allows for a constant communication stream between the Earth and the Moon. Unfortunately, this option would only allow for communication for over 900km on the Moon’s far side at its center or 30 degrees off the far side facing the satellite. The next option is to set up a relay satellite in low lunar orbit. This infrastructure is capable of covering most of the Moon, but a significant time delay for far side communications. The third option and involves placing a satellite in an orbit around the EM-L2 point. The coverage area provided by this communications satellite entirely covers the far side of the Moon and maintains nearly continuous communication with Earth. The shortcoming of this infrastructure lies in the difficulty of establishing orbit around EM-L2. The transmission data rate for the long stay missions will be 3.5 megabits/sec with a required constant power of 0.5W over the forty minute transmission time for near-side operations and a required constant power of 20W for far-side operations over the same transmission time. Please note that the equations used to determine these numbers are shown in Section 9.5.2.
Humans
have dreamed of travel to other planets for centuries. In particular, Mars has been the focus of
much interest. Arguably the first actual
mission design for Mars was presented in 1952 by rocket engineer Dr. Wernher
von Braun (History of Humans to Mars website, 2004). The plan was immense in scale and involved a
fleet of interplanetary spaceships carrying large crews to Mars. Because of the corresponding cost
requirement, von Braun’s design it did not become reality. Although the Space Race between
In 1989
political interest in Mars travel was revived with President Bush’s call for a
Space Exploration Initiative, and it resulted in a study known as “The 90 Day
Report.” This report resulted in a
projected mission to Mars cost estimate of $450 billion. The mission design required assembly in orbit
of a 1000 tonne spacecraft as well as a large orbiting facility to enable this
assembly (History of Humans to Mars website, 2004). Again the high cost ruled out this
humans-to-Mars initiative, however, at the same time, a radically low cost
architecture was being designed.
The Mars
Direct plan, developed by Robert Zubrin and David Baker at Martin Marietta
Astronautics Company was a revolutionary paradigm shift that has had a
significant impact on Mars mission design.
The focus of this plan was to “live off the land” as much as possible,
using the Martian atmosphere and soil to provide resources and in particular,
using the atmosphere to enable in-situ propellant production (ISPP) of methane
and oxygen for the return journey. The
resulting mass required in LEO dropped dramatically, as did the projected cost
that was reduced to approximately twenty to thirty billion dollars (Zubrin,
1996).
The Mars
Direct approach was well received at NASA under a new administration and NASA
performed its own study and corresponding mission design in the late
1990s. This study was known as the
Design Reference Mission (DRM) (Hoffman, 1997).
This study relied on Zubrin’s ideas including direct travel to the
Martian surface and ISPP, although it was scaled up in some respects, having a
crew size of 6 instead of 4. The cost
estimate of such a mission was fifty billion dollars. The significant difference was a two-stage
departure from Mars as opposed to the direct departure planned by Zubrin. The DRM had a Mars Ascent Vehicle (MAV) that
rose from the surface of Mars via ISPP to rendezvous with an Earth Return
Vehicle (ERV) for the trip back to Earth.
The trans-Earth injection was predicated on the use of conventional
fuel.
Other
agencies and companies have proposed Mars mission architectures in recent
years. One ESA scheme designed for the
EADS
Space Transportation has outlined a number of options for short and extended-stay
missions (Ransom, 2003) with short missions using Mars Orbit Rendezvous (MOR)
and landing vehicles for crew ascent from and descent to the surface and with
longer missions using ISPP and an architecture similar to NASA’s DRM. For the short stay missions, the main options
are: pre-positioning, the use of ISPP to fuel ascent vehicles, and dual Lander
architectures that allow the exploration of more than one site. Longer stay options include larger crew
sizes, and fully established infrastructure.
Global
Aerospace Corporation presents an interesting approach to infrastructure on
Mars (Nock, 2001) with the use of “Astrotels” cycling between Earth and Mars,
and “taxis” operating on rendezvous trajectories between these Astrotels and
transport hubs such as orbiting facilities.
This concept makes use of highly autonomous on-board systems to control
the operations of the vehicles when they are not crewed. As well, it details
the potential of harvesting LOX from the Moon as well as Phobos and using these
resources to fuel interplanetary transfer.
Two main
points emerge from this discussion. One is that ISPP is a relevant enabling
technology with a good potential for decreasing costs and allowing Mars
missions to be performed more frequently.
A second point is the sense that there is not a single correct
architecture. It is important to include
options for scalability and extensibility from current capabilities to reduce
cost and maintain flexibility in mission design; however, this can be
accomplished in a number of different ways, each with its own benefits and
penalties.
The following Mars mission architectures are predicated on the idea of an evolution of mission scale and complexity starting with a rationale for using Phobos as a preliminary step. This methodology is closely aligned with the message presented in President Bush’s New Vision for Space Exploration Program document. In addition, the mission architectures defined are reliant on successful demonstration of certain technologies and operations on the Moon. As our operations knowledge increases, utilization of complex technologies and other mission enhancements becomes feasible and these enhancements are incorporated into the mission profiles. The aim is that this evolution of missions in the space exploration system architecture will provide a means to create a sustainable transportation network and infrastructure for travel to Mars as well as to develop the capacity for missions to future destinations.
It is important at this point to mention how
the crew size used in the following mission profiles was determined. A manned mission to Mars will require trip
times well in excess of any mission thus far.
This extended time in relative seclusion from all other personal contact
suggests the need for a large crew for psychological considerations. However, this need must be balanced with mass
requirements for each additional crewmember.
During the mission, each astronaut will be required to perform a number
of functions during the mission due to long communications delays require a
higher level of autonomy. There appear
to be five relevant technical fields required for exploration: mechanical
engineer, electric engineer, geologist, life scientist, and physician. In addition to the primary specialty, each
crewmember would need to be cross-trained in another mission critical field and
be responsible for a variety of support tasks during the mission. However, with only a crew of five, a single
loss of a crewmember, even temporarily in the event of sickness, could
jeopardize the mission. Thus a crew of
six is recommended. (Hoffman, 1997)
The plan for launching the required equipment needed for missions to Mars into LEO is mentioned in this section. In order to achieve the enormous mass required in LEO for a manned Mars mission, a number of launches must occur. For a Short Stay mission, a number of elements are pre-positioned in Mars orbit. These elements are transported using electric propulsion and thus must be launched approximately two years prior to crew departure. Using the STS-derived cargo launcher, three launches are required to place the propulsion for Earth return, the Martian Landers and the surface habitat in LEO. Prior to the crew launch, the Habitation Module, which houses the crew during transit to and from Mars, is launched unmanned into LEO using two heavy lift cargo launch vehicles, such as an STS-derived launch vehicle. Finally, the crew travels to LEO in two separate vehicles: the Modern Command Module and the Crew Operations Vehicle. Both of these components can be launched by a single man-rated launch vehicle such as a man-rated heavy EELV. Thus, for a short stay mission, a total of five STS-derived cargo launches and a single man-rated EELV are required.
For an Extended Stay mission, an even larger mass is required to be pre-positioned at Mars due to a prolonged surface stay. Two years prior to crew departure, the pre-positioned elements using electric propulsion are launched via four STS-derived cargo launchers. Prior to launch, the Habitation Module is launched in two separate components on two STS-derived cargo launchers. Finally, the crew is launched in the crew operations vehicle and modern command module, using a single man-rated launch vehicle. Thus, a total of six STS-derived cargo launchers and a single man-rated launch vehicle are required.
Each mission defined below has some common assumptions built into the design. These assumptions are detailed here to emphasize commonality and avoid unnecessary repetition. For each mission to Mars a crew size of 6 and the use of chemical propulsion for crew transfer have been assumed. Electrical propulsion is used for pre-positioned elements such as cargo and Landers.
Since this report focuses on a strategy to achieve sustainability, certain specific mission details are not explicitly stated. These details are assumed. They include such items as radiation shielding and various features of the life support system such as air-regeneration facilities.
The general mission architecture is a Mars orbit rendezvous (MOR). The primary benefit of MOR is the decoupling of in-space transportation from descent and landing operations. MOR also allows flexibility with regards to the timing of landing, and this may prove to be important in the case of bad weather on Mars. Furthermore, because MOR involves landing from orbit as opposed to directly from the trans-Mars injection, landing precision can be improved. This allows architectures to include a pre-positioned surface component, which corresponds to a reduction in propellant mass.
The Martian moons, Phobos (Fear) and Deimos (Terror),
were discovered in 1877 and have since been the subject of long-range
observations made by Earth-based telescopes as well as by spacecraft traveling
to Mars. Aside from their inherent
scientific interest, these two Martian satellites could play a significant role
in the context of Mars exploration and in the extensibility of NASA’s space
exploration initiative.
Figure 25: Phobos
In constructing a Mars exploration strategy, one of
the principal objectives is to expand human knowledge and thus engage the
public. For this reason, it is
anticlimactic to send humans on a voyage of over a year in duration only to
enter Mars orbit before returning to Earth.
However, landing on Mars, which is likely to be achieved via aerobraking
and rendezvous or docking operations, is one of the most dangerous parts of the
mission. To mediate the risks of landing
and surface habitation and obviate the requirement of a human mission to
Martian orbit, it is suggested that the preliminary Mars mission have a surface
component on one of the Martian satellites, Phobos or Deimos.
New technology must be tested as much as possible
before the first manned mission to Mars, and this will be accomplished
primarily by using the Moon as a testbed.
Nevertheless, due to the inherent differences between lunar and Martian
exploration missions, the first human mission to Mars will still be, in some
respects, a venture into the unknown. A
preliminary mission to Phobos or Deimos would allow NASA to decouple the test
of the different components of the proposed space transportation system. Humans
would be sent on a Martian trajectory and would therefore prove the capabilities
of the transportation system and vehicles without having to perform the
dangerous task of landing on Mars.
Instead, the landing maneuver would be limited to a “docking” procedure
in the micro gravity (1/1000g) environment of Phobos or Deimos, and a surface
stay in a location without atmosphere similar to that of the Earth’s Moon.
A preliminary mission to Phobos or Deimos would also
lend itself to extensibility because it could be a stepping stone goal on the
way to other exploration initiatives. A
human mission to one of the Martian satellites would allow a telerobotic
presence on Mars whereby the crew could control rovers on the Martian surface
and respond to events of interest, taking advantage of the minimal
communications delay. This could also
aid in landing site certification. In
addition, because these two moons may be captured asteroids or at least very
similar in composition, a mission to these bodies could prepare NASA for a
future asteroid rendezvous mission while taking advantage of the moons’
predictable orbits (as determined precisely by the 1988 Soviet Phobos mission
(Brat, 2001). Looking further ahead, a
mission to one of the Martian moons may help to provide NASA with some of the
operational knowledge necessary for future exploration of the moons of Jupiter.
Recent initiatives by the Planetary and Space Sciences
Research Institute at the Open University in the
Finally, a successful mission to one of the Martian
moons would build public confidence and spark interest in the space exploration
problem. This would enable NASA to gain
support for subsequent missions to Mars.
Phobos has been chosen as the de facto destination between the two Martian satellites since it is larger and closer to Mars. A mission to Phobos would require remote sensing capabilities to survey the surface for potential landing sites during initial flybys. A landing capability would also be required to allow surface exploration, and dedicated scientific payloads would be used to collect samples and perform geological testing.
Since a crew would be sent to Phobos,
a high priority would be to minimize transfer time and hence the trans-Mars
injection would probably follow the opposition-class profile, making use of a
Venus flyby. The trajectory to Phobos
can be broken into seven sections as follows with various propulsive
requirements (Brat, 2001):
Of these maneuvers, only the TMI and
capture at Mars would have significant ∆V requirements, on the order of 4 km/s and 2.7 km/s, respectively. Energy requirements could be reduced via
aerobraking by approximately 1.2 km/s (Brat, 2001), but this would add to
mission complexity and could be maintained as an option if a test of
aerobraking is required. The return
trajectory requires a trans-Earth injection and ballistic entry at Earth’s
atmosphere.
The short-stay mission is the shortest Mars mission possible in terms of total mission duration; it is composed of approximately 600 days transit time and 60 days surface stay (Walberg, 1993). The crew travels to Mars via an opposition class free-return trajectory with a Venus fly-by in the Mars Transfer Vehicle (MTV), which is composed of a Habitation Module (HM) and a Crew Operations Vehicle (COV). Upon arrival at Mars, the MTV aerocaptures into Martian orbit and performs a rendezvous with two pre-positioned Mars landing vehicles (ML1 and ML2). Three crewmembers descend to the Martian surface in each Lander. This allows flexible timing for each landing, with the second being contingent on the success of the first. The landing is achieved using a heat shield for atmospheric entry after which parachutes are deployed to slow the spacecraft. The final stage of the landing is a powered touchdown that gives the crew as much control as possible over the landing so as to minimize risk of damage to the landing vehicle.
The crew remains on the surface for approximately 60 days. During this time, the crew lives in a pre-positioned surface habitat that could be extended by an inflatable module if more volume is required. At the end of the surface stay, the crew returns to Mars orbit in the two landing modules and docks with the MTV. The MTV docks with the pre-positioned return propellant module (SM2) and executes a trans-Earth injection maneuver. Entry back into low Earth orbit (LEO) is achieved via aerocapture, and the MTV docks with the Modern Command Modules (MCM) allowing half of the crew to transfer into this Earth re-entry vehicle. The other half of the crew returns to Earth’s surface in the COV. An operational overview of the short-stay mission to Mars is provided in Figure 26.
Figure 26: Short stay mission to Mars
During the surface stay, the crew will explore the Martian surface within close range of their landing site using EVA suits in conjunction with an un-pressurized rover. They will perform activities such as searching for water and life, collecting samples, imaging the surface, and recording weather phenomena. Requirements will include life support facilities in the surface habitat, radiation shielding, and consumables. An adequate communications network will also be required for this mission as discussed in further detail in Section 4.3.4.
To facilitate the short-stay mission, pre-positioning of mission elements is employed. In addition to the pre-positioned Martian Landers, surface habitat, and return fuel; surface equipment, such as an un-pressurized rover and scientific payloads will be pre-positioned on the surface. Pre-positioning requires a farsighted approach to mission planning, but this should be well within the scope of the space transportation system outlined here, and as discussed in Section 6.4.3.1.4.1, the benefits of pre-positioning justify the added constraints on mission timeline. Because the pre-positioned elements will make use of electric propulsion, they will have to be sent at the launch window approximately two years prior to that of the crew departure.
The description above outlines the basic structure of the short-stay mission; however, there are several options available for such a mission that would improve prospects of extensibility by testing enabling technologies for a subsequent mission. The first, most important option is to begin verifying and testing an in-situ propellant production (ISPP) plant. An ISPP plant would allow for fuel production on the Martian surface as described in further detail in Section 6.4.3.3.2.1. This option would require a pre-positioned propellant production module, a small nuclear power plant, and a hydrogen fuel stock for a Sabatier process. During the surface stay, the crew could check the functionality of the ISPP system and determine if it is reliable for propellant production. In addition, the propellant could be used to augment power for surface systems, such as the rover vehicle. This option would facilitate an extensible and sustainable infrastructure for further Mars missions.
Another option is to provide extended surface mobility. This could be accomplished by providing one or more remote-controlled rovers to assist in surface exploration. These rovers could act as “scouts” and be dispatched to areas of potential interest to determine which site holds the most interest for a follow-up visit by the crew.
Animation of short-stay Mars mission
The extended-stay mission offers the advantage of a longer surface stay with only a small increase in the total mission duration over the short-stay mission. It is also less complex from a trajectory perspective since it does not require a Venus flyby to provide the required ∆V. Instead, for this mission, the crew travels to Mars in the MTV via a fast transfer conjunction class trajectory. In most respects the mission architecture is taken directly from the short-stay mission. Assuming the short-stay mission precedes the extended-stay, the required operational knowledge to carry out the mission should be almost complete. An operational view is presented in Figure 27.
Figure 27: Extended stay mission to Mars
One of the main differences between the extended and short-stay missions is the ascent to Mars orbit in the landing modules. Assuming the option of ISPP during a short-stay mission is employed, and successfully demonstrated, the ascent propellant for the Landers will be provided in-situ using the Sabatier process. Details about the Sabatier process are available in Section 6.4.3.3.2.1. This will change the mission requirements slightly in that an ISPP unit consisting of hydrogen feedstock, a chemical plant and a nuclear power plant will have to be sent to Mars two years prior to the departure of the crew, and it will necessarily have autonomous communications ability to relay its status to ground control centers on Earth. This way, the crew will be assured that the propellant required for their ascent from Mars has been generated prior to their departure. In addition, since the Sabatier process produces methane/oxygen fuel, the landing module will have to be able to dock with the required propellant tanks and engine. All of these components may have been pre-positioned prior to crew departure or may be attached to the landing module itself. Using ISPP to fuel the Landers’ ascent will help to reduce the mass of pre-positioned cargo and thus save mass or enable the transportation of other science payloads. A further advantage of using ISPP is that the fuel produced may also be used to power combustion engine rovers and possibly life support systems as well. Finally, the functional test of ISPP at this stage will act as a stepping-stone towards the eventual goal of using ISPP to fuel the entire return journey to Earth.
Another principal distinction between extended and short-stay missions is the length of surface stay. For an extended-stay, the crew surface habitation module will need to be considerably larger than that of a short-stay mission. To this end, an inflatable module or an additional habitation module will be pre-positioned
Besides the obvious extensive life support requirements for a mission of this duration, the crew will require an extended means of exploration equipment for this mission. Along with EVA suits, a transport vehicle, open or pressurized, will be pre-positioned, to enhance surface mobility. An open, un-pressurized, rover is limited to ranges of 10km, such that the crew is always within walking distance of the surface habitat. However, this safety requirement highly constrains the amount of exploration that can be accomplished during this long duration mission. Therefore, although not a requirement, a pressurized rover capable of ranges on the order of 500km is a recommended option for the long-stay mission. This will allow the crew to explore a large area, searching for water and life, collecting samples to return to Earth, and taking various measurements.
In addition to the large-scale physical exploration of the Martian surface, the crew will have the opportunity to conduct more advanced scientific experiments including experiments that may require a longer duration, such as small-scale agriculture development. The construction of an inflatable greenhouse prototype is one option for the extended-stay mission. This could supplement the crew’s food supply for both the surface stay and Earth return trip.
Another important option to consider would be to include a pressurized surface transportation vehicle to satisfy the surface mobility requirement. This pressurized vehicle, with a mass of approximately 6000kg, would allow for a “shirtsleeve” environment for up to three astronauts on sorties on the order of 500 km and 400 hours duration. For safety reasons, two pressurized rovers would have to be sent to Mars to ensure rescue capability. It is likely that these vehicles would be sent separately from the crew via a direct launch in conjunction with other cargo such as extra food supplies or a power plant.
Further options that would facilitate movement towards a semi-permanent infrastructure include the development of ISPP to a high enough level of production and reliability, such that it would be used to supply the return fuel for the MTV in subsequent missions. Also, a drilling device that could be transported as pre-positioned cargo to tap into subsurface aquifers if it was determined in the short stay or robotic precursor missions that underground water may exist. Since surface life support systems require a substantial power supply, a nuclear power plant could be pre-positioned. Finally, a closed loop life-support system with bioregenerative components could be tested. Although this technology would not be implemented until its reliability is ensured, such a technology would help a self-sustained presence on Mars become more feasible.
Animation of extended-stay Mars mission
A comparison of the ΔV requirements for the Short and Extended Stay missions to Mars is shown in the following tables. Table 5 shows the ΔV requirements for the Short and Extended Stay missions to Mars assuming aerobraking and parachutes are not used. Table 6, on the other hand, assumes parachutes can be used in the missions.
Table 5: ΔV requirements assuming parachutes and aerobraking not used
Orbital Maneuvers |
Short DV (m/s) |
Extended DV (m/s) |
Trans-Mars Injection |
4098 |
4217 |
Mars Orbit Insertion |
3278 |
3050 |
Mars Surface Descent |
741 |
741 |
Mars Surface Ascent |
4140 |
4140 |
Trans-Earth Injection |
1415 |
2180 |
Earth Orbit Insertion |
2774 |
5485 |
Total DV (m/s) |
16446 |
19813 |
Table 6: ΔV requirements assuming parachutes used
Orbital Maneuvers |
Short DV (m/s) |
Extended DV (m/s) |
Trans-Mars Injection |
4098 |
4217 |
Mars Orbit Insertion |
3278 |
3050 |
Mars Surface Descent |
111 |
111 |
Mars Surface Ascent |
4140 |
4140 |
Trans-Earth Injection |
1415 |
2180 |
Earth Orbit Insertion |
2774 |
5485 |
Total DV (m/s) |
15816 |
19183 |
If it aerobraking is assumed for use in one of the Mars missions, the ΔV numbers in Table 6 for the stages of the mission in which aerobraking is used, such as Earth orbit insertion, are ignored for propulsive engine burn requirements. Instead, the initial mass required for the aerobraking maneuver is assumed to be 15% of the payload mass (Walberg, 1993).
The final mission class is the
extended-stay mission with the development of infrastructure. The idea behind this architecture is that if,
after previous short and extended-stay missions, Mars remains an interesting
destination either from a science or operations perspective. Another possibility is that Mars can serve as
a testing ground for the next exploration target. Subsequent Mars missions will then develop
infrastructure to facilitate surface stays and exploration as well as to
minimize mass that has to be transported from Earth.
In this case, the aim of the
mission is to use in-situ resources as much as possible. These resources are used to provide return
fuel, generate power, develop sustainable agriculture, and enable closed loop
life support. Missions to Mars will take
place more frequently, possibly with one crew traveling at every launch
opportunity
The initial architecture will
follow the proven MOR scheme for a long-stay conjunction class mission, but
assuming previous attempts at ISPP generation and fuelled ascent have been
successful, it is likely that the architecture design will become more similar
to Mars Direct (Zubrin, 1996). The
eventual outcome of this transition would be that the MTV would travel directly
to the surface of Mars without orbital rendezvous and ascend from the Martian
surface using ISPP fuel directly into a trans-Earth injection. The result of a move towards this
architecture would be a significant reduction in IMLEO.
For these extended-stay missions, pressurized transport vehicles will be pre-positioned allowing the crew to have significant surface mobility on the order of 500km. As previously mentioned, two pressurized rovers will be necessary to provide a rescue capability. The crew will conduct science experiments as described above. In addition, the scientific payloads will be chosen to explore areas that have been proven to be the most interesting in previous missions as well as any other new areas of interest. The ECLSS will be designed to achieve as close to 100% closure as possible, and the crew will derive most of their power from ISPP. Agricultural facilities such as inflatable greenhouses will be installed to provide or supplement the crew’s food supply. The crew habitat will take the form of multiple inflatable modules as well as pre-positioned Habitat modules sent direct from Earth.
The extended-stay mission with infrastructure will provide a testbed for further exploration technology development, and it is also possible that the ISPP facilities will allow Mars to serve as a way station for vehicles traveling to more remote destinations.
The four missions described above, from a preliminary mission to Phobos through to a mission for an extended stay on the Martian surface, are designed for logical evolution from one to the next. With this in mind, there is significant commonality between the missions. In all cases, one mission provides a testbed for the technology that is required for subsequent missions. A Phobos mission presents the opportunity to test the in-space transportation system that will be used to reach Mars during the short-stay mission. Furthermore, the short-stay mission allows the test of aerobraking, landing, surface habitation, and in-situ propellant production that will all be used during the extended-stay mission. Finally, during the extended stay mission, more advanced technologies such as agriculture can be tested in preparation for the next mission with infrastructure. The missions to Mars as described above would also be useful preparation for further destinations such as asteroids or Jovian moons.
For the short-stay sized Mars missions a
direct link can be set up between the Mars
Lander and one of the Earth’s Deep Space
Network (DSN) stations. This would allow
semi-frequent communication between Earth and Mars throughout the entire
mission. The data rate required for this
mission would be 1 gigabit/day and would require 8 watts of power per
transmission with a transmission data rate of 0.04 megabits/sec. After the mission is completed the
communication equipment that was landed on Mars will be left there for two
reasons, one if a future mission decides to use that spot as a landing or settlement
site then they won’t have to bring their own equipment, and in the unlikely
case that another future mission communication equipment fails the crew will
have the option of traveling in a rover to the old site and using its
equipment.
The extended-stay missions require the ability to communicate with much greater data rate and thus it might be necessary to create a relay satellite around Mars. There are two realistic options for the location of this satellite. The satellite could be placed in a Geostationary Martian orbit (GMO) around the landing site; the advantage of a GMO satellite is that it increases the time that the astronauts can communicate with the Earth, the disadvantage is that it can only really be set up for one portion of the planet. The other option is to position a satellite at the Earth-Mars L1 point thus decreasing the power required to send large communication streams to the Earth; unfortunately this would not add any extra time that the mission could communicate with the ground. As with the Moon missions there is an option in the case of emergencies to communicate with the Moon and use it as a relay station. The daily data rate for a large sized mission would be 10 gigabits/day, the transmission data rate would be 0.4 megabits/sec and the transmission power required would be approximately 80 W. Please note that the equations used to determine the numbers used in this section are in Section 9.5.2.
The purpose of this section is to highlight the commonality between the Moon and Mars missions within the framework of a baseline set of forms that will each perform certain functions in order to complete a baseline mission. The forms used to accomplish a baseline mission are described below.
The baseline mission suggests forms for specific functions. The purpose of developing a baseline mission is twofold. Firstly, with forms specified, the missions to the Moon and Mars can define functional requirements that are desired of each form. Secondly, specification of forms creates a framework for comparing missions and evaluating the attributes of each mission.
The method of selecting this baseline mission was based on the first draft of requirements for a Mars mission. A Mars mission was selected as a starting point and analyzed considering the impact of each mission decision on the extensibility of this architecture from a Moon mission. The architecture was designed considering reconfigurability, adaptability and extensibility of a Moon mission to a Mars mission. Each decision required specification of the baseline architecture and evaluation based on its influence on a Short, Extended or Extended+, Mars or Moon mission. Once each form was selected, specific details of the functions for each form were specified for the Mars and Moon missions.
A baseline mission was developed to create a complete list of forms required to accomplish the Mars and Moon missions. The baseline mission includes assumptions based on current technology limitations, safety concerns, and policy requirements to a lesser extent. These assumptions influence the technology of each form, but not the functional requirements of the forms. The focus of the baseline mission was not to develop another detailed description and analysis of the forms required for an extensible Mars mission. Instead, a framework is created that provides a method of comparing the functional requirements for each of the Mars and Moon missions. Details of the forms are provided in Section 6.4.3.1.
A schematic
representation of the baseline mission is shown visually in
Figure 28
and can be summarized as,
·
Pre-position in
LEO (HM, SM1)
·
Pre-position in
Destination orbit (SM2, ML1, ML2) (LL1, LL2 for Moon mission)
·
Launch MCM and
COV into LEO on man-rated launcher.
·
Sequentially dock
MCM and COV to HM (undock MCM and leave in LEO)
·
Dock SM1 to
COV/HM (This combination is known as the MTV)
·
Transit to
destination, undock SM1
·
Burn into Lunar
orbit or aerocapture into Mars orbit (mission specific)
·
ML1&ML2
(LL1&LL2 for Moon mission) sequentially dock and undock with MTV
·
Crew ascends to
destination surface and ascends to orbit
·
ML1&ML2
(LL1&LL2 for Moon mission) dock and transfer crew to MTV (ML1&ML2 stay
in orbit)
·
Dock SM2 to MTV
·
Transit to Earth,
undock SM2
·
Aerocapture at
Earth
·
Dock MCM to HM,
transferring crew of three
·
Remaining crew of
three enter COV
·
COV & MCM
Earth EDL
Many decisions regarding sustainability and
extensibility were made in determining this baseline mission. As such, a number of trade studies were
considered and presented in later sections of this report (see Chapter
6).
Figure 28: Schematic representation of the Moon and Mars Baseline missions
Pre-positioning crew habitation modules, fuel for return transport, Landers, and return capsules allow the mass of the module transporting the crew to the Moon or Mars to be reduced. Although pre-positioning modules reduces the injected mass required for a mission, additional difficulties are adopted. These difficulties include on-orbit docking capability and increased complexity in pre-positioned module functionality. Safety concerns are also raised because pre-positioning return fuel for the first mission to Mars will require a high level of confidence in the technology. As such, carrying return fuel for the first Mars mission increases confidence while providing a method of developing a stockpile of fuel, which could be used on Mars or as a safety measure for future Mars missions. A detailed description and justification for pre-positioning is presented in Section 6.4.3.1.4.
Based on the assumption of pre-positioning mass in Earth and the destination orbit, the following modules are pre-positioned in Martian orbit and correspondingly for Lunar orbit for the Moon mission:
1. Service Module #2 (SM2) – As an assumed function, this module must be capable of providing propulsion for transiting the crew from Martian (or lunar) orbit to Earth orbit.
2.
Two Lunar or Mars Landers (ML1 & ML2
or LL1 & LL2) – Two identical Apollo type Landers (slightly different
forms for Moon and Mars), each capable of transporting three crewmembers from
orbit to the surface. Functionally, only
minimal redesign is required for Lunar landing capabilities (see Section
6.4.3.1.6.). As such, the Moon can be a
“true testbed” for the Lander technology.
An option also exists to re-use the Landers for the Moon mission because
there is no heat shield requirement for a lunar landing (see Section
6.4.3.1.6.) Both
Landers must have propulsive capabilities in addition to the requirements for
pre-positioning. This extra propulsion
redundancy allows either the Lander to dock with the HM or the HM to dock with
the Lander in Mars orbit (or Lunar orbit).
Although there is increased mass associated with two three crew Landers
as compared to one six crew Lander,
·
This technology is modular and extensible to a
Mars mission (6 crewmembers) from a Moon mission (3 crewmembers).
·
Increased reliability and flexibility observed
because if one Lander were deemed unusable, the second Lander could still
perform the desired mission objectives.
A detailed trade study that discusses the above can be found in Section 6.4.3.1.6.2.
Similarly, pre-positioning all of the necessary cargo for Mars and Moon surface habitation, on the surface, in a separate mission, reduces the payload mass that must be launched from Earth or LEO. For surface habitation, if the environments were deemed similar in design requirements, an option may be to use a duplicate of the habitation module on the surface of Mars. Details of the surface habitation modules are given in Appendix 9.2.
The following items will also be pre-positioned in LEO.
1.
Habitation Module (HM) – This module
will be launched in pieces and assembled in LEO, allowing the overall volume to
not be limited by the minimum launch volume requirements. Since there are many modules, none, one or
many can be used for each mission, depending on the mission requirements for
duration and crew size. This module must
have propulsion capabilities to perform docking maneuvers in both Earth and
Mars (or Lunar) orbit.
2.
Service Module #1 (SM1) – This module
must be capable of providing propulsion for transiting the crew from Earth
orbit to Martian orbit or Lunar Orbit (Note that SM2 was used for (Mars or
Moon) to Earth transit). This module
will require minimal redesign to be truly extended for use on a Mars mission
(the same interface and platform could be used).
Since no crews are launched into LEO during the pre-positioning phase, all launches can take place on non-man-rated launchers. This serves to increase launch flexibility and lower overall cost.
By using a pre-positioning approach, only a single man-rated launch is required to deliver the six crewmembers (or three crewmembers for smaller Moon missions) to the Habitation Module. The launched modules include:
1.
Modern Command Module (MCM) – The Modern
Command Module is discussed in Section 6.4.3.1.1.1. This module will dock with the Habitation
Module and return crew back to Earth from LEO at the end of the mission. This
module is also capable of docking and delivering the crew of three to the
Habitation Module (HM). Note that this
module will not travel to the destination with the Habitation module, but only
serves as a means of transporting the crew to and from Earth (first and final
phases of the mission). Since some of
the defined Moon missions specify a crew of three instead of a crew of 6, a
three-crewmember transfer vehicle was chosen to reduce the mass launched on a
human rated launcher for the three crewmember Moon missions. This was also done to ensure that the Earth
to LEO and LEO to Earth transfer vehicle was not over designed for a
three-crewmember Moon mission. This can
be done because the mission scales via the number of Habitation modules (see
Section 6.4.3.1.3.) used and not the size of the LEO crew transport
vehicle. A detailed crew vehicle scaling
analysis and justification for this decision are presented in Appendix 9.1.
2.
Crew Operations Vehicle (COV) – This module
docks and delivers the remaining three crewmembers to the Habitation Module in
the same manner as MCM. However, this
module does not remain in LEO during the mission. Instead, this module travels with or without
the HM to the destination. Functionally,
this vehicle must provide all the necessary functions of a crew transport
module (i.e. docking capabilities, GNC, radiation protection, thermal control,
ECLSS, attitude control, communication equipment, etc.). The reasons for choosing an additional module
that travels with the HM are twofold:
·
Having a COV
separates the functions for crew transport, allowing the mission to scale down
to transporting a crew in only the COV to the Moon for a Short Moon mission, to
scaling up to transporting a crew in the COV & HM to Mars for the Extended
Mars mission.
·
Having the
smallest mass and volume re-entering Earth at the end of the mission is
beneficial to reducing the mass penalty of the mission. This is further described in Section
6.4.3.1.1.1.
All modules considered common interfaces and serve specific functions. Details of the launch logistics, timing and forms for both the human rated and non-human rated launches are presented in Section 6.4.2.2.
Now that all of the modules have been pre-positioned, SM1 docks with the HM in Earth orbit. To perform this docking, it is necessary that SM1 and HM both have propulsion capabilities for added redundancy. After this docking is complete, the COV docks with the HM and SM1. These three modules comprise the entire crewed vehicle that travel to Mars or the Moon, known as the Moon/Mars Transfer Vehicle (MTV) (see Figure 29).
Figure 29:
Mars/Moon Transfer Vehicle (MTV)
The MTV (COV/HM/SM1) attains Martian orbit by aerocapture. This requires the use of the COV heat shield and additional protection for the HM in contact with the atmosphere. It is assumed that the shielding could be modular (i.e. detachable) to ensure HM extensibility remains intact. Prior to entering Martian or Lunar orbit, SM1 undocks from the MTV, reducing the aerocapture mass (at Mars) or the mass of fuel required to enter Lunar Orbit.
Two pre-positioned Landers dock sequentially (ML1 & ML2 or LL1 & LL2), requiring the HM to have only one docking interface for the Landers. Three crewmembers transfer into each Lander, and then descend to the surface. In a similar manner to the Apollo Missions, the Landers ascend to Mars/Lunar orbit. Details of the Lander are presented in Section 6.4.3.1.6.
Once the Landers dock sequentially with the MTV, a second Service Module (SM2) consisting of the necessary propulsion to return the crew to Earth docks with the MTV in orbit. These three modules (COV/HM/SM2) comprise the entire vehicle that travels back to Earth. The vehicle is identical to the MTV that transited from Earth to Mars (or Earth to Moon) with the exception of SM2 now being used in place of SM1. Note that the Landers do not return to Earth with the crew for the Mars mission, but could possibly be reused for the Moon mission.
The MTV attains Earth orbit by aerocapture. For the Mars mission, the same MTV heat shield used for aerocapture at Mars is used at Earth. While in Earth orbit, the Modern Command Module (MCM) docks with the HM and the crews of three enter their respective modules (COV or MCM). MCM and COV undock from HM and reenter. If the MTV were capable of direct entry, an aerocapture maneuver would be eliminated and an argument could be made for MTV re-usability. However, at this stage it is difficult to foresee whether the entire Habitation Module could handle the heat load of re-entry, especially after experiencing the heat load associated with aerocapture at both Mars and Earth. For a Moon mission an argument could be made for MTV reusability because aerocapture is not required to enter lunar orbit.
NASA’s current direction is return to exploring the Moon, then explore Mars and beyond. It is with this in mind that we are proposing a high-level set of causally connected baseline missions. These missions are aimed at a continued expansion outward into the solar system, with no single ultimate destination. The experience and knowledge gained from one mission will be put to use on the following missions, thereby enabling future exploration. To recognize this fact is to understand that commonality must play a critical role in all extensible space exploration system. It is by searching for commonality between aspects in the forms and functions of these missions that they may be integrated to generate one consistent over-arching program plan.
Each of the proposed missions for the Moon and Mars contains a number of requirements. Clearly, the requirements of each module (or form) vary from mission to mission, but the objective of an extensible set of mission architectures is to utilize as many functions as is feasible for each form. Deciding the “as is feasible” is difficult, and as such, the following method of comparison aims to illustrate the functional requirements of each form that are not clearly demonstrated in the baseline mission.
The requirements were specified for three Moon missions: Short, Medium, and Extended. These were also specified for the four Mars Missions: Phobos, Short, Extended, and Extended+. Basic forms were selected and the functions were discussed for each form. Each mission was considered independently in the analysis below. For example, if an Extended+ Mars mission requires that the COV be equipped with an aeroshield, while the other three Mars missions do not, the Mars set of missions are assumed to require an aeroshield, however these situations are indicated in the table located below the Venn diagrams in the following figures.
This method of analysis allows functional traits of a form to be easily evaluated and compared with the other required functions in terms of importance across the entire set of mission objectives. If a form does cannot perform a specific function, a decision must be made as to whether or not extending the functionality of a form to include other functions is justified or whether an additional form should be developed to serve the functional requirements flow down from the Mars and Moon mission objectives.
The following sections will present an analysis of the form/function matrix by means of Venn diagrams. This is a helpful tool in deciding what functions each form should be able to perform, depending on the needs for specific Moon/Mars missions.
Although the shared portion of the Venn diagram is critical to both Mars and Moon missions, it is the regions on the right and left of the Venn diagram that impact individual missions within the respective Mars and Moon mission frameworks. Therefore, considering each of the functions in these areas of the Venn diagram allows the mission designer to decide when a new form capable of accomplishing the functions not included in the “overlap” region of the Venn diagram should be added to the network of modules.
The Crew Operations Vehicle (COV) is part of the crew exploration system. It is designed for transporting a crew of three from Earth to the destination and back, with or without the help of the Habitable Module (HM). Specific details of the COV were discussed in Section 4.1.1, but from the Form/Function Matrix results (see Appendix 9.2.1) a Venn diagram of functional requirements for the COV was determined and shown in Figure 30.
Figure 30: Functional requirements for a Crew Operations Vehicle
Considering Figure 30, the COV must be capable of providing attitude control and communications for a crew of three, as well as additional human space flight requirements for both the Moon and Mars missions. Based on this diagram, additional forms must be created in addition to the base form to accomplish the required additional functions.
Since it is beneficial to have a minimum number of forms in addition to an extensible network of forms, it is important that some of the functions required of a COV be captured by the baseline COV. For example, docking with the HM reduces the number of functions required of a COV because the HM can provide some of the functions listed for the COV. Since the functions listed in Figure 30 include the unique, additionally complex Extended+ Mars mission, it is difficult to truly evaluate how extensible the COV should be. For example, the Extended+ mission is the only mission that requires the COV be capable of ascending and descending to the surface of Mars. This single functional requirement is essential for only one of the seven missions. Furthermore, the particular mission that requires this added ability is well in the future. Thus, not only is the mission less likely to be carried out as planned, but the expected knowledge returns from this mission would diminish considerably upon discounting to present value. Considering the environment of uncertainty surrounding activities that are so far removed in time, this function should not be captured by the COV unless it could be added at low cost. An additional or alternative crew vehicle form might suffice in the future.
For the Short Moon missions, the COV is required to support one person in orbit. This extends the period of time that the COV must provide life support. For the Extended Moon mission, the COV is required to sustain unmanned orbit. This situation indicates that either an additional form should be added to the network or the additional functions be adopted by the COV. This important decision is made easier when all of the low-level functions can be considered together. The COV will be required to re-enter Earth ballistically for the Short and Medium Moon missions. This delays the requirement of Earth aerocapture technology until the Extended Moon mission, allowing the Moon missions to be a “true testbed.”
The Modern Command Module (MCM) is the form specified for the Earth launch module in the initial phase of the mission and Earth EDL in the final phase of the mission. The MCM resembles an Apollo-style command module, capable of transporting crew from the Earth’s surface to orbit and back to Earth from orbit, but without the required capability of transit to the Moon. Specific details of this module were discussed in Section 4.4.2.2.2, but from the Form/Function Matrix results (see Appendix 9.2.1) a Venn diagram of functional requirements for the MCM was determined and shown in Figure 31.
Figure 31: Functional requirements for a Modern Command Module
It is evident that the requirements for Earth launch are similar for all of the missions and independent of the mission objective. This indicates that the MCM is an extensible architectural element for these functions. The only specific requirements that differ between the Mars and Moon missions are the number of days of life support and the number of crew that must be transported. The number of days of life support was estimated based on the expected assembly time in orbit. Providing additional life support for the longest possible duration in the MCM is not a difficult function to incorporate in the baseline MCM. Rather than developing two functionally equivalent modules such as a three-person and a six-person module, many benefits exist when two identical forms (COV is a similar to the MCM) are used in place of a different form to perform the same list of functions. This premise was discussed in Section 6.4.3.1.1.1.
A Habitation Module (HM), as the name indicates, is a module that supports human life on long duration missions. While the transit durations for the Moon missions are short enough to not require a Habitation Module, the module could still be used to test the technology for future missions to Mars. Specific details of the HM were discussed in Section 4.4.2.2.1, but from the Form/Function Matrix results (see Appendix 9.2.1), descriptions of its desired functions are given in Figure 32.
Figure 32: Functional requirements for a Habitation Module
An interesting observation is the number of different forms that must have docking capabilities with the HM. Since the HM module will be docking with the COV, SM1, SM2, and MCM at separate times during the missions, the design of all these elements should require an identical interface. Such a similarity allows for more flexibility and adaptability for new strategic decisions. For example, if the timeline of a mission changes and, for example, docking with SM2 occurs before docking with the Landers, the system can still function adequately during logistic constraints because of the common interface. An option also exists to use a portion of the modular HM to support a crew of 6 on a Moon mission. This allows the functional requirements of a COV supporting a crew of 6 to be removed. For the Moon mission, an engine burn is required to attain lunar orbit. However, a heat shield is still required for the HM because an aerocapture maneuver will take place when the HM re-enters LEO on the Extended Moon mission.
It is important to mention that the HM has significantly different life support requirements for the Extended and Extended+ missions compared to the Short and Phobos missions. This is a result of different trajectories being used for those missions.
The Crew Service Module (SM) is the form that provides fuel to transport the crew traveling in either the COV for a Moon mission or MTV (COV/HM) for a Mars mission. This module could be functionally compared to the Apollo Service Module. Specific details of this module were presented in Section 4.4.2.2.1, but from the Form/Function Matrix results (see Appendix 9.2.1) a Venn diagram of functional requirements for the SM was determined and shown in Figure 33.
Figure 33: Functional requirements for a Crew Service Module
It is evident that the pre-positioning requirement given by a Mars mission is not a critical technology for a Moon mission; however, the Moon could use pre-positioning to test the technology. As such, two similar versions of an SM could be designed, one designed for pre-positioning use and the other for conventional propulsion. The first Moon mission would not require the technology and would thus use the conventional propulsion Service Module. However, later Moon missions could use pre-positioning technology using electric propulsion as a stepping-stone for the future Mars mission requirements. Furthermore, the amount of fuel required for a trans-lunar injection is significantly different than the amount of fuel required for a trans-Martian injection. Taking this into consideration may require that the Moon SM and the Mars SM have slightly different forms. The extensibility value of using similar structures is in the manufacturing savings resulting in the use of legacy hardware.
A Mars Lander (ML) or Lunar Lander (LL) share similar functionality. This similarity in functionality is exploited to incorporate commonality among the various Lander designs. Specific details of the Landers are discussed in Section 6.4.3.1.6. The Lander was based on an Apollo style Lander, capable of transporting a crew of three from orbit to the surface and back to orbit. From the Form/Function Matrix results (Appendix 9.2.1) a Venn diagram of functional requirements for the Lander was determined and shown in Figure 34.
Figure 34: Functional requirements for a Moon/Mars Lander
Consider Figure 34. For comparison purposes, it is clear that common functions are shared. When common functions exist, extensibility will benefit the overall group of missions to the Moon and Mars. Considering Figure 34, the Lander must dock with the COV or the COV/HM in both lunar and Martian orbit. As well, the Lander must deliver a crew of 6 to the surface for all of the Mars missions and some of the Moon missions. If two identical Landers are chosen instead of a single, larger Lander, the impact of this decision can be observed by evaluating whether or not the new option satisfies the functional requirements. If all of the functions are deemed satisfied, only then was the impact of the decision not critical. As can be expected, a wide range of requirements are made for the Landers, but many of these requirements are specified by only one of the seven missions, making it difficult to justify changing the baseline form. Indeed, the Landers are a mission critical piece of hardware and must be highly reliable. Therefore, when considering extensibility of such a device, it may be beneficial to target the Lander design for the most difficult landing mission, thereby ensuring a robust, if over-designed, form for the other missions. This has the effect of increasing net reliability while still maintaining an extensible form. The idea of designing a non-optimal form now such that it may be optimal when used in a different manner or location stands as one of the cornerstones of extensibility.
When this method of comparison of using Venn diagrams and form/function matrices was developed, it became apparent that many functions were required of each form. At this stage, the Venn diagrams do not capture the extreme detail required of an extensible Moon/Mars mission architecture. However, this technique provides a tool for designing an extensible transportation system. Although many simplifying assumptions were required to analyze a transportation system in this framework, the advantage of this method is seen when considering the impacts of a decision to not have a form perform a certain function. This method allows these decisions to be traced. The diagram highlights the functions that were not captured by the baseline mission architecture. Thus, pains must be taken to integrate these functionalities into the eventual design for the missions that require them. In doing so, it is important to recognize a fundamental engineering tension that exists between optimality and extensibility. A form that is designed purely with optimality in mind is restricted to the point design for which it was originally conceived. This makes the creative use and extension of such technology difficult. On the other hand, a form designed with only extensibility in mind will become “spread thin,” and unable to perform the functions required of it at certain stages or missions. Thus, a compromise must be made between these two extremes.
When designing for extensibility to missions involving high degrees of uncertainty, care must be taken to ensure that the current mission does not become so overburdened with extraneous requirements that it is prohibitively expensive to function as planned. For a “point design” system, it is a possibility that some functionality designed into the system may never be used. In addition, this “point design” system may need to be entirely re-designed before future missions take place. However, when designing for extensibility for near-term missions, the addition of an extra function now which will likely be needed in future missions may decrease the cost of developing and testing that functionality in the future, thus enabling further exploration in the long term and increasing operational knowledge.
The Presidential Directive on the Moon, Mars
and Beyond clearly sets the Moon as the first goal for today’s astronauts. The
Moon is intended to be used as a testing ground for missions to Mars and
Beyond. In the absence of existing infrastructure on the Moon, SSLMs may be
expected to occur in the near future. The primary purpose of these missions is
to serve as lunar “scouts,” which will search primarily for location information,
perhaps regarding possible resources that may be exploited on the Moon. They
possess the positive attributes of being relatively low in cost and have the
potential for high knowledge return given that they land in unexplored
locations. Prior to the first SSLM, unmanned robotic probes may be sent to a
number of promising landing sites. Similarly, lunar satellites could be
constructed to allow for continuous radio contact with the far side of the
Moon, if deemed necessary for communications with these probes. SSLMs will occur, each time in a different
location, until a decision is made to study particular locations in more depth
and for a longer period of time. Future Short Moon missions may include testing
of preliminary rover hardware, testing of new space-suit concepts, and gradual
extensions of the life-support capabilities of possibly up to 1-2 weeks. Once
sufficient experience with lunar operations has been established through the
SSLMs, MSLM missions may be launched.
Figure 35: Flow diagram describing elements of extensibility in integrated baseline
The primary purpose of MSLMs is to generate
scientific knowledge and to establish non-permanent infrastructure on the Moon.
These missions will be aimed at scientific exploration and resource evaluation
of promising sites found during the Short Stay Lunar Missions. These missions
are larger in scale than the SSLMs, and they possess the ability to carry more
equipment. As such, astronauts participating
in these missions will be tasked with operating larger scale scientific
apparatus and the scientific precursors to the first in-situ propellant
production facilities. These will be small scale at first, serving primarily as
technology demonstrators, but may be scaled up in future missions to allow for
some basic functionality. Included with Medium Moon missions will be an
unpressurized rover, which is designed to carry astronauts to locations beyond
their operational walking radius. These rovers may initially be tested on
SSLMs, and will probably initially be restricted to traveling distances that
astronauts can safely walk back from.
With sufficient experience gained by using this rover system, the range
of the rovers may be extended. MSLMs
will occur in a limited range of locations and multiple missions may be sent to
the same location. These missions will occur until a primary site is chosen for
a future semi-permanent lunar base. MSLMs are still limited by the fact that
astronauts may only stay on the lunar surface for limited periods of time and
the fact that most mass utilized will come from Earth. In general, these
missions will not have the capability to survive throughout the lunar night,
although later MSLMs may wish to gear some activities towards this kind of
sustenance.
The next step
beyond the MSLMs is dependant on the degree to which in-situ lunar resources
may be used. If these resources are present, it may be desirable to conduct a
series of Extended Stay Lunar Missions (ESLMs) designed to set up a
semi-permanent lunar base and to generate the capability to search the far side
of the Moon in more detail. This mission would be primarily aimed at allowing
more humans to live on the lunar surface for increasingly long amounts of time.
This report assumes that such resources are present and usable. In the event
that in-situ resource production is not directly feasible, such a long-term
base on the Moon would require a significant supply-chain from the Earth. The
requirement that this supply chain imposes is not directly in line with overall
mission sustainability. Thus, if in-situ resources are not available, the ESLMs
may not be launched. Instead, more MSLMs may be carried out to prepare for an
eventual Martian Short Stay Mission.
The primary purpose of ESLMs is for this architectural
design to align with the President’s declaration that more humans will remain
on the Moon for increasing lengths of time. These missions will be aimed at
establishing a semi-permanent habitat with at least six astronauts on the lunar
surface where astronauts may gain experience in living in non-Earth
environments for long periods of time. Since these missions will be aimed at
the construction of a habitat, some pre-positioning will take place by
necessity through resupply from unmanned probes and the use of cargo from
previous Moon missions. Therefore, this suite of missions will allow for the
buildup of accurate pre-positioning operational knowledge. If the targeting is
not accurate on the first few attempts, rovers from the MSLMs will be available
to allow the astronauts to reach the landing site. In the worst-case scenario,
they may return home as is done in a MSLM. Astronauts participating in these
missions will be tasked with operating larger-scale in-situ propellant
production facilities. The eventual goal will be to create a largely
self-sustained semi-permanent base. Some potential capabilities to be added
include a rescue vehicle stored in an easily accessible location so that
astronauts may escape to Earth in case of unforeseen circumstances. Included
with ESLMs will be habitat modules in which astronauts may live for increasing
lengths of time. ESLMs will initially occur only in one location. These missions will
likely occur until the first Mars mission is launched, at which point the semi-permanent
base may be turned over to international partners or the commercial sector for
further development. These missions will
have the capability to survive through the lunar night, and will therefore
require an independent source of power to be used during the lunar night. The
next step beyond the ESLMs is the Martian Short Stay set of missions. These
will be the first time humans will travel beyond Earth’s gravitational sphere
of influence with respect to the Earth-Moon system.
The primary purpose of the Martian Short
Stay is to demonstrate the ability of mankind to survive on the surface of
Mars. These missions nominally require
pre-positioning of cargo on the Martian surface, although the first such mission
may not utilize this capability simply because of the complexity of
pre-positioning maneuvers. If pre-positioning has already been successfully
tested during Moon missions, the technology used may be used on missions to
Mars. Prior to the first Short Stay mission, unmanned robotic probes may be
sent to a number of promising landing sites. These probes may also be used to
practice the accuracy of pre-positioning technology. Like an ESLM, Mars
missions utilize habitat modules. In the event that no in-situ resources exist
to be exploited on the Moon, the habitat capability must be developed for this
mission with no in-situ resources utilization knowledge to be extended from a
previous mission. Like the MSLM, Short Stay missions will possess unpressurized
rovers that may be used to explore over a relatively large range. They will
also be used to test and verify in-situ resource production and utilization
facilities for use on Mars.
Short Stay
missions will occur, usually in different locations, until a decision is made
to study particular locations in more depth and for a longer period of time.
Although Short Stay missions will likely occur more than once, the mass and
energy requirements to perform these missions will moderate the number of Short
Stay missions compared to the number of Short Stay Lunar Missions. Rather, upon finding an ideal long-term site,
Martian exploration may continue with longer term, shorter-transfer missions to
minimize the effect of microgravity and to take advantage of the resources
expected to be found on Mars. The major limitations of the Short Mars missions
are imposed by the mass required to support life on Mars and in transit for
such a long time. Any opportunity to
reduce mass discovered during lunar mission operations would likely be
implemented in this mission. For
example, if in-situ propellant could be produced on the Moon, it could
significantly reduce the mass required to get to Mars, even if that propellant
required a detour to the Moon.
Similarly, weight concerns impose an upper limit on the number of samples
that may be brought back to Earth.
Future Short Moon missions may include testing of longer-term habitation
facilities, testing of new space-suit concepts, and alternative propulsion and
in-situ propellant production concepts. Once sufficient experience with Martian
operations has been established, Extended Stay, and Extended Stay +
Infrastructure missions may occur.
The primary purpose of Extended Stay and Extended Stay + Infrastructure missions is to demonstrate the ability of mankind to survive on the surface of Mars for an increased duration. In the case of Extended Stay + Infrastructure missions, humans will establish a semi-permanent infrastructure on Mars to be used for science, operations research, or as a testbed for the next destination. These missions nominally require pre-positioning of cargo on the Martian surface, and therefore require the performance of a successful Short Stay Mission to ensure that pre-positioning technology is adequately developed. Prior to the first Extended Stay mission, Short Stay missions will have identified promising resource excavation sites, and resource processing activities. One of the purposes of an Extended Stay mission is to take advantage of these capabilities for refueling, for life-support and agriculture, and to explore Mars in a more comprehensive manner over a longer period of time than is possible with a Short Stay mission. An Extended Stay + Infrastructure mission will have the capability to be self-sustained based upon in-situ resource production, thus reducing mass in LEO as much as possible. To this end, the transit characteristics of an Extended Stay + Infrastructure mission will evolve from a MOR class mission to a Mars Direct mission. If in-situ production is not available, Extended Stay + Infrastructure mission will likely not occur because of the difficulties in maintaining a Martian supply chain from Earth. Like an ESLM mission, all Mars missions utilize surface habitation modules. Extended Stay and Extended Stay + Infrastructure missions will take advantage of knowledge gained from habitat technology used for the Short Stay missions. Extended Stay and Extended Stay + Infrastructure missions will possess upgraded rovers that may be used to explore a larger surface area. These rovers may be pressurized in the Extended Stay missions and will almost certainly be pressurized in the Extended Stay + Infrastructure missions. This represents extensibility from the ESLM. Extended Stay + Infrastructure missions will make use of inflatable structures and other innovative semi-permanent construction materials in the establishment of a Martian base. Extended Stay and Extended Stay+ missions will occur confined to valuable locations. These locations will often be dictated by those that are richest in exploitable resources and those that promise to yield the greatest knowledge returns.
Upon successful completion of Extended+ missions, NASA will have gained significant experience in the area of manned space exploration beyond LEO. This experience will help further exploration throughout the solar system. A mission to Phobos is just one way to start this expansion. The primary purpose of Phobos missions is to demonstrate extensibility on multiple levels. Beyond the Martian system, there are three places that NASA may choose to explore. These include:
For reasons previously mentioned, a pre-Short Stay mission to Phobos is an ideal technology demonstrator for the destinations listed above. Phobos missions may be conducted at any time within the baseline framework. The suggested time for the mission is before Short Stay Mars missions, when resources are not yet heavily invested in a semi-permanent Martian base. This will give NASA operation experience in navigating to Mars orbit, similar to how Apollo 8 achieved lunar orbit before the landing of Apollo 11.
Although this report only focuses on the Moon and on Mars as locations to be explored, the new NASA vision unequivocally states that the program does not end there. Exploration is an activity that will never cease and the potential to educate and to inspire from this exploration will never run dry. Although the next location to be visited by human astronauts is not certain, there is no shortage of secrets to be unlocked and mysteries to be discovered. It is for this reason that the President’s vision is entitled Moon, Mars and Beyond, for Beyond is the ultimate destination of mankind.
The ultimate goal of the design process is
to create an architecture that is flexible and robust in the face of change.
Having identified sources of commonality between the Moon and Mars, and
translated this commonality to operational and formal attributes, the next step
is to create a flexible architecture.
Final decisions regarding the architecture will depend on key trades
identified during commonality mapping.
Because we hope to create an architecture using a long-term view of the
system life-cycle, however, tools that capture the value of flexibility and
robustness will be needed.
This section presents three possible tools
for such analysis:
One defining
attribute of a sustainable system is a long expected life-cycle. Thus, predicting the circumstances under
which the system will operate throughout its life cycle becomes difficult as
uncertainty increases with time. Such systems must also incorporate subsystems
that operate throughout the architectural domain, from the mechanical to the
political, to the commercial. Again, significant uncertainty is present in the
system’s operating environment. As a result, the system must be prepared to
adapt to unexpected situations without significantly reducing the system’s
operational utility. Space systems must
balance budget constraints and risk. The difficulty of maintaining a delicate
balance has resulted in a low mission success rates.
This section
suggests a way to choose a flexible architecture that will adapt to different
scenarios, thereby helping the system to accommodate to changing environmental
conditions without significantly compromising performance. Two approaches are
available; either a closed “best design” that attempts to take into account
each and every possible change, or a strategy that will change and evolve to accommodate
the unforeseen. The former option is restricted to current projections of
future events, whereas the latter option is dynamic and adaptable in the face
of uncertainty. This report proposes the latter approach as a way to adapt to
changing environmental conditions. That
is, a baseline option is chosen now, while preserving the widest possible
options for the future. Decisions, which would otherwise have to be made at the
outset, are delayed such that, when the final choice is required, it is made in
an environment of decreased uncertainty.
The bulk of this
report focuses on the creation of a baseline strategy to go to the Moon, to
Mars and beyond. This baseline was designed with a set of implicit assumptions,
regarding the state of the system’s operating environment through time.
Ideally, the baseline is the strategy that is most likely to succeed given
present knowledge of future events.
An alternative set
of extreme environmental changes that would impact the baseline design (either
positively or negatively) was identified through brainstorming. These changes constitute scenario
descriptions. A set of responses to those scenario descriptions was then
developed. These responses constitute alternative paths that may be taken in
designing the system. Using this framework, one may also think of the baseline
design as a response to the most likely scenario. The scenarios also provide a means to
identify a possible set of architectural trades or options. These critical trade decisions are analyzed further.
In doing this selection, the amplitude of the field to be explored was severely
cropped in an effort to perform a somewhat deeper analysis on these interesting
features of the trade space. If all of these decisions were to be made at the
outset, based only on current understanding of future events, there is a high
likelihood that these decisions would not be the best choices in the future.
There is almost always more information that becomes available through time
that has the effect of changing the environment under which the system must
operate. Thus, it is beneficial to delay decisions to allow for a better, more
informed decision to be made later.
All potential
architectures can be fully described by a vector of the different decisions
that have to be taken in order to implement it. The meaning of the word
architecture in this case is not restricted to the physical form of the
objects. A space architecture includes
forms, transit points, budget, policy, etc. It is possible to apply Utility
Theory to analyze each of these vectors in the context of scenarios and
associated trades.
In doing so, one
must first explore how the present baseline reacts to a change in the operating
environment, and how appropriate decisions taken at points throughout the
system’s life-cycle could buy some insurance against negative scenarios, or
increased payoffs in case of positives ones. An investment increase will reduce
the expected utility, while potentially reducing risk by neutralizing negative
outcomes. Similarly, one could increase the payoffs by taking real advantage of
optimistic outcomes. This analysis is based on Real Options theory. The problem
of acquiring utility is complex; however, and in its complexity lays its value.
The utility is not an absolute number; rather it is a scaled numerical
representation that reflects a synthesis of the opinions of a group of
stakeholders around an issue. The tool proposed to make this synthesis is
called the Analytical Deliberative Process. This tool is a formal framework
that helps a group of people to discuss a set of opposing measures and to
synthesize their potentially contradictory opinions. In doing so, utility
values may be determined such that the analysis can proceed. However, one
important point should be clarified. This utility number is an artificial
creation. It is not a “silver bullet”. Rather, it expresses a set of opinions,
many of which may be subject to change through time. Computers are often used
to calculate these values based upon subjective input, and it is therefore a
temptation to treat these utilities as incontrovertible data, although they are
subjective assessments of individuals’ opinions.
By assessing the
different performance metrics that each of the architecture vectors presents,
it would be possible to get a measure of the net utility that each architecture
holds for the stakeholders. In order to select the architecture that will
provide the highest expected utility, a method using Decision Analysis is
proposed.
As time passes,
some decision points are encountered. Similarly, there are points in time in
which the architecture’s operation depends on some element of chance. Some of
these chance points may have been predicted through previous analyses, and thus
the baseline would remain unchanged.
Still, others may not have been anticipated and an out-of-baseline
approach might be necessary. It is crucial to understand each of these possible
branching points, and to study them through a conceptual scenario analysis,
such as the one described above, in order to be able to apply the real options
approach. The approach provides a
thoughtful way to decide where, when, and how much to pay for risk reduction
and benefit magnification.
In order to
proceed, three tools are proposed:
Both tools help to
understand how to deal with a complex set of options and requirements, which are
sometimes in conflict.
This is a rigorous and replicable method that provides a protocol, under
which a community of “experts” may arrive at an answer to a factual question.
This concept of a consultative body of experts, which assigns these utility
values, is important, since it allows for the incorporation of viewpoints
ranging from members of the technical community to members of the political
community to the public at large.
The process uses the following steps:
1. Decompose the problem into a hierarchy:
All the different
requirements have to be sorted into a hierarchy.
In order to calculate the value of a certain
approach, several higher-level Impact categories are devised. They may or may
not be independent, in the sense that a change in any of them may affect
others. Each of these impact categories includes several objectives that use
performance metrics in order to be measured. Some of these performance metrics
will be formulas, and will have exact values, whereas others will be more
subjective, and be measured in ranges. In each case, a number of utility levels
are chosen. This yields a tree-like structure in which the main trunk depicts the
utility value.
2. Experts rate the importance of the different
levels and state their preferences. They are asked to make pairwise comparisons, starting from the utility levels on the lower
branches of the tree.
3. In order to synthesize the results, these
previous pairwise comparisons are used to fill the so-called Ranking matrices.
The eigenvalues of these matrices are the weights that the experts have chosen
for the different branches that feed into the utility values.
4. To evaluate the
consistency of a judgment, a confidence index is calculated. This index assesses
the coherency with which the experts have made their judgments. This index
helps to build confidence on the coherence of the opinions expressed.
The same approach is used to generate
weights for the different branches of the tree: performance measures feed an
objective, objectives feed an impact category and, finally, impact categories
feed into the utility value.
Using these methods, experts may determine how to rate the different
approaches. Utility values may be calculated using some metric, such as mass to
be launched to LEO, or a subjective measure, such as the desired level of
extensibility through a constructed range scale.
Decision Analysis Theory is a structured way to rank the decision
options available to the decision maker. It enumerates the immediate and later
choices available, characterizes uncertainty, quantifies their desirability,
and provides rules to help the decision maker to choose the “best” alternative.
Choices to be made, and chances to be taken are organized in a tree-like
hierarchy, with the immediate choice at the trunk of the tree, and the
posterior ones following in some sort of order (e.g. chronologically).
At each position where a chance event is encountered, there is a chance
node, and at each position where a decision has to be taken, a decision node is
assigned.
p=0.1 p=0.4 Alternative B Alternative A Utilities
Figure 36:
Decision analysis tree.
Decision nodes are drawn as
squares, and chance nodes are drawn as circles.
In this example, the decision maker will choose Alternative A because it
expected utility is 0.72, which is higher than the utility of Alternative B
(0.22).
In order to evaluate the desirability of each option, a certain utility
is assigned to each final outcome of the tree as shown in Figure 36. This utility function could be an absolute number,
but the previously explained Analytic Deliberative process can be used to
assign these utilities. Similarly, at each chance point, probabilities are
assigned to each possible outcome. These probabilities could either be assigned
by a complete assessment of the probabilities, or by asking experts about the
likelihood of the different events and applying the Analytic Deliberative
process once again. Taking into account these utilities and probabilities, it
is possible to travel backwards in the tree, and thus to calculate expected
utilities at each chance node.
A short example follows, using a decision node that is defined by the
option of using
EM-L1 as a transit
point on the way to the Moon. This example also considers the probability of
having water on the lunar South Pole. The utility of including the fuel necessary
to transit through EM-L1, and then to construct a lunar base, is the
calculated. The numbers and opinions expressed in this example are not rigorous.
Rather, they have been used to show how a method of this kind could be used,
even when exhaustive analysis and complete studies of the involved
probabilities have not been done. The primary purpose of this example is to
show the mechanics involved in the decision-making process.
The following impact categories have been arbitrarily chosen:
In order to measure the above impacts, the following objectives have to
be fulfilled
·
Cost
o
P1 = The utility gained from reducing mass,
o
P2 = The utility gained from reducing schedule time
·
Schedule fulfillment
o
P3 = The utility gained from launches going
according to schedule
·
Longer term projections
o
P4 = Probability of maintaining a permanent base,
o
P5 = Probability of producing and shipping fuel
Using the AH process matrix, and pairwise comparisons, experts have
arrived at the following conclusions:
P1 Minimum weight
A LEO weight between 40MT to 60MT is most desirable with a utility of
0.5
A LEO weight between 60MT to 250MT is less desirable with a utility of
0.4
A LEO weight between 150MT to 600MT is least desirable with a utility of
0.1
P2 Shorter Schedule
A schedule between 3 and 8 years is the most desirable with a utility of
0.75
A schedule longer than 8 years is less desirable with a utility of 0.25
P3 Schedule fulfillment
95% of launches on time is most desirable with a utility of 0.6
50% of launches on time is less desirable with a utility of 0.3
Less 50% of launches on time is least desirable with a utility of 0.1
P4 Probability of maintaining a permanent base
An 80% probability of maintaining a permanent base has a utility of 0.8
Less than 80% probability of maintaining a permanent base has a utility
of 0.2
P5 Probability of producing and shipping fuel
An 80% probability of producing and shipping fuel has a utility of 0.85
Less than 80% probability of producing and shipping fuel has a utility
of 0.15
Using the same method, experts then decide that the weights between
factors P1 and P2 should get a utility for the cost. These are 0.7 and 0.3
respectively. Similarly they have found that their AH process gives a weight of
0.8 to P4 and 0.2 to P5
Therefore we have now
Cost utility = 0.7 P1 + 0.3 P2
Schedule fulfillment utility = P3
Longer term projections utility = 0.8 P4 + 0.2 P5
Next, experts are asked again to evaluate their pairwise preferences
between these three impact categories. They return values that result in the
following weights:
The following value formula may therefore be derived:
Total Utility = 0.5 x (0.7 P1 + 0.3 P2) +
0.2 x P3 + 0.3 x (0.8 P4 + 0.2 P5)
Total Utility = 0.35 P1 + 0.15 P2 + 0.20 P3
+ 0.24 P4 + 0.06 P5
It is important to note that while it would be possible to obtain firm
data for some of these performance measures, the lack of unequivocal data does
not hinder the analysis. In any case, the result will be a mix of opinions and
engineering calculus that reflect the experts’ best knowledge of the situation.
This concludes the Analytic-Deliberative Process.
To analyze the Decision Analysis Tree, one must enumerate the chances
and events that will occur, in chronological order:
First Event
Decision: Should an architecture to travel to the Moon include L1 as a
transit point? This will require an extra 10% of propulsion capability that
will not be used in case the mission is not directed to pass through at L1.
This implies an extra 3% of IMLEO, regardless of whether the capability is used
or not.
Second Event
Chance: Is there water in the South Pole? In this case, the experts are
asked to rate the probabilities, result in an 80% probability of the presence
of at least 100MT of water in the lunar South Pole.
Third Event
Decision: Should a
mission be sent to the lunar South Pole? If this is done from L1, the cost is
trivial. If this has to be done from an equatorial Moon orbit this is implies
an expensive burn to change orbital planes, and an increase of 200% in IMLEO.
Fourth Event
Decision: Should a
permanent base be established at the lunar South Pole?
The decisions and
chances described above may be graphically represented in Figure 37.
Figure 37:
Graphical representation of decisions and chances for the example to decide
whether to have the capability to go to L1
Thus, the above
tree may be used to determine what decisions need be made at what points. This
method is not a “silver bullet” solution. Rather, it proposes a formal
framework for analysis, argue about options, and quantify the sensitivity to
factors. A diverse set of variables also helps to take into account the
sometimes-contradictory opinion of a group of stakeholders.
Real Options is a
method to value flexibility in system design. It evaluates the costs of the
enabling decisions taken today to perform certain functions in the future.
Since these decisions are enablers, they are a necessary condition. In
addition, a trigger must act in the future to decide whether to exercise the
option or not. In some sense, creating a real option is buying insurance. Thus,
a price is paid, whether or not this option is exercised. Once this price is
paid, one has the right, but not the obligation, to exercise the option.
In the example
explained above, the option bought will be the ability to use L1. Its benefits
are realizable only if a decision is made to land on the South Pole. To assess
the cost of the option, the following exercise may be executed.
First, compare the
utilities using the Decision Analysis tree, assuming that the mission path
never lands on the South Pole. As shown in Table
7, the expected utilities are 0.382 for having the
ability to go to L1 and 0.417 for not having that ability. In this case, the
choice of L1 ability is not ideal, since the polar opportunity is never
exploited.
Table 7:
Expected utilities from the Decision Analysis tree for the L1 capability
decision
|
Lands on South Pole |
Forbids South Pole |
Able to go to L1 |
0.574 |
0.382 |
Not able to go to L1 |
0.417 |
0.417 |
On the other hand,
when a decision is made to go to the South Pole, a utility of 0.417 with L1
capability is traded with a potential utility of 0.574 without L1 capability.
Given the decision
is not made to go to L1, the cost of the real option is therefore the
difference between the utility that the option is not taken (0.417) and the
utility of taking the option erroneously (0.382). Thus the cost of the option
is a utility of 0.035. Similarly, the potential benefit of taking this option
is 0.157. In this conceptual situation, the potential reward is significantly
larger than the potential cost, indicating that a decision should be made to
utilize L1.
Another way to
frame the decision of using L1 to access the Moon involves using the notion of
“expected mass” rather than utility. The following example uses real options
thinking to value the benefit of creating a system that has the option to use
L1 to explore the Moon.
Background and
assumptions:
An exploration
system starting in LEO is assumed to be composed of both a COV and a Lunar
Lander. Two operational architectures are considered:
1.) LOR: The COV and Lander enter lunar orbit.
The Lander descends and ascends to and from lunar orbit.
2.) L1: The COV enters L1 orbit, the Lander
descends and ascends to the Moon from L1.
If continuous
access to the poles is desired together with continuous free return (a likely
need if a lunar base is to be built at the poles), then the system
architectures present two possibilities:
1.) Use of L1, with a Lander that can descend
and ascend to and from L1
2.) Use of LOR, with a plane change burn once in
lunar equatorial orbit
The plane-change
burn at the Moon requires considerable ΔV and makes the L1 options more appealing.
Under what
circumstances would it be beneficial to have the option access L1? Framed as a
real option, one can examine the mass savings and penalties from LEO for using
L1 compared to the base case in which the equator is accessed. Decision
analysis is used here, although more complicated modeling techniques could also
be used to increase the accuracy of the valuation.
If a mission to the
Moon targets the equatorial region, it is clear that lunar orbit is the best
location from which to descend. A mission to the lunar equatorial region,
employing a Lander from L1 requires about 11% more Δv than from lunar orbit. If lunar pole access is required, however,
this architecture demands a plane change in lunar orbit, resulting in extra
mass in LEO compared to the base case. Conversely, if the pole does not need to
be accessed and the Lander is equipped with L1 capability, the architecture
requires extra mass in LEO compared to the base case.
Assuming that mass
in LEO is a surrogate for cost, we can calculate the expected mass of a mission
to the Moon, based on the probability of a decision to access the poles.
Expected Mass (EM):
EM = (Mass in LEO
for Pole Access)*P + (Mass in LEO for no Pole Access)*(1-P)
where P is equal to the probability that the
pole will be accessed for a given mission.
Figure 38: Decision tree for L1 capability example
Then, sensitivity
analysis can be used to determine where the value of the option, exceeds that
of the base case. The sensitivity analysis reveals that the option to use L1
becomes more valuable (less costly) than the base-case if there is more than a
30% chance that the poles will need to be accessed during the system’s
life-cycle.
Figure 39: Value of L1 capability
Of course, a number
of other factors will affect the decision-making process. The number of
missions to each location is an important factor in determining mass savings.
Also, this analysis does not consider the fact that mass savings, like cost,
might need to be “discounted” over time.
Still, options
valuation presents a powerful method to quantitatively justify a decision that
is currently sub-optimal, but may increase in optimality as circumstances
change. By framing the L1 decision as a real option, system architects can
design flexibility into the architecture and produce a more sustainable system.
Real options can be
used to evaluate commonality in the design of the transportation system for LEO
to beyond. The two major transportation architectural designs for the transfer
to the Moon and Mars are a staged and a cycler system. Traditional design
methodologies would evaluate both systems, but would eventually choose only one
design for the transportation system. Instead of following a traditional
approach, this section will describe how the design of the space transportation
system could be evaluated as a real option and how commonality can exist
between the designs of transportation systems with different destinations.
A staged system is
similar to the design of Apollo where stages were used to get from LEO to lunar
orbit and then from lunar orbit back to Earth. In the staged architecture, each
stage is discarded once it has been used; the use of stages in the
transportation design maximizes the efficiency of the rocket equation because
the design eliminates any dead weight that would have to be carried throughout
the mission.
A cycler
architecture differs from the staged architecture in that a cycler architecture
has no stages, consequently carrying around dead weight. One stage provides the
Δv for the entire mission. Where
a staged design would have two or more sets of engines and fuel tanks, the
cycler has only one set of engines and fuel tanks. The other main difference is
that the cycler is reusable after it re-enters Earth orbit. A common example of
a cycler is the modern automobile. Between destinations the physical form of a
car doesn’t change except for fuel. From a high-level perspective, the cycler
design has the advantage over the staged design because it can be reused from
mission to mission, while the expendable, staged architecture would have to be
rebuilt for each mission.
In order to
determine which design approach to take in the development of the
transportation system, a sample Mars mission was used as a baseline and the
minimum required mass in LEO was calculated. The characteristics of the sample
Mars mission can be found in Table
8. The total mass required at LEO for the staged
mission is roughly one-fourth of the cycler architecture mass. This result is
partially due to the fact that the cycler has to carry around additional dead
weight on the return trip that requires significant additional fuel for the
return trip, which in turn leads to the need for increased fuel for the
outbound trip. The most likely main reason for the large mass difference is
because the staged architecture is not required to reestablish Earth orbit on
the return trip and therefore the staged design does not have to carry the
additional fuel mass to perform a re-orbit burn. Because of the reusable nature
of the cycler, the vehicle must perform an additional burn in order to
reestablish Earth orbit. The additional burn to reestablish Earth orbit is
about 5 km/s and results in a significant mass increase in the amount of fuel
required by the cycler.
Table 8: Staged vs. Cycler transportation vehicle design
* All masses in Kg |
Staged
Design |
Cycler
Design |
Mass COV |
5,700 |
5,700 |
Mass Habitation Module |
55,000 |
55,000 |
Cargo Mass to Destination |
39,180 |
39,180 |
Cargo Mass Returned |
9,180 |
9,180 |
Mass of Fuel for Stage 1 |
176,267 |
>10,000,000 |
Mass of Stage 1 |
488,000 |
>10,000,000 |
Mass of Fuel for Stage 2 |
138,138 |
>10,000,000 |
Mass of Stage 2 |
170,394 |
>10,000,000 |
Total Initial LEO Mass |
740,000 |
>>10,000,000 |
Since the required
burn to reestablish Earth orbit is a significant constraint on the design of
the cycler, one must consider whether the burn at the Earth is truly required.
The cycler is required to re-enter Earth orbit, but the cycler is not required
to perform a burn in order to reestablish Earth orbit. It could be possible for
the cycler to perform some form of aero-braking in order to minimize or
possibly eliminate the need for a burn to establish Earth orbit.
After reevaluating
the required mass at LEO for the cycler assuming the use of aerobraking, the
staged architecture is preferred over the cycler architecture. However, the
mass required at LEO for the cycler architecture has been reduced by a factor
of three. The use of aerobraking did not change the preference of the staged
over the cycler architecture, although it significantly improved the required
initial cycler mass LEO. The results of the case in which aerobraking was
performed can be found in Table
9.
Table 9: Staged vs. Cycler design comparison with
aerobraking
* All masses in Kg |
Staged
Design |
Cycler
Design |
Mass COV |
5,700 |
5,700 |
Mass Habitation Module |
55,000 |
55,000 |
Cargo Mass to Destination |
39,180 |
39,180 |
Cargo Mass Returned |
9,180 |
9,180 |
Mass of Fuel for Stage 1 |
176,267 |
722,000 |
Mass of Stage 1 |
488,000 |
794,000 |
Mass of Fuel for Stage 2 |
138,138 |
281,000 |
Mass of Stage 2 |
170,394 |
309,000 |
Total Initial LEO Mass |
740,000 |
1,202,422 |
Perhaps, the
requirement of a burn was not the deciding factor in the mass difference. Instead, one might consider the
inefficiencies of the mass fraction. How would the required mass at LEO change
if the return fuel could be pre-positioned at the Moon or Mars? In order to
compare both transportation designs on an equal level, pre-positioning must be
applied to both the staged and cycler transportation designs. In the case of
pre-positioning for the staged architecture, the transfer vehicle would only be
required to carry the first stage on the outgoing leg. It could then drop off
the first stage and pick-up the second stage for the return flight home. In the
case of the cycler, the fuel for the return flight home could be provided at
the final destination, but instead of dropping a stage like the staged
architecture, the cycler would simply refuel using the pre-positioned fuel
provided at the destination. The design of the cycler would require that the
fuel stage be sized accordingly to accommodate the leg of the trip which
required the largest fuel mass. The design choice would result in either the
inbound or outbound trip with a sub-optimal use of the mass fraction equation.
It turns out that the extra structural mass is insignificant when compared to
the mass of the entire system.
The concept for
re-fueling the cycler opens the idea for the development of in-situ propellant
production. At a top level, it is conceivable that some form of cycler system
could also be used in the design of in-situ propellant production delivery.
Using a cycler for the delivery of in-situ produced fuel would be a tremendous
advantage because it would require only one delivery vehicle be developed, as
opposed to multiple vehicles in a staged system. However, in this case in-situ
propellant production was not considered in the design and return fuel was
pre-positioned similar to the staged system.
After recalculating
the design for the staged and cycler designs assuming pre-positioning of return
fuel, the preferred architecture again is the staged architecture. The total
mass required for the staged architecture went from 740,000 kg to 464,000 kg
and the mass required for the cycler architecture went from >>10,000,000
to 6,000,000 kg. The results can be seen in Table
10.
Table 10: Staged vs. Cycler design comparison with the
pre-positioning of return fuel
* All masses in Kg |
Staged
Design |
Cycler
Design |
Mass COV |
5,700 |
5,700 |
Mass Habitation Module |
55,000 |
55,000 |
Cargo Mass to Destination |
39,180 |
39,180 |
Cargo Mass Returned |
9,180 |
9,180 |
Mass of Fuel for Stage 1 |
176,000 |
1,500,000 |
Mass of Stage 1 |
193,000 |
1,650,000 |
Mass of Fuel for Stage 2 |
138,000 |
3,000,000 |
Pre-positioned Mass of Stage 2 (LEO) |
170,000 |
4,200,000 |
Total Initial LEO Mass |
464,000 |
6,000,000 |
The final trade to
consider is the situation in which both aerobraking and pre-positioning of
return fuel are used. In this case, the staged and cycler architectures combine
architectural elements of the pre-positioning and aerobraking cases. The
results were different from the previous cases: the cycler architecture is the
preferred architecture when the number of mission exceeds two. The mass
required for the staged architecture went from 740,000 kg to 464,000 kg, while
the mass for the cycler architecture went from 1,200,000 kg to 472,000 kg for
the first mission and 359,000 kg for each additional mission. Here the benefits
of the reusable nature of the cycler dominate the design of the transportation
architecture. These results are shown in Table
11 and the total mass in LEO per number of missions are
plotted in Figure
40.
Table 11: Staged vs. Cycler design comparison with
aerobraking and pre-position return fuel
* All masses in Kg |
Staged Design |
Cycler Design |
Mass COV |
5,700 |
5,700 |
Mass
Habitation Module |
55,000 |
55,000 |
Cargo Mass
to Destination |
39,180 |
39,180 |
Cargo Mass
Returned |
9,180 |
9,180 |
Mass of
Fuel for Stage 1 |
176,000 |
176,000 |
Mass of
Stage 1 |
193,000 |
193,000 |
Pre-positioned
Mass of Fuel for Stage 2 |
138,000 |
144,000 |
Pre-positioned
Mass of Stage 2 (LEO) |
170,000 |
178,000 |
Total
Initial LEO Mass |
464,000 |
472,000 |
Additional
Mass needed |
464,000 |
354,000 |
Total LEO
mass for |
928,000 |
826,000 |
Figure 40:
Total Cycler and Staged transportation systems LEO mass per number of flights
assuming aerobraking and pre-positioning of return fuel |
When comparing the
transportation designs for Mars to the design for the Moon, it can be assumed
that the resulting trends for Mars are the same for the Moon. Coincidently,
when the Δv requirement for a
Moon mission and the Δv
requirement for a Mars mission, assuming pre-positioning and aerobraking are
compared the total required ΔVs
are almost identical (~8km/s round-trip). Therefore, if pre-positioning and
aerobraking are used in the design of the transportation system, the
transportation system design for a Mars and Moon mission are almost identical.
It is conceivable that one vehicle could be designed such that it could provide
the transportation for both a Moon and Mars mission.
Now that the
results of these four trades have been evaluated, should the transportation
system be designed as a cycler or as a staged system? The answer is that the
transportation system should at first be designed as a staged system, as
expected. However, at the same time that the staged design is being developed
or used, research into aerobraking and pre-positioning should be examined. If
at any time it is discovered that either pre-positioning or aerobraking is
unlikely, then the design of the transportation system will remain as a staged
system. In the event that both pre-positioning and aerobraking have been found
to be feasible, NASA should then and only then switch to a cycler
transportation architecture. Therefore, in order to have the ability to switch
between architectures, NASA needs to develop a staged architecture that has
common elements that could be used in the development for the cycler system.
This commonality could be accomplished through a modular design for the
habitation module, COV, and transportation, most likely propulsion, system.
The design choice
of building a staged system with common elements to a cycler system is an
example of a real option. NASA only needs to develop a staged system, with commonality
in mind, and research aerobraking and pre-positioning in order to have the
capability for a cycler system. Only once both pre-positioning and aerobraking
have been found to be feasible should NASA make the decision to spend the
resources to switch the design of the transportation system to a cycler.
The theories and tools about decisions
analysis, which were presented in the previous section, were not used for all
the trades we considered. This part of chapter 6 presents the trades that were
considered for Earth to LEO options (launch site, Earth launch system for both
human and cargo, crew escape system and entry, descent and landing), followed
by trades for In-Space transportation (crew exploration vehicle, rover,
habitation module, pre-positioning, planetary landing systems, crew module
scaling, Moon options, Mars options). This voluminous chapter reflects all the
background studies that were performed during this class to support our
decisions for the baseline mission.
The policy that is
more likely to be applied for this human exploration program is that all
critical launches should be made from US territory. Taking into account that
all our architectures benefit from a launch from low latitudes
Directed by the new
policy, independent launch systems were considered for the crew and the heavy
cargo needed for human exploration missions. Also following the new policy,
only expendable launchers will be considered.
To transfer humans
from the surface of the Earth to a destination in LEO is a capability that, due
to the early retirement of the shuttle, has to be reacquired in the
A strong reason why
humans should be launched separately from the cargo is that the
The Crew Operations
Vehicle (COV) and functionally-similar Modern Command Module (MCM) are required
to transport a total of six people in some configuration or grouping, and must
have rendezvous and docking capabilities in LEO. It is a reasonable to assume
20 metric tons will be required in LEO.
The two launchers that can soon be adapted to launch such a human
mission are the EELVs. The Delta-IV provides the added advantage because it
uses the RS68 engine, which can be used to replace the SSME on an STS-derived
heavy launch vehicle for cargo. As much
work as possible from the Orbital Space Plane concepts that used capsules
instead of winged vehicles shall be reused in this design. Therefore, the launcher of choice for humans
is a human-rated version of the Delta IV.
The decisions
regarding the launch of cargo for the Mars and Moon missions have a very
important impact on the overall cost and feasibility of the Exploration
Program. We argue that an STS-based heavy launcher should be developed and
employed for this mission.
Since cost is a
bounded variable in this program it makes sense to include some cost estimates
in the overall evaluation of the different architectural options. One of the
main choices that face the program is the decision to develop a heavy launch
capability. It has been argued that, if
using modularity and extensibility, the mission’s hardware can be broken down
into parts, of about 20 metric tons, that are manageable by the heavy version
of the Delta IV (an EELV launcher). We will now compare this option to an
STS-derived heavy launch architecture.
A typical payload in LEO for a small lunar
human exploration mission is 118 metric tons.
To obtain the
number of Delta IV Heavy (DIV-H) launches that would be needed we cannot just
divide 118 by 20. A “penalty factor”
must be applied that accounts for the extra mass stemming from the rendezvous
and docking systems as well as the less structurally efficient geometry. This
factor has been chosen to be 1.3. This gives roughly 6 DIV-H launches.
All cost figures
were corrected for inflation into FY14 dollars using the Consumer Price Index
and the prognosis of the Office of Management and Budget and the Congressional
Budget Office. The cost of a DIV-H launch in FY99 is $170 million. Correcting
for inflation in FY14, the total cost of the six launches would be $1.4
billion.
This figure can be
benchmarked with the cost of the launch of a Saturn V rocket which was $431
million FY67, which makes almost $3 billion FY14.
The cost estimates
per launch of a Shuttle C were estimated to be $85 million FY85, which is $182
million FY14. This valuation, as is the case with programs that do not get to
the stage of operation, may be incorrect by as much as a factor of four. Therefore, we will assume that the cost of a
STS launch is approximately the cost of a shuttle flight and the more reliable
values that are given for a launch of the Space Shuttle will be used. This
value is highly dependent on the flight rate, so again caution should be
exercised. Assuming a flight rate of 6 per year, the value that is commonly
accepted for a shuttle flight is $245 million FY88, which is $477 million
FY14. For a flight rate required for a
crew of six, the STS-derived is probably too ambitious. If we assume a flight
rate of 4 per year, that is a Moon trip every three months, then a single
flight would be about $715 million FY14. It should be noted that this value is
substantially less than the value obtained for the launch using 6 DIV-Hs. A
flight rate of 2 missions per year gives a break even in the cost making. From a cost point of view, both architectures
are equally attractive at a price tag of $1.4 billion FY14. Naturally the same
argument applies even more strongly to the case of Mars missions.
Another advantage
of the STS architecture over the DIV-H is that, since all the hardware for a
small Moon mission is launched at one time, automatic rendezvous and docking
capabilities are not a critical new technology to be developed for the lunar
missions.
This capability of
automatic rendezvous and docking will be necessary when human exploration
missions to Mars will be attempted. The mass budget in LEO required for even
the simplest human exploration mission to Mars are in the range of 200-600
metric tons. This can be reasonably done with 2-6 STS based launches per
mission, which is a flight rate of Mars missions of roughly 1 per year.
Therefore, for an
Apollo-class Moon mission, one STS would launch most of the mass and a roughly
20 tons capsule, launched separately, would carry the humans to dock with the
rest of the mass.
Among several
architectures for an STS-derived vehicle the one that seems most attractive is
an external tank two SRBs and three disposable 3RS68 as well as a newly
developed J2 class upper stage. Such a vehicle can deliver roughly 100 metric
tons in LEO. A detailed explanation and performance curves can be found in
Appendix 9.1.3.
It has been argued
that a new launch system will be better than an STS-based design because most
of the problems with the STS are not a consequence of the Orbiter’s design, but
are rather related the parts that would be kept in any STS-derived. For
instance, note the problems with the O-rings in the SRBs and the foam in the
ET. To provide a complete answer to that question falls beyond the scope of
this study, however the results of this study can be used as a baseline to know
what a new design should consider in terms of cost and performance as compared
to an existing STS based system.
Human spaceflight
escape systems have been developed for on-pad abort and boost phase
emergencies. Once in-orbit, the escape
mechanism for the crew is the same as the normal re-entry sequence into the
Earth’s atmosphere. Throughout the
history of human spaceflight, ejection seats and escape towers have been
developed to provide this additional layer of safety to the crew. However, the option of including such systems
must be traded against significant mass penalties (Nuttall, 1971).
The first manned
orbiter, Vostok I, had an ejection seat escape device. In addition to providing an escape mechanism
for the cosmonaut on the pad and during the boost phase, this ejection seat
functioned as the normal means of landing after post-orbital descent.
The first
The US Gemini
escape system also utilized ejection seats just as the Vostok I. This ejection system was flight-tested up to
20,000 ft and Mach 1.75. The decision to
use ejection seats instead of the escape tower incorporated into the design of
Mercury was driven by the hundreds of kilograms in mass savings. Unlike the Vostok I, the dual rocket-powered
ejection seats were only used in landing emergencies (on-pad, pre-orbital
ascent, post-orbital descent over land).
Normally, the Gemini capsule was lowered to the ocean by a large parachute.
As in Mercury, the
Apollo Launch Escape System (LES) utilized an escape tower. Providing an emergency escape capability to
the crew from the on-pad launch sequence to the end of second-stage ignition,
the LES engines weighed 5,500 pounds with a total structural mass in excess of
9,000lb. Its maximum operational
parameters were 320,000 feet and a Mach number of 8.0. The LES consisted of three solid-propellant
rocket motors. After the firings of
explosive bolts to separate the command module from the service module in an
escape sequence, the launch escape motor would pull the 11,000 pound command
module to safety using a 155,000 pound thrust solid rocket (Townsend,
1974). The tower-jettison motor was employed
to separate the escape tower from the command module prior to parachute
deployment (Lee, 1971).
As for Apollo, the
Soyuz Emergency Escape System (EES) utilizes an escape tower of
solid-propellant rocket motors to pull the crew capsule to safety in an
emergency during on-pad launch operations of the boost phase. The EES is operational throughout all phases
of powered flight trajectory prior to orbital insertion (Kolesov, 1969).
A seated tractor rocket escape system was proposed for the STS in the
wake of the Challenger accident. The
tractor system is lighter and less voluminous than an equivalent ejection seat
system, however, aerodynamic “blow-back” causes unsuccessful extraction at
altitudes above 15,000 feet (Ondler, 1989).
Utilizing Lockheed
Martin’s Pad Abort Demonstration (PAD) platform consisting of sensors and
mannequins in a simulated crew cabin to measure accelerations and motions
generated, NASA will conduct seven integrated PAD test flights during 2005-2006
to test an escape tower system of four 50,000-pound thrust RS-88 rocket
engines. These tests aim to trade
various propulsion systems; parachute deployment, vehicle configurations, and
landing techniques for a future tower escape system (Orbital Space Plane/Crew
Exploration Vehicle).
Ejection and
tractor rocket seats are lighter than tower escape systems by an order of
magnitude—weighing hundreds of pounds instead of thousands. However, because the tower escape system is
jettisoned during ascent while ejection and tractor seats are carried in the
service module throughout the mission, LEO payload mass reductions for tower
escape systems are approximately only two to five times greater than ejection
and tractor seats.
Figure
41:
Minimum LEO payload mass penalty for EELV tower escape
Figure 41 displays the lost LEO payload mass when a 5,500lb
(2495kg) tower escape system is added to various EELV designs. (A tower escape system of this mass is a
minimum estimate for systems capable of saving service modules in the 6,000kg
range). It is assumed in these
calculations that the escape system is jettisoned with the first stage.
For Delta-IV
vehicles, the average impact is ~6% reduction in payload to LEO. Specifically, a Delta-IV designed to launch
6,760kg would have the payload mass reduced by 441kg, a Delta-IV designed to
launch 9,070kg reduced by 610kg, and a Delta-IV designed to launch 20,500kg by
1,210kg. The Atlas-V vehicles have an
average 5% reduction in payload mass to LEO.
For Atlas-V launchers designed to launch 10,300kg, 12,500kg, and
20,520kg, the payload mass reductions are 621kg, 530kg, and 880kg,
respectively. For an EELV-derived
heavier lift vehicle capable of placing 50,000kg in LEO, adding the escape
system would have a much lower effect—reducing the payload capability by only
251kg (or about 0.5%).
Figure 42:
Launch escape mass as a function of crew module mass (Source: Orbital Science
Corp.)
Figure 42 displays how tower escape system mass scales
parametrically with the crew module mass according to a model developed by
Orbital Sciences Corporation. With
escape system ranging from 6,000-8,500lb, the 5,500lb escape system mass
selected to calculate the LEO payload mass reductions for EELV architectures
(see Figure
41) is clearly on the low-end of the scale.
To appreciate the
available options for entry vehicle shapes, one must first determine the
criteria affecting selection. Decreasing
development cost has been identified as a major constraint in the project. Another highly important requirement is
minimizing mass. A subjective criterion
to represent minimum mass for given vehicle’s shape is volumetric
efficiency. A third objective might be
to limit the peak deceleration forces on the crew. Clearly, there are other criteria affecting
the shape of the entry vehicle, but these have been identified as the most
important. Next, the relative importance
of the identified criteria is determined:
Minimum development cost 0.5
Volumetric efficiency 0.3
Peak entry deceleration 0.2
Total 1.0
Soyuz, Apollo, and
heat shield with afterbody all have the same development cost, whereas the
lifting body will have the highest development cost. The biconic most closely resembles a
cylinder. Soyuz and Apollo are almost as
efficient as the biconic, but their blunt conic shape is more conical. The heat shield and afterbody shape is
cylindrical, but the cylindrical diameter is smaller than the heat shield, so
some volume is wasted. The lifting body
sacrifices volumetric efficiency in the interest of streamlining. The comparison method used for assessing peak
entry deceleration indicates the lifting body would have the lowest and the
Soyuz would have the highest deceleration.
Figure
43:
Entry vehicle shape pair-wise option comparison
The selection
criteria weightings generate an overall score, which is mapped onto a number
line in Figure
44. Notice that
the criteria-weighting factors directly influence the final rankings. Importantly, these rankings are subjective
assessments that should not suggest an “optimal” option. Experience and intuition might confound the
option space. For example, a winged body
might be too difficult to equip with thermal protection for high-speed lunar or
Martian returns. For the purposes of
this report, an Apollo-class entry vehicle, termed the Modern Command Module
(MCM) shall be used for Earth return.
Figure
44: Comparison scale for entry vehicle
Descent and landing
is the flight phase designed to reduce the horizontal and vertical velocities
to a desired value for surface touchdown.
The thick atmosphere of the Earth allows a spacecraft to follow the
aeroentry phase with an inflatable, parachute, or parafoil deceleration all the
way to the surface. The Apollo Command
Module used parachutes to an ocean splashdown, and the Russian Soyuz capsule
rides a parachute until a retrorocket fires just before a land-based touchdown.
To achieve mission
objectives, atmospheric entry is constrained by three fundamental requirements:
deceleration, heating, and accuracy.
Although a vehicle’s structure and payload limit maximum deceleration,
we must consider the requirements of a human-rated exploration system. Well-conditioned humans can withstand a
maximum of about 12 Earth g’s for a
short time. A system designed for
de-conditioned crew must produce less than 3.5-5 Earth g’s (Hale, 1994).
Friction between
the speeding entry and atmosphere generates heating that must be dissipated
during the few minutes of atmospheric entry.
The thermal protection system must withstand the total heating and the
peak-heat rate encountered during entry.
A third important
mission requirement is accuracy. The
spacecraft’s capability to maintain a predetermined trajectory depends on its
inertial navigation systems and available ground support.
Details of the
calculations can be found in Appendix 9.2.2.7.
As an alternative
to the heavy ablative heat shield systems, researchers have refocused studies
on inflatable technologies such as the ballute and Inflatable Braking Device
(IBD). The IBD is an aerodynamically
shaped cone to increase the surface area of the entry vehicle. The increase in surface directly affects the
ballistic coefficient, β, of the vehicle, thereby decreasing the maximum
heat and deceleration loads during entry.
A Primary IBD is
inflated just before reaching the atmospheric interface. Once the maximum deceleration, pressure, and
heat flux are passed, a Second IBD inflates to replace the parachute system at
the appropriate altitude. A Third IBD may
also be inflated to increase the size of the vehicle to achieve the required
terminal velocity. Depending on the
design, the landing system can either be one of the conventional landing
systems or can be replaced by one of the stages of the IBD that cushions the
impact.
EDL systems based
on the conventional approach benefit from a strong and proven heritage, but
depend on the use of a heavy heat shield and a dedicated landing system for
re-entry. To offer a unique perspective
on such a system’s purported benefits of inflatable over conventional EDL
systems, a parametric comparison study was performed on for an Apollo-class
Earth return vehicle. Graphical results
are shown in Figure
45. Because of
the infancy of research in inflatable technologies, sizing was scaled from a
proposed post-Beagle2, robotic mission to Mars.
Figure
45: Parametric comparison of inflatable versus
conventional Earth re-entry technology
To subjectively
assess the trade space of possible descent and landing system combinations, a
pair-wise option comparison was performed in Figure
46. The
methodology used in this study follows that of Section 6.4.2.4.1. The selective criteria and relative weighting
chosen for this comparison were:
Minimum mass 0.40
Minimum development 0.25
Maximum cross-range 0.25
Minimum peak
deceleration 0.10
Total 1.00
The mass, cross-range,
and peak deceleration of the descent and landing systems were calculated using
the methods previously stated. The
minimum development cost was subjectively assessed based on current technology
readiness level (NASA/TRL).
Figure
46: EDL pair-wise option comparison
The analytical
calculations assume the use of an Apollo-class Earth re-entry capsule. The conventional systems rely on an ablative
SLA561V heat shield. The EDL systems
that include inflatable devices substitute parachutes with the Third IBD. The conventional system that includes
retrorockets for touchdown deceleration also includes the release of drogue
parachutes. Notice that the purely
inflatable system has the least mass, the conventional system with parachutes
require the least development time, the inflatable system with retrorocket
decelerators produces minimum peak deceleration, and the inflatable system with
parafoil technology has the maximum cross-range capability.
The options are
subjectively ranked, using the selection criteria weightings listed above. Notice that the ranking of the entry and
descent architectures follows the distribution of the system mass. This ranking should be used only as a
qualitative estimation that is inherently dependent on the relative selection
criteria weightings chosen.
Table
12: EDL
option ranking and system mass for an Apollo-class Earth re-entry vehicle
The parametric
comparison of Apollo-class inflatable and conventional Earth EDL systems
yielded interesting results. The
inflatable system had 15% to 20% less mass, 40% to 45% less maximum
deceleration, 20% to 25% less average surface temperature than a conventional
heat shield system with a cross-range capability that is roughly equal to a
parafoil system. Additionally, the
inflatable device can serve as an air-cushion for ground-based landings or as a
flotation device for water-based landings.
Although the inflatable systems seem promising, they are untested. Consequently, the system’s cost, reliability,
and safety are difficult to estimate. A
test flight in February 2000 from
The chosen EDL
architecture directly influences the choice of landing site. The Apollo Command Module landed in the water
to reduce the touchdown impact of its unpowered descent. Similarly, Soyuz fires a small solid-motor
thruster just before touching down on land.
In addition, uncertain atmospheric density, navigation errors, and
unanticipated winds can push an uncontrollable vehicle, such as a spacecraft on
parachutes, away from its intended landing location. The Apollo Command Module landed in the
There are roughly
three recovery possibilities: land, sea, and lake/coastal. Recovery operations on land can be relatively
fast and inexpensive by utilizing existing infrastructure. Land-based touchdowns require a sink rate
below 7.5 m/s, whereas water landings can sustain velocities of about 9.5
m/s. Because land-based landing g-loads
can be 2 to 3 times higher than water-based landings, a crushable nose or inflatable
air-cushion is required. To distribute
the impact forces over the entire lower surface of the spacecraft, a
self-leveling honeycomb might be used to plastically deform to absorb the shock
of the landing. Possible materials for
the honeycomb might include lightweight metal alloys, carbon-carbon composites,
and high density styrene polymers. This design would provide significant mass
savings over conventional landing mechanisms because a composite honeycomb
weighs a fraction of aluminum or steel.
Sea-based landings
tolerate higher impact velocities, reducing the mass needed to decelerate the
entry vehicle. The mass needed for
flotation bags might partially offset this benefit, unless an inflatable
landing system is used. Water also
provides immediate cooling of the overheated spacecraft. The Apollo program demonstrated that recovery
operations at sea can be costly, and can be adversely affected by poor weather.
Sea and
lake/coastal-based landings share similar properties, except that lake or coastal-based
landings have lesser infrastructure costs.
A lake-based landing could use existing Coast Guard recovery
capabilities, instead of deploying a large Naval Carrier Battle Group. Possible landing sites for such a landing
might be the Gulf of Mexico or the
A number of trades
were examined in determining the forms for the extensible Moon/Mars mission
architecture. The space transportation
system is a network of modules that was developed from the trades described
below. It was assumed that the space
transportation system does not require the use of the International Space
Station (ISS) as an assembly or return point.
This was done to ensure that NASA can divest itself from the ISS and STS
to meet the Space Exploration goals within budgetary and political constraints.
The Mass
Transportation Vehicle (MTV), as presented in the baseline mission, is made of
the Crew Operations Vehicle (COV) and the Habitation Module (HM). It is part of
the Crew Exploration System (CES), which consists of all the forms necessary to
support manned exploration of the solar system.
The different
phases that the MTV is required to perform is in space transportation. The first trade study, which led us to
dissociate the Earth to LEO transportation from the In Space transportation is
described below and shown in Figure
47.
-
From
Earth to LEO and back to Earth
-
In-space
travel: LEO to another destination in space
For this initial
trade, we considered only small range exploration (up to 8 crew and 40 days),
which excluded Mars exploration. The results and lessons learned from this
short study led us to the choice of separating the forms as much as
possible. Most notably, it will be
highlighted that separating the function of in-space and Earth to LEO
transportation is a beneficial choice.
In this initial study, the CES (Crew Exploration System) performs the
following: it goes from Earth’s surface, travels in orbit or further (but
middle range) and comes back to the Earth surface’s surface at the end of the
mission.
The basic functions that are required are listed below:
-
Support
and shelter the crew during launch
-
Provide
crew escape in case of a launch emergency
-
Provide
a habitat for the crew during in-space transportation
-
Provide
energy to displace the crew module during in-space phases
-
Perform
landing on a selected site
For performing these functions, we have studied two forms: service
module and crew module (called HM here) – the names are internal to this trade
study.
Figure
47:
An interactive
model was developed to determine the mass of various components of a MTV based
on the number of crew and the mission duration.
The model is described in Appendix 9.1.1.
A strong inter-level dependence
exists between technologies used for the various functions to be
performed. For example, the mass of the
re-entry system depends on the volume of the capsule, on the type of deceleration
device, etc. Figure 48 shows the decision tree for key technologies/options
that have been traded. Links between
levels represent preferable or feasible options. For example, the lake/coastal landing site
option requires a re-entry system with a large cross-range capability for
precisely landing into a small body of water (inflatable or parafoil).
Three main categories of HM have been identified, as shown on the figure
below:
·
Combined
(one unique form which
transports the crew for all the sections; Earth to LEO, In-Space and Landing)
·
Separate
(two different forms; one
performs In-Space transport and the other performs Earth to LEO and landing)
·
Flexible (the same core transports the crew, but
minor modifications are made so that it is able to land or perform In-Space
transportation)
For each category,
the vehicle could be expendable or reusable (see
Figure 49).
Figure
48:
Elements of the MTV, assuming a crew of three for a ten-day mission
Figure
49:
Classification of existing crew transport modules
To assist in decision-making,
three metrics were used: mass, TRL, rank.
The Technology Readiness Level (TRL) methodology is a NASA metric based on a
nine-stage process ranging from the basic principle being observed and reported
(#1) to flight proven through successful mission operations (#9). The rank-measurement enabled a normalized
comparison across elements. For example,
the launch escape system shouldn’t be chosen for the same reason as the EDL
elements. Each element of the CES (each
row in the network) has its own criteria.
Ranking also enables comparison between each option in a row with
different metrics, while trying to assess each option as objectively as
possible for trading different measures of performance/priority. Ranking also allows weighing metrics by
priority. Depending on the priority, you can weigh the metric so that the final
ranking reflects priorities.
Three of the many
combinations were considered in detail:
·
Modern Apollo CM - MCM - Tower
Escape, Modern CM and SM, Conventional Re-entry, and Sea Recovery.
·
Improved Soyuz - IS (Best Rank) - Ejection
Seat, Soyuz DM & OM & SM, Inflatable Re-entry, and
New Type - XTV (Lowest TRL) -
Ejection Seat, OASIS CTV & CTM, Inflatable Re-entry, and
The mass of these configurations is shown in Figure 50.
Figure
50:
Configuration masses (10-day to 40-day missions)
The Improved Soyuz (IS) architecture is a separate-expendable type of Habitable Module. It is comprised of all of the best-ranked components for the launch escape system, habitable volume, service module, re-entry system and landing site.
This architecture has the
Ejection Seat Launch Escape System, which is ranked as the best launch escape
system based on the metrics of mass penalty, reliability, cost and weight. As can be expected, if the relative weights
are altered, the final launch escape system ranking may change. The separate expendable Soyuz crew module had
the best overall ranking in the following categories (minimum launch mass,
minimum development cost, autonomy and flexibility). Similarly, the best-ranked technique of
re-entry was inflatable re-entry. This
was based upon the minimum mass, minimum development cost, minimum deceleration
and maximum cross-range. The method of
landing that was best ranked was a water landing. This was based upon the minimum recovery
time, least weather affected, minimum infrastructure cost and maximum landing
speed.
Scaling was performed for both
the launch escape system and the crew module.
Based on the analysis, which proposes that mass of the vehicle is a
function of both the mission duration and number of crew, it is clear that the
length of the mission from Earth to LEO does not greatly affect the overall CES
mass. It should be noted that the
mission durations highlighted here are completed arbitrary and were chosen to
illustrate that the required increases in structure and CES component mass will
not be the primary factor affecting mass increases.
The “Modern” Apollo uses all of
the same methods of re-entry and transportation modules as the original Apollo,
however a structural analysis was performed, which determined a new vehicle
mass based on modern materials. For this
architecture, the COV mass is greater, but it is still moderately dependent on
the mission duration.
Using the same methods of launch
escape and re-entry as the Soyuz based architecture discussed earlier, the
Oasis XTV-CTM combination was chosen as a third architecture to present because
of its lowest TRL, but also second-best rank after the Improved Soyuz. For this architecture, the mission duration had
a greater influence on the overall architecture mass compared to the other
architecture. The vehicle structure
comprises a much greater proportion of the overall mass than in the case of the
“Modern Apollo”. Since the mission
duration is related to the habitable volume and the external vehicle surface
area scales the structural mass, the overall mass is more greatly affected for
this case. Even though this
configuration had the lowest TRL, which could indicate the use of advanced or
modern technologies, other configurations had lower masses.
A summary of the three
configurations is shown in Figure 51.
Figure
51:
Three COV configurations for launch from Earth to LEO
The mass of the COV
was approximated as 5708kg from the Model described in Appendix 9.1.1. It was
assumed that the EDL mass could be neglected and the additional mass required
to aerobrake at Mars was 15% greater than the COV mass. Similarly, the additional mass required to
aerobrake at Earth was 6% greater than the COV mass (Larson, 1999).
The main lesson
learned from this trade study was that the forms used for Segment 1
(Earth to LEO and back to Earth) and Segment 2 (In-Space transportation)
should be separate and expendable. Such a separation leads to a high
rank and low mass within the framework of this trade study. A vehicle that performs all the functions at
once (such as the Shuttle) is sub-optimal and leads to additional mass,
especially when it is reusable.
6.4.3.1.1.2.1. Transportation Form
for Segment 1
The conclusions
highlighted in the previous Section guided the from selection process for
Segment 1 (Earth to LEO and back to Earth) of the baseline mission. The
paragraph below explains why the Modern Command Module (MCM) form
was selected for the baseline mission.
When determining
the type of COV to use for launch and re-entry,
the mass of the three configurations without their respective Service Modules
were determined. From this analysis, the
Modern Apollo Command Module was observed to have the lowest mass (5,200kg) and
was selected as the form for Earth launch at the start of the mission and Earth
EDL at the end of the mission. Since
this vehicle has approximately three times less the habitable volume per person
as compared to the OASIS XTV, this may indicate that separating the function of
crew habitation and re-entry is beneficial to overall mission mass reduction.
6.4.3.1.1.2.2. Transportation Form
for Segment 2
For Segment 2
(In-Space transportation), a modular approach was taken to ensure increased
commonality between the forms required to complete a Moon and Mars exploration
mission.
Surface exploration
of the Moon and Mars will require a diverse array of robotic capabilities. Mobility systems such as rovers are critical
to achieving scientific missions and accomplishing a variety of operational
requirements. Use of rovers will
increase effectiveness and safety while reducing costs. Tasks to be performed include instrument
deployment, soil manipulation, and human transportation.
Four categories of
rovers exist. Automated, autonomous
rovers are equipped with artificial intelligence for hazard avoidance and are
capable of collecting and communicating scientific data. Autonomous rovers on precursor missions may
identify potential landing sites for human missions and refine Lunar/Martian
scientific goals. Remote controlled
rovers may be utilized by human operators on Mars for missions of varying
duration. Unpressurized rovers can be
driven by astronauts but should only support exploration within walking
distance of living quarters. The Apollo
Lunar Rover Vehicle and Soviet Lunakhod are legacy unpressurized rover systems
(
Table 13: Rover functional requirements
Category |
Unmanned
Precursor |
Short-Stay |
Medium-Stay |
Extended-Stay |
Automated,
Autonomous |
X |
|
|
|
Remote
Controlled |
|
x |
x |
x |
Unpressurized |
|
x |
x |
|
Pressurized |
|
|
|
x |
To support these
operations, modular rover architecture is proposed whereby rover for instrument
deployment and soil manipulation tasks can be assembled from an inventory of
modules to accomplish a specific task.
This inventory of modules includes actuated joints, links,
end-effectors, sensors, and mobility units.
Initial configuration and reconfiguration can be done autonomously or by
an astronaut (Farritor, 2000). A modular
architecture for robotic surface operations may represent a new paradigm in
NASA robot design but it does not rely on developing new technologies.
For our design
space—multiple missions to the Moon and Mars—a modular rover architecture was
selected for a variety of reasons:
1) Efficiency: for missions requiring a wide variety of tasks a single modular system
is superior to creating a dedicated robot for each task, additionally,
packaging modules on launch vehicle may be more efficient than packaging
assembled robots in terms of mass and volume
2) Adaptability: modules enable construction of novel
robots, including robots for tasks that are not foreseen
3) Reliability: failed modules can be replaced, different
configurations can potentially accomplish the same task, Mars missions should
place a premium on this reliability
4) Extensibility: fits spiral development model of
increasing capability over exploration program life
5) Cost: standardized modules will limit non-recurring research and development
A variety of design
architectures are possible with open and pressurized rovers to transport
astronauts on the Lunar and Martian surface.
For mobility, tracks, screw drives, legs, rockets, and balloons are all
available, although wheels offer the greatest overall performance when
considering energy, ground pressure, ground clearance, reliability, and human
factor requirements. For rover structure
and pressure shell, an inner pressure shell of aluminum alloy and an outer
shell of aluminum and carbon/graphite epoxy offer a strong baseline
design. The communications system must
maintain contact with all manned rovers at all times for navigation, scientific
investigations, and safety. As a legacy
system with proven reliability, hydrogen-oxygen fuel cells are an ideal power
source. For life support, an open system
is recommended given the relatively short excursions and high mass penalty
(~1000kg) to recover consumables (
An open,
unpressurized rover is limited to sorties of 10km for safety considerations
(within walking distance of their surface habitat). These vehicles will typically support a crew
of two. Therefore, although not a
requirement, a pressurized rover capable of sorties ranging from 50-500km and 12-400
hours duration is a recommended option for extended-stay missions (
The Habitation
Module (HM) will sustain human life for an extended-duration Mars mission. This module will be launched in two pieces
and assembled in LEO, allowing the overall volume to not be limited by the
minimum launch volume requirements (see Figure
52). Since there
are two modules, none, one or both can be used for each mission, depending on
the mission requirements for duration and crew size. This module must have propulsion capabilities
to perform docking maneuvers in both Earth and Mars (or Lunar) orbit.
Figure 52: Mars/Moon Transfer Vehicle (MTV)
Following the Mars
study performed by (Larson, 1999), the mass of the HM was calculated as
~55,000kg for a crew of six, depending on a number of critical factors (mission
duration, type of radiation protection, life support, supplies, aeroshield and
power requirements).
The HM is composed
of separable modules that promote significant modular spacecraft design
flexibility. Six of these modules are
combined in two groups of three and platform, forming one large volume required
for a Mars mission. Based on Larson
(1999), it was assumed that a habitable volume of 20m3 per person
was required for a 6 crew, 6-month mission.
For this analysis, 30m3 was specified per person. It was also assumed by Larson (1999) that 33%
of the total volume was assumed to be habitable. Therefore, a total volume of 540m3
could be created by 6 octahedrons as shown in Figure 52, each with a 5.6m diameter. As shown in Figure
53, this volume agreed well with other pressurized
spacecraft volumes.
Figure
53:
Historical space habitat pressurized volume (Kennedy, 2002)
Based on Larson
(1999), an estimate for spacecraft mass (based on current technology) could be
made:
, (1)
where m is
the total vehicle mass, N is the number of crew, D is the mission
duration in days and V is the total spacecraft volume in m3. Based on the missions in question, the
results compared well with the calculated 55,000 kg mass for the Habitation
Module.
A summary of the
module masses is given in Table
14.
Table 14: Baseline module masses
Pre-positioning
items needed for a space exploration mission is the act of sending hardware or
any other required cargo to its respective destination in advance of the “main”
portion of the exploration mission. In
the case of this space exploration architecture, the “main” portion of the
mission is the launch through the landing of the human crew.
Non-time critical
components with lifetimes of appropriate length that are not required for the
crew to have available during transfer to their destination are good candidates
for pre-positioning.
The pre-positioning
of mission components will likely be accomplished with an efficient propulsion
system technology such as electric propulsion.
Once this propulsion technology is successfully demonstrated for large
masses, electric propulsion will then be used to send more exploration
equipment. However, it should be mentioned that until the ability of the
crew to live off re-supply provided by electric or other efficient propulsion
systems, or “live off the land” has been demonstrated, the crew should carry
all of the material with them needed for safe return.
One major reason to
pre-position items for a space exploration mission is to take advantage of
being able to transport this cargo using a more efficient propulsion system
than would be used otherwise. This
results in a reduced overall mass of the pre-positioned module. For the purposes of this project, it is
assumed that this propulsion system will be a form of electric propulsion.
Electric
propulsion, while more efficient, has much less thrust than engines using
chemical propellant. This requires a
longer time of flight to get the cargo to its destination. This necessitates that the cargo being
pre-positioned is not time critical.
This increased
efficiency of electric propulsion over chemical is due to a significantly
increased specific impulse, Isp. The famous “rocket equation,” shown below, is
used to exemplify why this is beneficial.
(1)
where ΔV is the change in velocity
provided by the engine burn, g0
is acceleration due to gravity on Earth, mi
is initial mass, and mf is
the final mass after the engine burn.
Solving for the
final mass:
(2)
This shows that an increase
in Isp results in a
decrease of initial total mass required.
This is the resultant benefit of a more efficient propulsion
system. The relaxed time of flight of
the pre-positioned mission phase, compared to the crewed segment, makes this
advantage possible.
The initial mass
benefit due to pre-positioning for this project can be seen in Table 15 and Table
16 for missions to the Moon and Mars, respectively.
Table
15: Mass benefit using pre-positioning for a
Medium Moon mission
Total
Mass in LEO Without Pre-positioning (kg) |
Total
Mass in LEO With Pre-positioning (kg) |
101,000 |
80,000 |
Table
16: Mass benefit using pre-positioning for an
Extended Mars mission
Components
Pre-positioned |
Total
Mass in LEO (kg) |
SH |
745,000 |
SH, Landers |
463,000 |
SH, Landers, Earth-return fuel |
379,000 |
For missions to the
Moon, Table
15 shows a significant mass savings due to
pre-positioning non-crewed mission components at the destination. For missions to Mars, Table 16 shows an increasing benefit as more mission
components are pre-positioned. The use
of efficient propulsion systems such as electric propulsion combined with a
relaxed time of flight requirement allow for such a mass savings.
6.4.3.1.4.2.2. Launch Vehicle Selection
A more subtle
benefit for pre-positioning non-time critical mission components is the ability
to use less expensive, non-human-rated launch vehicles to launch these
pre-positioned components. The launch
vehicle used to launch the crew will likely be a new or partially new
heavy-lift launch vehicle design. The
human-rated launch vehicle is likely to cost more per launch than launch
vehicles such as the Evolved Expendable Launch Vehicles.
For example, EELVs could be
used to launch “packages” to be pre-positioned in advance of the main crewed
portion of the mission as opposed to launching portions of this pre-positioned
cargo with the crew on a more expensive heavy-lift launch vehicle.
In addition, if enough mission components are pre-positioned using launch
vehicles such as EELVs, a reduction of the payload mass requirement for this
new man-rated launch vehicle could be realized.
An inherent
advantage with pre-positioning is the reduction of mission risk. This risk reduction is possible for two main
reasons. First, mission planners on
Earth would know in advance of the launch of the crew if the pre-positioned
components were successfully deployed in their desired locations. Second, mission planners would also know if
these components are functioning properly before the launch of the crew.
If a pre-positioned
mission critical component is found to not be functioning properly before the
crew launches on their mission, the mission planners have several options to
solve the problem. One option is to try
fixing the problem via communication with the malfunctioning component and
delay the launch of the crew if required.
Another option is to launch a replacement component either using
chemical or electric propulsion depending on the mission schedule. Finally, a replacement component could be
sent with the crew when they launch as scheduled.
Finally, the risk of launching many mission components using existing
launch vehicle technology such as EELVs will likely be less than sending the
same cargo using a new or partially new heavy-lift launch vehicle.
Some drawbacks
exist for incorporating pre-positioning into a space exploration mission
design. First, pre-positioning may
require an increase in the number of required Earth launches. This may increase the risk of a launch
failure. Second, the designs of the
pre-positioned components will need to accommodate increased component lifetime
requirements to account for an increased time of flight and the time between
the arrival of the component(s) at the destination and the arrival of the crew. A final penalty to pre-positioning mission
components is the increased time required of mission control personnel for
monitoring and control of pre-positioned components. These personnel will likely need to monitor
these components for an extended period of time before the launch of the
crew. If everything were launched at the
same time, less time would be required for monitoring mission components.
Varying the mission
duration and the number of crew greatly affects the overall mass of the
habitable crew vehicle. It was necessary
to develop a method of accurately scaling the habitable crew vehicle to predict
the required launch to LEO crew vehicle mass.
By comparing the
masses of various crew modules (Gemini, Mercury, Apollo CM, OASIS CTV and
Soyuz), while holding the habitable volume constant, the various architectures
were compared. By answering the
questions shown in Figure
54, a new vehicle mass can be obtained.
Figure
54:
Flowchart of scaling analysis
It should be noted
that the heat shield mass was determined in a separate analysis and was not
considered in this analysis. Using
modern materials and structural analysis on the Apollo CM, the mass might be
reduced. For comparison, the materials
used for the OASIS CEV were used for the Modern Command Module (MCM) and Crew
Operation Vehicle (COV) structures. From
a detailed analysis described in Appendix 9.1.2., Figure 55 was obtained.
Figure
55:
Vehicle mass scaling (broken line: 3-day mission, solid line: 30-day mission)
The masses shown in
Figure
55 do not include a heat shield because a comparison can
be made between functionally equivalent forms.
Details of heat shield sizing are described in Appendix 9.2.2.7.2.
As the mission
duration increases, the habitable volume increases, causing the vehicle surface
area to increase correspondingly, to maintain the vehicle proportions. As explained earlier, when the surface area
increases, all of the mass breakdown components that scale with surface area
increase. Since the majority of the mass
breakdown components that were scaled for the Apollo CM scaled with the vehicle
surface area, there was a significant overall vehicle mass dependence on the
mission duration and/or crew. When the
mass per surface area was reduced (as was the case for the Modern Command
Module analysis), the influence of mission duration increased the overall mass
of the Modern Command Module only slightly (see Figure 55).
FoxTrot by
Bill Amend
Figure
56: The reality of designing an EDL system
(Amend, 2004)
EDLA phases and
options for lunar and Martian missions, including Earth return, are summarized
in Figure
57. The connected
boxes identify a possible human mission scenario. Although aerocapture has never been employed
on a human mission, it should substantially reduce the total required vehicle
mass (see Section 6.4.3.3.4). Human missions require drag- and
lift-modulated entry at Mars and Earth to lower the peak acceleration, augment
the trajectory precision, and increase control.
Figure
57: Trade space for EDLA missions (Larson, 1999)
The propulsive
Δv requirements for the phases of lunar and Martian EDLA are summarized in
Table
17.
Table 17: Propulsive Δv requirements for Martian and lunar EDLA
The Martian
de-orbit Δv involves
transferring from a 500 km circular orbit to a 500 km × 20 km Mars transfer
orbit. The initial descent phase entails
aeromanuevering and parachute deployment that does not require any propulsive
Δv. The powered descent phase includes ± 4.5 km
lateral translation capability for dispersion accommodation and landing target
site redesignation. The lunar de-orbit
maneuver is from a 100 km circular orbit to a 100 km × 17.5 km transfer
orbit. The powered descent and landing
phase includes initiation, braking, pitch-up/throttle-down, and vertical
descent to surface. The Martian and
lunar ascent/rendezvous Δv
include boost and circularization at 500 km and 100 km altitudes,
respectively. To build a sustainable
EDLA architecture for the variably sized Mars and Moon missions, the Lander’s
functionality requirements for each mission type are integrated in Table 18. The
commonality with the Lander required for the Earth return portion of each
mission is shown in the figure as well.
This architecture design has deliberately chosen Lander designs such
that significant portions of the designs are common amongst all missions.
Table
18: Integrated Lunar and Martian Lander
functionality requirements
The commonality
shown between Lander designs in Table
18 allows for the leveraging of at least a portion of
engineering design, manufacturing, and testing costs across all missions.
Although
significant commonality between Lander designs has been purposely designed for
this space exploration architecture, several differences between Lander designs
do exist. These differences are due to
the significant differences between the landing environments on the Moon and
Mars. The differences in ∆V requirements and Lunar and Martian
atmospheres play major roles in determining the final Lander designs for each
destination. Schematics for Earth
return, Lunar, and Martian Landers are shown below in Figure 58, Figure
59, and Figure
60.
Figure
58:
Earth return capsule design
Figure
59: Lunar Lander design
Figure
60: Martian Lander design
The three Lander
designs can be seen to have some identical components, some similar components,
and some differences. For example, the
crew capsule on top of each of the Landers is identical. The engines for the various Lander designs
are similar to each other. The actual
engine masses can be seen in Table
19 in the next section.
The differences among the Landers are few. A major difference is the requirement for
floatation devices on the Earth return capsule.
This is not required for the other two Lander designs. In addition, the Earth return Lander will
require a parafoil while the other two Landers do not. Finally, the Martian Lander has increased
complexity due to the fact that it has a deployable heat shield and landing
structure. This is not required on the
other Landers.
A major decision
made during the Lander design for this project was to determine how many
crewmembers the Landers should be designed to accommodate. If missions with varying numbers of crew
members are going to be done, perhaps it would be better to have greater
numbers of Landers that each accommodate fewer crew members. On the other hand, if most or all missions
will be done with the same crew size, it may be beneficial to simply design the
crew compartment to be common for all missions.
Since three and
six-person crew sizes are being considered for this project, a trade study was
proposed to weigh the benefits and drawbacks from designing a three-person
Lander versus a six-person Lander. For
missions with a crew of six, two three-person-sized Landers would be required
to transport the crew instead of one Lander to accommodate the entire crew.
Table 19 below is a comparison of the Lander component masses
required for Earth return, lunar, and Martian Lander designs for three and
six-person crews. An important
assumption made was that the Landers have life support systems capable of
supporting the crew for several days.
Table
19: Three and six-person Lander component mass
comparison
|
Earth Return |
Lunar Lander |
Martian Lander |
|||
|
3-person vehicle masses
(kg) |
6-person vehicle masses
(kg) |
3-person vehicle masses
(kg) |
6-person vehicle masses
(kg) |
3-person vehicle masses
(kg) |
6-person vehicle masses
(kg) |
Crew
capsule |
3617 |
4281 |
3541 |
4411 |
3541 |
4411 |
Parachute |
225 |
450 |
|
|
410 |
425 |
Parafoil |
275 |
300 |
|
|
|
|
Heat
shield |
438 |
573 |
|
|
950 |
986 |
Landing
structure |
|
|
40 |
55 |
50 |
65 |
De-orbit
stage |
95 |
124 |
50 |
62 |
436 |
535 |
Descent
stage |
|
|
3942 |
4915 |
2049 |
2553 |
Ascent
Stage 1 |
|
|
2319 |
2889 |
3430 |
4272 |
Ascent
Stage 2 |
|
|
|
|
4067 |
5066 |
Total mass |
4650 |
5728 |
9893 |
12332 |
14932 |
18318 |
Total mass/crew |
1550 |
955 |
3298 |
2055 |
4977 |
3053 |
The last row in Table 19 makes the best argument for simply designing one
six-person Lander if crews of only six are to be used for all missions. This is because the mass per crew for
six-person-sized Landers is significantly lower than that of the three-person
Landers. However, if a crew of three
were to use a six-person Lander, this benefit would be lost. In fact, a significant mass penalty would
result from a crew of three using a six-person-sized Lander.
A benefit from a
crew of less than six people using a six-person Lander would be the option of
using the extra internal volume to bring additional life support and other
equipment for either enhancing mission capabilities or simply providing extra
redundancy to provide additional safety margin for the crew.
For a six-person
crew, a major benefit for using two three-person Landers as opposed to using
one six-person Lander is the ability for a portion of the crew to survive and
complete their mission even if there is an accident in which a major Lander
failure results in the loss of that portion of the crew. If a crew of six is landing in one Lander and
the Lander fails and the entire crew is lost, that would effectively end the
mission and it would be deemed a failure.
Having two Landers adds redundancy to mission success by removing a
single-point failure. If a remaining
crew of three were able to successfully complete their mission and return home,
the mission will most likely be deemed a success.
Many of the lunar
and Martian mission architectures are dependent on the pre-positioning of cargo
on the planetary surface. Indeed, a
major objective is to separate cargo from crew as much as practicable for
reasons of safety and cost-benefit. For
the crewed phases of the mission, the Landers have been assumed to be partially
controllable by the astronauts. For
instance, on Apollo 11 Neil Armstrong took manual control of the powered
descent system when he noticed the target landing site was in a boulder
field. Unfortunately, unmanned cargo
Landers cannot rely on such human situational awareness and adaptability. The autonomous landing systems for these
Landers will require a high degree of accuracy to position cargo modules in
close proximity, while avoiding terrain hazards and other modules. NASA/JPL Mars Exploration Rover’s Spirit
successfully landed within its 77 km long footprint at a distance of only 9 km
from its dead-center target. NASA has
renewed focus on autonomous descent and terminal guidance. As part of NASA’s New Millennium Program, the
“Smart” landing technology capabilities roadmap is summarized in Figure
12. The roadmap matches well with the
time-scale of developing the Moon and Mars exploration system.
Figure
61: NASA’s missions and “smart” landing
technologies roadmap (Thurman, 2003)
On-orbit assembly
is a difficult task in more ways than one.
Choosing a launch vehicle, selecting the timeline and determining the
launch logistics for assembly are three difficulties of using this method of
module assembly.
According to
(Larson, 1999), four considerations are important when designing a launch
sequence,
From a
transportation perspective, it was assumed that the space transportation system
architecture does not require the use of the International Space Station (ISS)
as an assembly or return point. This was
done to ensure that NASA could divest itself from the ISS and STS to meet the
Space Exploration goals within the future budget cap. When pre-positioning becomes paramount to a
mission architecture, the assumption of vehicle on-orbit assembly is critical
to mission success.
On-orbit assembly
prior to crew arrival increases the safety of the mission. Once the transfer vehicle is assembled and
the crew is delivered to orbit, there will be some on-orbit
"check-out" functions that the crew will need to accomplish to
complete assembly, verify functionality, and prepare for injection to the
destination. This may require a few days
during which the crew would be exposed to a high radiation environment. Launching a crew to a fully assembled vehicle
ensures that radiation exposure is minimized.
This also ensures that the vehicle has been correctly assembled and
checked over prior to crew arrival. If any problems occur during
assembly, the vehicle is more readily accessible, and if problems occur during
check-out the crew can return home easily.
Spacecraft
platforming, reconfigurability, extensibility, and assembly are discussed
elsewhere in the Report.
The lunar baseline
mission architectures were evaluated for two trajectories: one using lunar
orbit at a staging point, as discussed in the lunar baseline mission
description, and one using the Earth-Moon L1 (EM-L1) rather than lunar orbit as
a staging point. The different trajectories were compared in terms of total
mission mass in LEO, and the results are shown in the figure below.
Figure
62: Comparison of Mass in
LEO for Different Missions
For the purpose of
comparison, all missions segments use the minimum Dv trajectories, employing a Hohmann transfer
where appropriate. Also, calculations assume a lunar parking orbit of 100km
where applicable, and include additional Dv required to establish and leave EM-L1 halo
orbits where applicable. The appropriate parameters to reproduce this
calculation are summarized in Appendix 9.3.
The results shown
for architectures using lunar orbit apply for lunar missions to any latitude
landing site if a free-return trajectory constraint is not imposed. They also
apply to lunar equatorial landing sites if a free-return trajectory is imposed.
The results shown for architectures making use of EM-L1 apply for missions to
any latitude landing site, regardless of free-return trajectory constraints.
Because the use of a trajectory utilizing the EM-L1 consistently results in a
greater mass in LEO, the baseline mission uses lunar orbit as a staging area.
However, this analysis does not include consideration such as accessibility to
high latitude landing sites when free-return trajectory or scheduling
constraints are imposed. These issues are dealt with separately in Section 6.4.3.2.3.
A list of suggested
landing sites is shown in Table
20.
Table 20:
Suggested landing sites
There are three
primary factors influencing whether to use lunar orbit or the Earth-Moon L1
point in a lunar mission architecture: choice of landing sites, free-return
option, and scheduling.
If lunar missions
are targeting sites within plus or minus five degrees of the equator, an
architecture utilizing a lunar equatorial orbit provides the option of a
free-return trajectory. In addition, it
offers the ability to descend from and ascent to lunar orbit every two hours.
Lunar missions that
do not target equatorial sites can achieve intermediate inclination and polar
lunar orbits by making minor targeting maneuvers early in the trajectory. This
targeting requires a negligible amount of Dv, however it removes the possibility of a
free-return trajectory. Making use of intermediate inclination and polar orbits
also introduces scheduling constraints. For example, from an intermediate lunar
orbit, a spacecraft can only descend, ascent, or enter a return to Earth
trajectory every 27 days. From a polar orbit, a spacecraft can descend to or
ascent from an intermediate or equatorial landing site every 14 days, and can
descend to or ascend from a polar landing site every two hours. An opportunity
to enter a return to Earth trajectory from polar orbit occurs every 14 days
(Larson, 2002).
If a lunar mission
is targeting a non-equatorial landing site and a free return trajectory is
deemed necessary, two possible solutions exist. The first solution is to make
use of the Earth-Moon L1 (EM-L1). The second solution is to enter a lunar
equatorial orbit and then initiate a propulsive maneuver to change orbital
planes.
The EM-L1 provides
access to all lunar landing sites with the option of a free-return trajectory
if no burn is initiated at the EM-L1, and the continuous ability to descend
from, ascend to, and enter a return to Earth trajectory using the EM-L1.
However, an architecture utilizing the EM-L1 requires an increase in total
mission DV of
approximately 11% (plus or minus 2% depending on the trajectory used) and four
extra propulsive burns as compared to an architecture using lunar orbit.
The second option,
entering a lunar equatorial orbit and then initiating a propulsive maneuver to
change orbital planes, is expensive in terms of total mission D v. If the lunar mission is targeting a
landing site no more than 39° from the lunar equator, then the total increase in D v is less than that required to use the
EM-L1. However, if the landing site is greater than 39° from the lunar equator, using the EM-L1 is
less expensive in terms of D v.
The advantages of
using lunar orbit are clear for missions targeting equatorial sites, but the
best option is not clear for missions targeting non-equatorial landing sites.
Unfortunately, as described in the discussion of landing sites, many
interesting sites are not located along the equator. Thus, mission planners
must weigh the importance of a free-return trajectory and scheduling
constraints in terms of increased DV to determine the appropriate architecture.
While total mission
Dv provides a general means of comparing lunar
orbit and EM-L1, mission architecture can play a significant role in
determining the influence this metric has on the total mission mass in LEO. For
example, Figure
63 shows the total mission mass in LEO assuming a
free-return trajectory requirement for the manned segments of a mission to a
lunar pole. While EM-L1 may be beneficial in terms of total mission Dv if the mission is targeting a polar landing
site such as the
Figure 63:
Mass in LEO for mission to lunar pole with free-return trajectory requirement
The proximity of
the Earth and Moon, and the likelihood of multiple lunar missions to serve as
test beds for future Mars missions raise the question: what benefit, if any,
can be accrued through use of a reusable Lunar Lander?
A trade study was
performed to compare a non-reusable Lander to a reusable Lander in terms of
cumulative mass in LEO required to pre-position and ready one three-person
Lander in lunar orbit. The results are shown in the figure below.
Figure
64:
Comparison of a non-reusable and reusable Lunar Lander
The mass in LEO for
a non-reusable Lander was calculated assuming electric propulsion to
pre-position one Lunar Lander with enough chemical fuel to descend to and
ascend from the lunar surface. The mass in LEO for a reusable Lander was
calculated assuming electric propulsion to pre-position one wet Lunar Lander
for the first mission. Mass in LEO of subsequent missions was calculated
assuming electric propulsion to transport fuel for the pre-positioned Lander’s
descent and ascent. Because a reusable Lander will likely have a greater mass
than a non-reusable Lander, a mass penalty was included in the study. For
example, a 30% mass penalty means the calculations shows that an expendable
Lunar Lander has a mass 30% greater than the non-reusable Lander mass.
Figure 64 shows that if a reusable Lander can be designed with
only a 30% mass penalty, then a reusable Lander is beneficial after only two
uses. However, if the mass of a reusable Lander is nearly double the mass of a
non-reusable one, then five missions are necessary before benefit is accrued.
While Figure 64 suggests that a reusable Lunar Lander provides
benefits in terms of mass in LEO, this study does not include the mass in LEO
for tools and extra equipment to service the Lunar Lander when necessary. Also,
mass in LEO as a comparison metric does not capture all cost associated with
designing and building a reusable spacecraft; there are additional development
and operational costs not captured in this study.
A human mission to
Mars is an enormous undertaking that poses significant design constraints. There are many factors that have influenced
the baseline mission and many more that will emerge. In order to understand these design
constraints, a deeper investigation into the factors that influence the
trajectory and the mission must be conducted.
The Mars environment
will directly affect the design of any mission to the planet. Therefore, it is necessary to examine the
environmental attributes and determine how they impact the design.
Martian seasonal
change can drastically affect in weather and the working environment. The northern hemispheres’ spring is 94 sols
long and autumn is 143 sols. The Martian day, or sol, is 24 hours and 39.6
minutes long. Gravity on Mars will
affect all activities on the surface as well as the health and well-being of
the crew. The gravity on Mars is approximately 3.758 m/s2, which is
slightly more than one-third of Earth’s surface gravity.
Landing site
selection will depend on a number of factors; altitude and site characteristics
will be paramount. Altitudes range from
+27 km on Olympus Mons to –4 km in
Three types of
planetary surface material exist: rocks, regolith, and fines. Mars’ surface material can shield against
radiation and micro-meteoroids. Thus,
about 0.5 m of Martian surface materials should be enough to stop the primary
dose from solar particle events, although you could use several meters to
prevent more radiation.
Previous excursions
to Mars, such as VL-1, VL-2, and Sojourner spacecraft exceeded operation
lifetime. Thus, planners should not
expect long-term chemical degradation if the system is properly designed.
Atmospheric
conditions will have the biggest impact on mission design. Weather and atmospheric composition will
drive all surface operations. Wind
speeds vary by season and are lowest during the summer (around 2 m/s – 7 m/s),
and reach speeds around 5m/s – 10m/s during autumn. Global dust storms, which can reach speeds of
30 m/s, occur in the southern spring and summer (
The severity of the
Martian temperature will have a major effect on both mission design and
planetary surface operations. The average
recorded temperature on Mars is -63° C (-81° F) with a maximum temperature of
20° C (68° F) and a minimum of -140° C (-220° F). Barometric pressure varies at each landing
site on a semiannual basis. Carbon dioxide, the major constituent of the atmosphere,
freezes out to form an immense polar cap, alternately at each pole. The carbon
dioxide forms a great cover of snow and then evaporates again with the coming
of spring in each hemisphere. When the southern cap was largest, the mean daily
pressure observed by Viking Lander 1 was as low as 6.8 millibars; at other
times of the year it was as high as 9.0 millibars. The pressures at the Viking
Lander 2 site were between 7.3 and 10.8 millibars. In comparison, the average
pressure of the Earth is 1000 millibars.
Thus, the comparatively low barometric pressures will pose serious
design constraints and limitations.
The Mars
environment will direct the design of any mission to the planet. Thus, having examined some of the environmental
factors, it is necessary to understand how they impact the mission design. .
Examining the
environmental conditions given above we can infer some general
conclusions. The soil can support
Landers, stations, and rovers, however footpad and wheel dimensions must be
sized according to load. Structures can
be anchored in the soil for additional stability against seismic events, wind,
etc. In addition, the soil can be used
to provide some level of radiation and environmental shielding. The effective shielding against radiation is
0.5 m to 3 m. The effective shielding
against micrometeoroids and orbital debris is 0.5 m.
The design of
equipment must account for the harsh environment and atmospheric conditions on
Mars. The physical effect and chemical
effect of soil and dust on mechanical and electrical systems is unknown,
however mechanical devices will need lubrication and sealing.
Using solar energy
to provide a power supply is advantageous on Mars. Mars has longer days and no significant eclipses. The surface can be used as an electromagnetic
ground. In addition, the local geology
contains usable quantities of critical resources such as CO2 and
water from the atmosphere (Larson, 2003).
There are certain
criteria for landing site selection that is dictated by both the mission
objectives and the planet’s environment.
The latitude should be between 30-600S for power constraints
and poleward of 300 for near surface water. Positioning the landing site near fluvial
activity, where subsurface water may be expected (i.e. gully locations), sites
of past seismic activity. The elevation
should be at a maximum of 1.3 km above the Mars’ datum. The site should be a smooth flat plain,
relatively devoid of large obstacles and the landing ellipse must fit within
the flat plateau region.
The following is a
list of potential sites that satisfy the above criteria:
• Dao Vallis (33°S,
267°W)
• Gorgonum Chaos
(37°S, 168°W)
• Nirgal Valles
(30°S, 39°W)
• Elysium Planitia
(37°N, 252°W)
•
Although a human
mission to Mars may be accomplished without developing new technologies, to
develop a sustainable initiative, it will be quite beneficial to investigate
some or all of these technologies. Some
of these technologies will allow for a lower mass to be launched from Earth,
while others will allow for a semi-permanent infrastructure on Mars to remain
relatively self-sufficient. A review of
some of the important technologies follows.
In-situ propellant production (ISPP) promises to be a
viable new technology that will have a significant effect on missions to Mars
and the corresponding mission architectures.
The Mars Direct plan developed primarily by Robert Zubrin in the late
1990’s introduced a realistic ISPP scheme, which makes use of a readily
accessible resource: the Martian atmosphere (Zubrin, 1996). ISPP would allow propellant to be produced on
Mars for the ascent to Mars orbit or even for the return journey to Earth, thus
dramatically reducing the required IMLEO.
Before the Mars Direct proposal, mission plans estimated the IMLEO to be
on the order of 1500 metric tonnes or ten times that of the Apollo
missions. By utilizing ISPP, the IMLEO
estimate is reduced to 250 tonnes for Mars Direct and 450 tonnes for a slightly
scaled up version, Mars Semi-Direct, with a crew of 6, as presented in NASA’s
1997 Design Reference Mission (Heidmann, 2003).
The ISPP method, which produces methane/oxygen (Isp
~370s), as described by Zubrin, is still the most widely recognised possibility
for Mars. This method entails sending
hydrogen feedstock, a small nuclear power plant, and a chemical processing
plant to Mars in advance of the human crew.
The ISPP plant is autonomously set up to begin fuel production.
The production of fuel is performed by the Sabatier
process, which converts carbon dioxide from the Martian atmosphere into methane
and oxygen via the following reaction:
4H2 + CO2 à CH4 + 2H2O.
The resulting water
is electrolysed into hydrogen and oxygen.
The oxygen is stored and the hydrogen is recycled back into the Sabatier
reaction. This series of reactions
provides a mass leverage (produced fuel mass to imported fuel mass ratio) of
12, which can be increased further by simply improving the methane/oxygen molar
ratio to that required for propellant, by utilizing a Reverse Water-Gas Shift (RWGS) to generate
oxygen as follows:
CO2 + H2 à CO + H2O.
Using this process,
the mass leverage is increased to 20 (Heidmann, 2003). Yet another option is to use RWGS to produce
methanol that has a lower Isp than the methane propellant but uses a
more efficient process and does not require cryogenic storage. Since methanol-based propellants have a
relatively simple engine design, such a fuel may be very useful for surface
power such as rover fuel cells (Heidmann, 2003).
The development of
ISPP technologies requires several obstacles to be surmounted. Considerations of the Martian environment,
such as dust and reduced gravity will affect the functioning of the chemical
processing plant and nuclear power plant.
The design of these components will have to be carefully implemented
such that these factors are taken into consideration. Secondly, ISPP will rely heavily on
autonomous systems. A spacecraft will
have to survive the hazards of launch and interplanetary travel, land
successfully, deploy its various components, and begin to produce power even
before the chemical processing plant can begin to function. Furthermore, autonomous verification and
communication of the status of propellant production will be required to
validate the functioning of ISPP. The
above obstacles suggest that the first human mission to Mars not rely on ISPP,
but instead verify and test the technology such that future missions can
benefit from this technology.
In-situ resource utilization (ISRU) has the potential advantage of producing a significant savings in both mass and cost for Mars missions. Concerns about developing in-situ resources include technological readiness of many processes, their safety and reliability, and their sometimes serious effects on mission design. However, with the idea of developing a sustainable exploration, some examples of potentially advantageous ISRU are given below.
First, the use of solar radiation for power generation could provide a sustainable surface infrastructure with a low cost, reliable power supply. Also, the low gravity and near-vacuum environment is good for most material processing. To this end, carbon dioxide components can be used not only for life support, but for the creation of plastics, which allows for the return to commercial beneficiaries. Nitrogen from the atmosphere will be added to the air in the life support systems and as well as the soil in a greenhouse. The soil itself can be used for radiation shielding, and can be melted and sintered for the purpose of construction. Any metal extracted from the soil can be used in construction (Larson, 2003). In addition, water ice permafrost can be extracted from the soil to be used in propellant as well as life support. These in-situ resources can be utilized to enable a low cost, semi-permanent infrastructure on Mars.
To support a long-duration mission, it will be necessary to recycle
resources from the life support system.
The closure-level of a life support system is the percentage of waste
products that are recovered as useful resources. A high closure level requires less re-supply
but may also add cost for technology development, increase power requirements,
and increase complexity.
A life support system must provide resources for meals, hygiene, medical
activities, and science experiments. The
primary consumables required for astronauts are water, oxygen, and food. Four major life support functions exist: 1)
managing the atmosphere to maintain nitrogen-oxygen pressures, provide a
comfortable temperature and humidity, and remove contaminants such as carbon
dioxide “bubbles,” 2) distributing water to the crew and processing waste
water, 3) treating waste and recycling consumables, 4) produce and process
food. For missions longer than a few
weeks, regenerable technologies will be utilized to remove CO2 and
recover wastewater. For missions on the
order of several months, it will be necessary to also employ
oxygen-regeneration technologies.
In general, food, water, and air represent most of the mass of a life
support system. Closed-loop life support
technologies (also known as advanced life support or ALS) can significantly
reduce water and clothing masses, where the clothing mass is reduced at the
expense of an aqueous laundry. This is a
viable trade for long-duration missions (
One element of the trade space for closed-loop life support systems is
whether or not to employ such a system for the pressurized rovers that
astronauts will use to travel hundreds of kilometers on the surface of the Moon
and Mars. Pressurized rovers provide a
shirtsleeve environment with a breathable atmosphere—oxygen is provided and CO2
and water vapor are removed. It is
possible to reclaim all CO2 and water vapor for recycling. However, the mass and power requirements of
such a system may exceed the mass of the consumables saved on any one excursion or even multiple trips. Given that roughly 1000 kg of equipment is
required to recover 1 kg of nitrogen and oxygen per crewmember per day, nearly
50 week-long
surface trips by three astronauts would be required to gain mass savings from a
closed life support system in a pressurized rover. Whether utilized in rovers, or just in the
stationary surface habitats, closed life support systems allow for a
semi-permanent infrastructure to be sustained.
Nuclear Propulsion is a relatively well-developed area of research that can provide significant mass savings in LEO. The safety concerns of launching a nuclear powered engine must be balanced with the benefits of having a high specific impulse form of propulsion.
Similar to electric propulsion, nuclear propulsion can provide constant thrust, unlike chemical propulsion. However, the thrust is significantly higher than electric propulsion and allows for competitive transit times. This decreases the major concern involved with the use of electric propulsion, which is the significant transit times during the outward spiral through the Van Allen belts. A nuclear propulsion engine can provide a specific impulse on the order of 960 sec (Walberg, 1993).
Figure
65: Comparison of nuclear propulsion to chemical
propulsion for baseline trajectories
Assuming that a
nuclear propulsion engine is employed for the human travel to and from Mars and
that pre-positioning is still provided by electric propulsion (since flight
times are less of a concern), Figure
65 shows that a significant decrease in initial mass in LEO exists when
nuclear propulsion is utilized. Thus, to
develop a truly efficient and sustainable transportation system to Mars, and
beyond, nuclear propulsion will provide a low mission mass cost, and therefore,
warrants further consideration (Walberg, 1993).
Power requirements
for a Mars mission are projected to be between 25 kW-100 kW for an initial
surface habitat, 10 kW for a piloted rover, and 160 kW for a habitat that
process in-situ resources. Any
sustainable exploration architecture must meet these requirements in an
economical manner. Power systems on Mars
face a variety of challenges: 12.3 hour nights, variations in the day/night
cycle by season and latitude, and low temperatures (especially, during the
night and winter). Atmospheric dust
poses a threat to solar arrays on long-duration missions. In such cases, scrubbers may be needed to
limit degradation. The primary challenge
to power systems on the surface of the Moon is the 354 hour night. Other design issues arise from lunar dust
that is produced from normal surface operations, structural issues in a
one-sixth Earth g environment, and daytime temperatures, which can exceed 100˚C
(Landis, 1999).
The primary
electrical power source on the surface of the Moon and Mars may come from a
combination of solar cells, batteries, radioisotope thermoelectric generators
(RTG), or nuclear reactors. Photovoltaic power sources provide a long-term
source of power at an estimated degradation rate. For relatively short manned missions in the
past, energy storage systems have proved sufficient to power habitual modules
(e.g., Mercury, Gemini, Apollo, STS).
Nuclear power sources provide a long-term source of power and are
appropriate for missions operating with little sunlight (e.g., polar regions of
Mars, lunar nights).
The principal metric for evaluating solar and nuclear power sources is
specific power (watts per kilogram).
Photovoltaic specific power ranges from 25-250 W/kg while RTG specific
power ranges from 5-20 W/kg. For a Mars
mission requiring 100 kW of power, this translates to 400-4000 kg of solar
arrays or 5000-20000 kg of RTG. However,
nuclear power outperforms solar cells in terms of radiation hardness,
stability, maneuverability, shadowing sensitivity, and obstruction of
view. In fact, the
Nuclear power offers the greatest flexibility for human solar system
exploration because of its utility in virtually any space environment. Given the likelihood of extended Martian
missions and the construction of a lunar base that will require electricity
during the long lunar nights, photovoltaic sources are not optimal.
One option that would meet initial power requirements and provide plenty
of room for extensibility is the deployment of a nuclear reactor. NASA is developing a nuclear reactor called
the SP-100, which provides 825 kW at a specific power of 41 with a lifetime working
at full power of around 7 years. Baseline
versions of the SP-100 provide 100 kW with a specific power of 30 (3000 kg),
50% less mass than the best RTG projections.
With nuclear reactors, it is necessary to shield the crew from
radiation. This shielding makes up a
significant fraction of the power systems overall mass. Lunar or Martian soil may be employed to
reduce the shielding materials transported from Earth. Launch pad safety is of utmost importance
when dealing with nuclear reactors and it may be necessary to launch the
reactor in an un-powered state to minimize radiation. Of course, obtaining permits and completing
environmental impact statements for nuclear reactors add cost and may delay
program schedules.
For redundancy and safety, it is optimal to have a back-up power system. Solar arrays could accomplish this task. Although they would not provide all of the
power necessary for the entire spacecraft, they would provide enough
electricity for emergency life support.
In space, the more
material you discard, the more you have to bring with you, thus the longer the
trip, the more Herculean the logistics requirements become. Human missions to Mars will most certainly
have one main design characteristic: a closed-loop life support system that includes
plants and microorganisms (bio-regenerative).
The most likely way to grow plants on Mars is the design of a
greenhouse.
Significant
attention must be paid to the rather extreme conditions of Mars. These conditions include the soil chemistry,
the gravity field, radiation, and temperature.
Special kind of glass that can allow in light by filter harmful
radiation will be required. An
alternative to this would be to utilize solar arrays to provide power for the
lights inside the greenhouse and to generate the temperatures necessary. In addition, extra nitrogen, along with
various other chemicals will be added to the regolith (Cowing, 2002).
Along with
providing crew with food, the plants would provide a secondary use as natural
carbon scrubbers for breathable air. This technology is highly capable and is
easily testable on Earth and on the precursor missions to the Moon. Due to infrastructure cost, this technology
would be more useful for the medium and extended missions, and thus will
increase in importance as we transition to a semi-permanent infrastructure.
In order to develop
a baseline mission scenario, both the architecture of the mission and the
choice of trajectory must be considered.
When referring to the mission architecture, we are investigating the
sequencing of events and the main form to function relationships that will
guide the mission design. To compare
different mission architectures a nominal trajectory is assumed. For an accurate comparison, all specific
impulses are assumed to be 425 sec, and all engines have a structural factor of
0.1. In addition, all pre-positioning is
performed assuming electric propulsion with a specific power of 150 W/kg, an
efficiency of 0.7 and a specific impulse of 3200 sec.
Figure 66,
shows the comparison between different mission architectures, assuming a
short-stay mission. For each mission
architecture, all maneuvers at Mars are powered, and a direct entry at Earth is
assumed. In addition, all missions
assume a surface habitat is pre-positioned.
The following is a brief description of each mission according to the
labels given in Figure
66. A NOVA mission refers to a
direct entry at Mars without orbit insertion.
A Mars orbit rendezvous (MOR) is an Apollo-class architecture in which
separate Landers are carried from Earth.
The transfer vehicle enters Mars orbit and the crew uses the Landers to
travel to and return from the Martian surface.
The MOR2 and MOR3 architectures have a similar design, but allow for the
pre-positioning of the Landers and pre-positioning of the Landers and return
fuel, respectively. Figure 66
confirms that a Mars orbit rendezvous, with pre-positioning of the Landers,
surface habitat, and return fuel is the most mass efficient way to perform the
transfer.
Figure
66:
Initial Mass in LEO for Various
MOR3 was chosen as
the baseline architecture to decrease the IMLEO. This allows for a more robust design that can
carry greater payload and produce a more sustainable transportation system
architecture. However, since MOR2 has
only a slightly higher IMLEO, if through testing on the Moon, or for just
safety concerns, pre-positioning of the return fuel is eliminated for the first
few Mars missions, then MOR2 could easily replace the baseline architecture
with few changes to the rest of the baseline mission design. However, for a sustainable architecture,
pre-positioning of the return fuel will provide a significant benefit as the
missions progress, and therefore is chosen as the baseline design.
Having shown that a
Mars orbit rendezvous with pre-positioning of both the Landers and the fuel is
a mass-efficient architecture, different trajectories were evaluated. In order to compare trajectories, a baseline
architecture was assumed for each comparison.
For each trajectory, the Mars Transfer Vehicle (MTV), with crew, departs
from LEO, transfers to Mars, and injects into a 1-sol Martian orbit. The crew transfers into two pre-positioned
landing vehicles and descends to the Martian surface. Once on the surface, the crew will live in a
separate habitation, which has been pre-positioned with the relevant
consumables for the surface stay. At the
end of the surface stay, the landing vehicles ascend and dock with the transfer
vehicle. In addition, the transfer
vehicle docks with the return fuel. The
crew returns to Earth in the transfer vehicle, leaving behind the landing
vehicles in Martian orbit. Upon Earth
return, the transfer vehicle injects into Earth orbit and the crew transfers
into two Earth return capsules. In order
to accurately compare different trajectories, the specific impulses of all
chemical engines are assumed to be 425 seconds, the structural factor is
assumed to be 0.1, and the payload to initial mass factor for aerobraking is assumed
to be 0.15. All pre-positioning is
accomplished by an electric propulsion engine with the same characteristics as
listed in Section 6.4.3.1.4.
Figure 67
compares an Opposition-class mission with and without a Venus fly-by maneuver
during one direction of transit. In
addition, both types of missions are compared with and without utilizing
aerobraking. As can be seen in Figure 67,
without a Venus fly-by, the opposition class mission, even employing
aerobraking at both Earth and Mars is prohibitive, for the above described
mission. The addition of the Venus
fly-by maneuver lowers the required initial mass in LEO (IMLEO) to a reasonable
value, especially with aerobraking, for only a small increase in time of
flight. Other issues remain in question
for this maneuver, such as radiation exposure due to the closer pass to the
Sun. However, from a strictly mass
related standpoint, a short-stay mission is not possible without a Venus fly-by.
Figure
67:
Comparison of Opposition-class mission with and without a Venus fly-by
Figure 68
displays a comparison between different trajectories for an extended-stay
mission. As we can see in Figure 68,
aerobraking significantly reduces IMLEO for each mission. An all-propulsive maneuver yields a
prohibitively high IMLEO for a fast transfer; however employing aerobraking
makes this class of missions feasible.
Figure
68: Comparison of Conjunction-class missions
According to Figure 68,
the time of flight for the fast transfer is only slightly shorter than that of
the regular conjunction class mission, and requires a greater IMLEO. However, the travel time is significantly
shorter for a fast transfer then for a conjunction class mission. Since the travel time is important for other
considerations, such as zero gravity and radiation exposure during flight, and
since when aerobraking is employed, the initial mass in LEO is only slightly
higher than that of a regular conjunction class mission, the fast transfer is
chosen as the trajectory for an extended stay mission.
Figure 69,
summarizes the IMLEO requirements for the feasible mission designs: Opposition-class trajectory with Venus
fly-by, and fast transfer. Once again,
we see that aerobraking yields a significant reduction in initial mass in
LEO. If we refer to Figure 67 and
Figure 68,
we notice a large difference in the amount of IMLEO for a manned Mars
mission. However, if we include the
IMLEO for all pre-positioned elements, as in Figure
69, once aerobraking is employed, the IMLEO is virtually the same for
both a short and extended stay mission.
In addition, the use of aerobraking and parachutes, followed by a
powered touchdown for the Landers, reduces the amount of fuel that they must
use, and therefore further reduces the initial mass in LEO.
Figure
69:
Comparison of Mars trajectories
By analyzing the
Mars trajectories, we have concluded that a Mars orbit rendezvous (MOR) with
pre-positioning is essential for reduced mass in LEO and is thus the baseline
architecture for all other comparisons.
For a short-stay mission, a fly-by of Venus is essential for a
reasonable IMLEO for a human Mars mission.
Since this maneuver is complicated and potentially exposes the crew to
significant increases in radiation dosage, as well as other complications, the
human factors elements must be studied in greater depth. By examining the Mars trajectories for an
extended stay, the increase in IMLEO for a fast transfer is abated by the
significantly lower travel time, as compared to a conjunction class
mission. In addition, the reduced
microgravity and radiation exposure makes this trajectory desirable. Aerobraking has been shown to significantly
reduce the initial mass in LEO for every type of trajectory and is essential
for a manned mission to Mars. When the
short-stay and extended-stay trajectories are compared, including all
pre-positioned elements, we notice that the initial masses in LEO are
comparable, when aerobraking is employed.
IMLEO is further reduced when a parachute and powered touchdown are
utilized.
Scenario planning is a method that may be used to determine the degree to which a given system responds to changes in the environment. By proposing scenarios, many of them examples of the extreme cases, and evaluating the ways in which the system would change or would fail to change, critical contingencies may be planned into the final design. This section outlines a set of seven scenarios and the anticipated responses to them.
It is important to recognize that each design, including the baseline presented in this report, carries with it implicit assumptions about the state of the present and future environment. These assumptions about the environment constitute a de facto scenario in which the system is designed to operate nominally. Due the high degree of uncertainty, and associated high probability of change surrounding these assumptions, it is necessary that the system designed be able to adapt to changing environmental factors. Seven extreme changes in the system’s operating environment were selected as scenarios against which the baseline system’s performance could be evaluated. Scenario planning is primarily used in this fashion to identify architectural options and trades. In doing so, the system may be made adaptable or robust in the face changing environmental conditions. In this section, the baseline strategy was evaluated using each of the following scenarios. Where possible, options were exercised that allowed for significant adaptability or robustness to the constraints imposed by the scenario’s parameters. The performance of the baseline strategy upon the use of these options serves to demonstrate the degree to which this strategy is sustainable and extensible in the face of drastic change.
A foreign power successfully demonstrates a
mission to the Moon similar in style to the Apollo-11 mission of 1969. Upon
successful completion of their mission, the foreign power announces its
intention to establish a permanent colony on the Moon and to land a human crew
on Mars within the decade. The
With such a rapid influx of money with the
requirement of exploring Mars and colonizing the Moon within a decade, NASA’s
best chance for success would be realized by following the Apollo model:
develop a simple mission statement and optimize a point design. Technology development would be kept to a
minimum to meet schedule constraints. In
this case, the current prioritization of cost over schedule and performance
would change: schedule would be fixed, performance would be optimized, and cost
would be variable.
The mission outline is as follows: to meet
schedule requirements, a 70-80-ton launch vehicle will be developed utilizing
components of the STS. Money will be
poured into man-rating this new vehicle, and existing STS construction plants
and launch facilities will be utilized.
Funding for the development and operation of the International Space
Station would also increase to accelerate research into countermeasures to the
adverse impact of microgravity on human physiology. Because the mission to Mars is a priority,
all hardware developed (landers, crew transport, rovers) will be optimized for
Mars, not the Moon, and thus will most likely be over-designed for the lunar
missions.
To demonstrate colonization capabilities on
the Moon, construction of a lunar surface base would have to commence within
ten years. Specifically, getting to
the Moon, and demonstrating technologies are a priority, but not
science; therefore, the medium-sized missions will be sacrificed (see section
4.2.2) in favor of a few short missions and a long mission (lunar base) as soon
as possible, establishing semi-permanent human presence. In order to credibly demonstrate a sustained
human presence capability, the base will support developing in-situ resources
with the ultimate goal of becoming self-sufficient. Rather than selecting a fixed construction
site before detailed investigation of the lunar surface, one possibility is for
the initial surface bases to be pressurized rovers on the order of 10,000 kg
with habitation, laboratory, EVA airlock, and re-supply elements. In this case, short-stays on the order of two
weeks enabling exploration of the lunar surface would begin the colonization program;
with infrastructure build-up occurring once a site has been selected for a
permanent base. Fuel cells would power
the initial pressurized rovers; a nuclear reactor would be the primary source
of power for a permanent, fixed base.
Science operations not in support of in-situ resource utilization and
other colonization technologies will be kept to a minimum, as these are not
imperative demonstrations for a semi-permanent Moon base or for a Mars mission.
For landing humans on Mars and bringing them
safely back to Earth, a short stay mission minimizing duration and consisting
of 600 days of transit time and 60 days of surface operations via an opposition
class free-return trajectory with a Venus fly-by would serve as a baseline goal
(see section 4.2.4). To meet schedule
requirements of successful completion within a decade, the short stay Mars
mission would have to be launched within eight years. Given this short time window, all chemical
propellant will be used for propulsion, and no pre-positioning will be
used.
Although the schedule is accelerated rapidly
in this scenario to accomplish two specific objectives within a decade, it is
important to note that our overall exploration architecture remains unchanged:
a sustainable human presence in outer space.
Missions in this scenario flow from existing Moon and Mars baselines
with infrastructure increasing capabilities over time.
This
scenario illustrates a choice that must be made with regards to how extensible
the final system should be. A typical engineering situation will require a
trade-off between extensibility and optimality. In the case of a drastic shift
in
NASA’s primary launch vehicle is destroyed
during operation due to a technical failure, killing the entire crew
compliment. All usage of that particular vehicle ceases until the problem can
be isolated and fixed, a process that may take as long as five years. American
astronauts are involved in space exploration activities during the catastrophe,
thus requiring that they find another method to leave and return to Earth as
soon as possible.
A launch system failure is a terrible
situation. Due to the danger inherent in launch activities, the risk of
something like this happening is always present. Much can be done in advance to
reduce the impact that such an accident would have on the overall exploration
mission, for example, launch the crew and cargo separately. Design of lunar
transfer and Martian transfer vehicles to employ a docking mechanism and an
orbit that allows for foreign vehicles to dock on it will ensure that in the
case of a launch system failure, NASA has the option of turning to
international partners for support. If these cautious measures are taken in
advance, the impact on the exploration agenda can be minimized.
If the launch failure involves only the
heavy cargo vehicle it will be a setback, but only a minor one. In this case
the vehicle is unlikely to be grounded for more than a year. However, if the
decision where made to use the same vehicle for humans and heavy cargo and that
vehicle failed catastrophically, then the consequences would be very serious
and the leadership of the United States of America in human space flight would
be very vulnerable.
One way to incorporate robustness against
this scenario would be to create a redundant second launch vehicle design.
There are two ways that this could be done:
The first way involves heavy cooperation
with international partners who already have human launch capability.
Essentially, the
The
second way involves creating a competitive structure within American business.
In this situation, the
Nuclear propulsion technology emerges as a
viable replacement to chemical propulsion. The technology is more efficient,
can generate a higher-specific impulse, and has a higher amount of total thrust
at liftoff. While it is initially very expensive and has not yet been
flight-tested, it is expected to be approved for flight within 2 years. The
sensitivity of nuclear technology prevents its development in cooperation with
foreign nations and its use on foreign launch vehicles. If mishandles it may
cause catastrophic failure. Although extensive testing suggests that the
technology is highly reliable, the public is wary given that significant damage
may result if it is misused.
The use of nuclear propulsion would benefit a space exploration system program in several ways. First, since nuclear propulsion has a higher specific impulse than conventional chemical propulsion systems, initial mass required in low Earth orbit would be reduced for each mission using nuclear propulsion. This benefit would have a large impact on the mission design. One benefit would be an increase the amount of non-propulsion system mass in the launch vehicle. This increased payload efficiency would allow for potentially more redundancy or scientific hardware to be launched for the same launch cost. On the other hand, this reduction in propulsion system mass could also be a cost savings measure since it could reduce overall payload mass and in-turn reduce launch costs.
In addition, if high-thrust nuclear propulsion systems become available, a significant reduction in time of flight from Earth to the desired destination would occur. This would provide several benefits. First, it would minimize the exposure of the crew to microgravity by reducing their transit time. Second, this would allow the crew to spend more time at the destination than they would if they used a chemical propulsion system to transfer to the destination.
If the use of nuclear propulsion is realized and these cost reduction and human factors benefits become clear, nuclear propulsion technology will likely be incorporated into the space exploration system architecture. The incorporation of this technology depends on what stage the space exploration program is at when nuclear propulsion technology becomes viable. If nuclear propulsion becomes viable during the design phase of the program, it is likely that the initial, short missions to the Moon will still use chemical propulsion but the larger-scale missions to the Moon would test nuclear propulsion technology in preparation for its use on missions to Mars.
If nuclear propulsion technology becomes viable when the first manned missions to Mars are being launched, it is likely the technology will not be used for the first Martian mission. However, nuclear propulsion technology may still be incorporated into the space exploration architecture. For example, a precursor mission to the Moon could be used to test the propulsion technology. This would once again utilize the Moon as a “test bed” for missions to Mars. Once the technology has been successfully tested, the “Extended Stay” or “Extended Stay + Infrastructure” Mars missions could be launched using nuclear propulsion.
In conclusion, if the potential cost savings from incorporating advanced nuclear propulsion technology can be realized, the space exploration program may be able to increase mission frequency or enhance program sustainability. The reduction in the cost of launching each mission may allow for more missions to be flown for the same cost. Alternatively, a reduction in program costs would make the program more sustainable by reducing the portion of the total NASA budget consumed by the space exploration program. Politicians may be more willing to fund the program if they see more value for reduced costs.
Although the initial cost of designing nuclear propulsion modules for use in a space exploration program may be high, the cost of incorporating the new propulsion technology into the exploration program should not be large if the propulsion architecture is modular. This will allow propulsion modules to be exchanged as long as common interfaces are used. This is another exploitation of the advantages of designing an extensible space exploration system by using modular components. Generally, modularity is more difficult and more expensive to design into a system. If, on the other hand, the propulsion systems are designed such that they are “built-in”, incorporation of the benefits of this new technology may be even more expensive.
A Near Earth Asteroid impacts the Earth’s
atmosphere, exploding harmlessly over the ocean, causing noticeable changes in
weather patterns (e.g., tsunamis, storms, etc.). Scientists unanimously agree
that if the asteroid had exploded over a populated area, significant death
would have resulted. The
It is estimated that asteroid impacts occur approximately once every century, with the most recent such event taking place in Tunguska forest, Siberia in 1908. Small asteroids on the order of one meter in diameter burn up in the atmosphere with impact energy equivalent to tens of kilotons of TNT (Rabinowitz, 1998). The threshold size for asteroids capable of global disaster is believed to be ½ to 1 km, and the impact frequency for these asteroids are once every 1000 centuries on average (Rabinowitz, 1998). It is hypothesized that there are between 1000 and 2000 Earth-approaching asteroids larger than 1 km, however only approximately 100 have been discovered so far (Rabinowitz, 1998).
Current programs for asteroid (also called
Near Earth Objects, NEOs) detection and cataloging involve observations at
optical telescopes worldwide including efforts by MIT’s Lincoln Near Earth
Asteroid Research (LINEAR) Project, JPL, and SPACEWATCH at the
In the case that NASA were charged with developing
an asteroid detection and deflection system, three concerns would be of primary
importance, namely: characterization of asteroids, development of early warning
capabilities, and procedures for interception and deflection
It is likely that a good part of NASA’s
budget increase would be used to fund programs such as MIT’s LINEAR project and
the requisite telescopes for detection and cataloging. This would aid in
asteroid characterization, as well as a reduction in uncertainty of ephemeris
data that would assist in the development of an early warning program.
Furthermore, a series of small missions akin
to the NEAR project to orbit and rendezvous with asteroids could be
initiated. It is very unlikely that
these missions would be manned since the time scale would be too short to
develop the redundancy and safety assurances necessary. Also, it would be much easier to launch these
missions if the mass could be kept relatively low – not having a crew would
mean that life support systems and a return to Earth (and the required
propellant) would be unnecessary. This
way, the mass could be kept to a level were the probe could be launched on a
common launch vehicle such as a Delta IV (Smith, 2001). These missions would aid in the
characterization of asteroids as well as the development of operational
knowledge required for interception and deflection. Deflection schemes could also be tested on
such missions.
The increase in NASA’s budget would have a
beneficial effect on the exploration initiative as well as the asteroid
protection scheme. Budgets for certain
enabling technologies such as nuclear propulsion would be likely to increase
since in the case of a short warning time, it would be necessary to reach the
dangerous asteroid as quickly as possible.
Nuclear propulsion would facilitate a mission to Mars or the Moon by
significantly reducing IMLEO. The
exploration program may also be affected in that its principal destinations
would be modified. Along with the series
of unmanned missions to NEOs, a manned mission to Phobos would be a practical
alternative: both as a precursor Mars mission, and as a possible asteroid
characterization mission (since Phobos is thought to be a captured
asteroid). This would be possible since
the ephemeris data for Phobos is much more certain than that of most NEOs and
therefore sending humans to Phobos would be less risky than a similar asteroid
mission.
The shift of focus towards asteroids and
away from exploration for its own sake would have a significant effect on the exploration
program. It is possible that planned
lunar or Mars missions would be postponed while effort is diverted into smaller
asteroid rendezvous missions. On the
other hand, given the increased total budget, NASA may be able to expand its
workforce and maintain the exploration program at full strength with minor
changes such as emphasis on particular enabling technologies or destinations.
A decision on whether or not to land on
Phobos, and thus develop knowledge and technology that may be applied toward
future asteroid operations, represents the option for this scenario. On the one
hand, Phobos may not be compelling to the public compared to the allure of
Mars. Landing on Phobos may be viewed by many as a waste of resources that
could otherwise be diverted to a Mars exploration program. On the other hand,
if asteroid missions become a priority, Phobos would provide a testing ground
for these operations.
An American expedition to the Moon discovers
reserves of resources at the Lunar Poles, allowing for the large-scale
extraction of hydrogen, oxygen, and water ice. Potential rates of extraction
and production could sustain a lunar colony of 30 people indefinitely.
International interest rises, and coalition of developing and space-faring
nations proposes the development of a permanently manned international lunar
base.
The American Expedition
discovers the water resources in year 2010. This event results in an increased
interest in the Moon, and the modification of previous plans, installing a
permanent base by year 2016. International partners participate in this effort,
and the base is used largely for science and a deeper exploration of the lunar
satellite.
This deeper exploration leads to
the discovery and valuation of additional resources, which would include
additional underground water resources, He3 repositories and Rare Earth mining.
The fact that water can be cheaply dissociated using solar energy brings an
important economic value to this discovery.
The fact that some of the cargo
for the Mars and beyond missions will be shipped from the Earth and some of it
will be shipped from the Moon under this scenario increases the importance of
an EM-L1 node in Space Exploration beyond the EM neighborhood, and propitiates
a case for a Space Station parked at this point (Kent, 2001).
Using present day technologies,
the following table shows the advantage gained in sending one kilogram of
propellant to the L1 point, which is the most likely storage point for
propellant to be sent out of the Earth Moon Neighborhood.
|
Mass at Planet surface |
Mass at planet orbit |
Mass at L1 |
Departing from Moon Pole |
2050 kg |
1200 kg |
1000 kg |
Departing from Earth Equator
and using electric propulsion |
64300 kg |
2250 kg |
1000 kg |
This advantage allows increasing
travels outside of the Earth neighborhood to be easier to afford, and at the
same time gives the Moon station a revenue case that helps to sustain the costs
of the scientific base.
Additional water resources and
increased Moon travel frequency prompt market forces to intervene, and make
feasible the establishment of a tourist’s hotel on the Moon, and eventually a
stable civil population.
By year 2022, freight to the
Moon is handled by private corporations, while NASA only operates the Moon
scientific base, and focuses its exploration efforts on Mars and beyond. Fuel
production also is handled by private corporations, which have already invested
heavily in the Hydrogen Industry on Earth.
There is therefore an increased
interest on the Moon that leads, to some degree, to its trivialization. The
exploration quest continues, but beyond the Earth neighborhood, helped by the
fact that fuel is easier and less costly to obtain, and that two lower escape
velocity nodes are present on the system: the Moon itself, with a stable
population, and the L1 point as a supply node that holds deposits of fuel and
supplies.
It is arguable that the
additional interest, and resources that the Moon base will require, could
reduce the investment rate for the Mars exploration on the short term. On the
other hand, the trivialization of space travels that this scenario could imply
will allow a cheaper and faster exploration of the remaining solar system
assets on the longer term.
The primary option associated with this scenario is the use of the EM-L1 point. If the ability to transit through this point is originally included in the lunar transit architecture it may be easier to reach the lunar poles without having to do an energy-expensive plane change in lunar orbit. If, on the other hand, a decision is made not to transit to the lunar poles, this extra option is not utilized, and constitutes a deviation from an efficient design. A similar option to be considered is the value of creating a space station of some sort at the L1 point. Doing so may allow for refueling to occur on the way to the Moon, Mars and other celestial bodies. On the other hand, stations require maintenance. One possible way to get around this constraint would be to create a station that is somewhat autonomous. In this situation, the station would not require a constant human presence, although such a temporary presence may be desirable for other reasons, such as regular routine maintenance and microgravity-related research.
Following discoveries of microbial fossils
on Mars, unmanned probes find strong evidence of one-celled life currently
existing in the Martian subsurface soil. The Public’s interest is piqued and
NASA receives a 5% budget increase to hasten Mars exploration efforts. Some
groups on Earth protest the government’s decision, stating that the Martian
biosphere should not be contaminated by human presence.
With the discovery of possible life, the
Moon missions would be de-emphasized, and the timelines for Mars would be moved
up. It is likely that more precursor-manned and unmanned mission would be sent
to Mars to further investigate the life phenomenon. The budget increase would be used to fund
these efforts.
There are currently a significant number of
missions geared towards investigating lunar resources and towards building up a
lunar infrastructure/habitation knowledge base.
Since one of the purposes of the lunar missions is to serve as technology
test beds for subsequent missions to Mars, lunar missions may be cut down in
size, with emphasis placed on those missions and technology demonstrations that
are deemed critical for enabling human life on Mars. Since the Lunar missions are to serve as
testbeds many of the systems used for Martian explorations, the exploration
system is highly flexible to this change.
Since most of the lunar missions and tests
are scheduled to occur in the timeframe of the next 20 years, the introduction
of this scenario would significantly reduce this schedule. This development
would have the effect of shifting timelines for Mars missions almost 20 years
ahead. This would have the effect of
making the projection of a Mars short-stay mission more feasible since it
relies heavily on current technology and requires relatively little technology
testing.
To placate concerns on Earth about contaminating the Martian biosphere as well as protecting any human crew from contamination, precursor missions such as the Phobos and Deimos missions and more robotic missions could increase in value. Landing site certification would become paramount, so as to ensure that the environment is not destroyed by astronaut landing activities. An ideal landing site would be close to the signs of life, but the landing site itself should be chosen so as to will not irreversibly disturb the local organisms. Further unmanned investigation would give mission planners a better idea of any contamination risk.
One trade that may be drawn from this scenario is a decision on the degree to which lunar missions will focus on activities that are not Mars-related, such as exploration of the lunar poles and of other sites of scientific interest. The fundamental decision that must be made is one of the science agenda versus the exploration agenda. On one extreme, every square meter of the Moon could be mapped and cataloged in a search for scientifically interesting phenomena that are associated with the lunar surface. On the other extreme is a situation in which NASA simply lands on the Moon as a technology demonstration before going directly to Mars in the name of exploration. A similar question may be asked of the degree to which life on Mars is studied. On the one extreme, it may warrant such in-depth analysis that humanity does not go beyond Mars for decades. On the other extreme, the presence of life may simply be confirmed for its news value before NASA proceeds to other locales. This scenario highlights a situation in which an option to explore the Moon was sidelined in favor of the option to explore the possibility of Martian life. In evaluating this choice, a decision must be made so as to ensure that the knowledge, which is most important to the key stakeholders is delivered.
Motivated by election-year debates, a
nationwide referendum reveals that the 60% of the population of the
A situation of this magnitude essentially puts a moratorium on all space exploration activities. If such an eventuality were to occur, the only way exploration activities would be reinstated requires direct intervention of the President or of Congress. In either case, this would probably have to be motivated by public demand or by outside political pressure (e.g., for foreign policy reasons). Although outside political concerns are beyond the control of NASA in any situation, the degree to which NASA could rekindle the public’s interest in exploration is directly related to the knowledge that has already been gained. If there were significant knowledge available, NASA, which is responsible for educating the public on its space exploration activities, would be able to present this body of knowledge. Like most learning, this education would probably raise at least as many questions as it does answer. Thus, the natural curiosity and inquisitiveness of many people would contribute to public support for re-instatement of the space exploration initiative. Therefore, the worst-case scenario would involve this situation occurring immediately, before any new knowledge has been gained. In this situation, very little could be done by NASA to overcome the public’s disinterest, particularly since NASA, like any other government agency, does not advertise. As NASA begins to collect more knowledge and to return more results and information to the public, it becomes less likely that the exploration initiative will be cut by public demand. In any case, NASA, as a government agency, does not advertise or push a specific political agenda. Thus any movement for space exploration would have to be independently initiated. Therefore, even in the very unlikely event that public support were to drop dramatically after a momentous act (such as the first human landing on Mars), there would be little NASA could do about it directly. It is therefore incumbent upon NASA to design a knowledge delivery system in such a way that it would prevent this scenario from occurring. One sure-fire way to do this is to keep the public’s interest high. This could be accomplished through a regular series of reports and outreach activities in which NASA discusses its most recent accomplishments. The four year long election cycle provides an ideal length of time during which NASA may set a series of major milestones, which would be designed to engage the public and to maintain interest in space exploration. For example, at the end of one four-year cycle, a CEV prototype could be flown and docked with the ISS. Four years later, the first lunar CEV could land on the Moon. Each of these major milestones could be punctuated with a series of smaller yearly milestones. At the end of the fist year, for example, the CEV concept would be revealed to the public. The year after, an unmanned mockup would be flown. The year after that, test flights would occur, culminating in the ISS docking at the end of the fourth year. Not only would this approach constantly engage the public, thereby promoting sustainability, but also it would foster regular technical progress as NASA proceeded with the exploration agenda milestone by milestone, thus also promoting extensibility.
The primary trade that must be analyzed for this scenario is the degree to which NASA should concentrate on public education, inspiration and awareness campaigns. Keeping the public constantly informed is a difficult and time-consuming process. Furthermore, it opens the doors for criticism from the public. On the other hand, it may be argued that public criticism would improve NASA’s operations in the long run, if taken constructively. Furthermore, NASA is ultimately supported by the direction of the President, who is, in turn, supported by a mandate from the public. If the public is allowed to lose interest in NASA, it is likely that funding will be cut, as the public’s priorities shift away from space exploration and towards other, more immediate concerns. Ultimately, the American people are the primary beneficiaries of NASA’s efforts and some of the prime recipients of the knowledge gained through exploration programs. If the public decides that exploration is not worth the associated cost in tax dollars, the program will be cut.
In conclusion, the following two recommendations are made to NASA:
Future
systems must be designed with sustainability in mind, ensuring maximal life
cycle value (benefit at cost), as opposed to the traditional point design
approach that optimizes missions based on a fixed set of requirements. An initial design framework has been presented
as an example of proactively designing sustainable attributes into the
exploration system. While NASA can
certainly improve the process, the key message is that sustainability is not
accidental; it must be actively pursued, and the short-term costs associated
with designing for sustainability must be accepted in order to reap the
long-term benefits.
This paper lays the foundation of a methodology for designing sustainability into space systems. It must also be stressed that a system must be sustainable throughout time. This requires that any potential design method must have the ability to be reevaluated throughout time, so that the design has the ability to react to uncertainties in the future.
Tools
such as form/function mapping, commonality mapping, scenario planning along
with formalized decision analysis, such as utility analysis and real options,
were described and demonstrated in the proposed design process. These tools are only provided as examples of
structured methods for complex system design, offering the potential for proper
valuation of nontraditional system attributes such as sustainability.
The
methodology and tools described in this paper support the design of NASA’s new
exploration system. As has been
mentioned previously, the goal of this system is to deliver knowledge to all
stakeholders. This goal must be kept
continually in mind if the space exploration program is to be successful and
therefore, sustainable.
It is the authors’ hope that the readers take away a vision that current design methods are not sufficient to meet the goals of the new space exploration initiative and that a new design methodology must be developed. The new methodology accounts for sustainability and evaluates designs based on knowledge. While the proposed design method is certainly not the only solution, it is intended to be a starting point for further improvement, and ultimately, a catalyst that enables NASA and the nation to move into the next era of space exploration in a sustained and consistent fashion.
The interface for
the CEV model is shown in Figure 70.
Figure 70: Interface used for the Excel CEV model
Figure 70 presents the interface used for the model. It is a view from the “choices” worksheet in the “architecture” Excel file. The user is required to enter the number of crewmembers and duration of the mission, and also to make choices for each of the five options:
- escape system
- habitable module
- service module
- EDL architecture
- landing site
All of the options cannot be linked together, so all of the possible combinations are shown in
Figure 71. This architecture Excel file is linked to other Excel worksheets that are responsible for updating the masses for the option chosen. These are:
- “CES history” which give numbers for some the service module, and for some habitable modules
- “escape system” which give numbers for the escape system
- “re-entry_landing” which calculates the masses required to land the corresponding crew modules.
- “scaling” which computes the scaled masses of the options: 1-combined-expendable-Apollo, and 5-Flexible-OSP/XTV.
Figure 71: Linking possibilities among CEV options and
ranking criteria and weights
For the service
module, mass numbers were found from studies or real systems. These dry masses
were not scaled depending on the number of crew and days of the mission.
For the habitable
module, the Combined-Reusable-Shuttle, the Separate-Expandable-Soyuz and
Separate-Reusable-Kliper configurations, were linearly scaled given a number of
crewmembers. The duration of the mission was not included in the scaling.
For the
Combined-Expandable-Apollo and the Flexible-OSP/XTV configurations, a detailed
scaling was performed, including the number of crew and duration of the
mission.
The reason for not
scaling all the habitable modules configurations is that we lacked detailed
mass breakdown for them.
All the existing spacecraft masses (and the breakdown and other spacecraft information) were found in one of the following references: (Wade, 2004) or (Zak, 2004) or (Braeunig, 2001).
For each system, a rank was determined according to the technique described in Section 6.4.2.4. It was a way to try to evaluate systems with the few available data while being as objective as we could. The criteria for evaluating each system (and their corresponding weights) are described on the left of
Figure 71.
A Report was
written by the Orbital Aggregation & Space Infrastructure Systems (OASIS)
titled, The Revolutionary Aerospace Systems Concepts Preliminary Architecture and Operations
Analysis Report (2002). This
Report aimed to “identify
synergistic opportunities and concepts among human exploration initiatives and
space commercialization activities while taking into account technology
assumptions and mission viability in an Orbital Aggregation & Space
Infrastructure Systems (OASIS) framework.”
This
Report provided detailed information about a proposed crew exploration vehicle
and the component mass breakdown. The
methods of analysis were explained as well as engineering design details of the
structural components and hardware.
Additional resources were used to augment the scaling analysis,
including a NASA Report titled, “JSC
Lunar Transfer Vehicle (LTV) design concept, Crew Transfer Vehicle Element
Conceptual Design Report, EX15-01-094.”
For LEO crew
transfer, the Crew Transfer Vehicle (CTV) was described by OASIS (2002) and
summarized below. This module is
designed for short sleeve environment transport from LEO to the Lunar Gateway
and back, and to transfer crews between the ISS to any other crewed orbiting
infrastructure (see Figure
72).
Figure 72: OASIS CTV Internal Layout
System requirements and mass properties have
been derived from other OASIS elements as appropriate. CTV-unique system
requirements and mass properties (e.g., for human habitability systems) have
been derived from the NASA JSC Lunar Transfer Vehicle (LTV) design concept (Crew
Transfer Vehicle Element Conceptual Design Report, EX15-01-094).
Vehicle
specifics include,
·
Normal
mission duration of 4.5 days for transfer from the ISS to the Lunar Gateway (9
day total transfer time from ISS to the Lunar Gateway and back ISS).
·
Systems
shall be sized for a 22-day extended contingency mission.
·
Crew of
four (deemed sufficient for operational requirements and mission science).
·
Vehicle
remains at ISS and is designed to travel to Moon L1 and return crew to ISS with
a one-time return to the surface of the Earth in an emergency situation.
·
Internal
volume shall be sufficient to meet NASA minimal habitable threshold
requirements of 4.25 m3/person for a 22-day mission (NASA-STD-3000,
Man-Systems Integration Standards).
·
Systems
shall meet all other human habitability and life support design requirements
specified in NASA-STD-3000.
·
Designed
for launch by a Shuttle-class launch vehicle.
A preliminary estimate of CTV system mass is
provided in given in Table
21 based on derivations from the HPM and the NASA JSC
Lunar Transfer Vehicle. Note that the
heat shield analysis was considered elsewhere.
Table 21: CTV mass estimation (OASIS, 2001)
From NASA Standards (8.6.2.1 Mission Duration Design Considerations), the duration of the mission has an overall effect on the required envelope geometry. Increasing mission duration requires a greater physical envelope to accommodate mission tasks and personal needs. Crew accommodation needs are additive, so the total required habitable volume per crewmember increases with mission duration. Guidelines for determining the amount of habitable volume per crewmember for varying mission durations are shown in Figure 73.
Figure
73: NASA Habitable Volume Standard 8.6.2.1
Initially, it was thought that his standard would provide a reasonable approximation of the total volume required for a given mission duration. However, it was thought that as the number of crew increases, the habitable volume per crewmember should decrease. Therefore, Figure 74 was assumed to better approximate the habitable volume requirements for a given crew, for a given mission duration.
Figure 74: Habitable volume for various crew sizes as a function of mission duration
The habitable volume per person for the Apollo Command Module and the OASIS CTV are shown as a reference. By comparing the masses of various crew modules (Gemini, Mercury, Apollo CM, OASIS CTV and Soyuz), while holding the habitable volume constant, the various architectures were compared.
Based on the mass
breakdown described in the OASIS Report, each component needed to be considered
independently. Based on the type of
component, its scaling relationship was chosen accordingly.
The following
components of the mass breakdown were assumed to be independent of the number
of crew:
The following
components of the mass breakdown were assumed to scale in direct proportion
with the number of crew:
The following
components of the mass breakdown were assumed to scale in direct proportion
with the external vehicle surface area:
The following components of the mass breakdown were assumed to scale in direct proportion with the habitable volume:
A more detailed discussion of scaling is provided for specific components of the mass breakdown.
The mass of the Power Systems was derived from the HPM, which consisted of,
Therefore, a total average power level of 3.580 kW was predicted. When scaling these values it was assumed that the cryogenic cooling system (the largest power load for the propellant management system - PMS) was independent of the number of crew. However, the power load will be significant when in-space propellant requirements are specified later in the project. The Guidance, Navigation & Control (GN&C) was assumed to be independent of the number of crew. This was also assumed for Communications & Tracking System (C&T) and the Electrical Power System (EPS).
As the number of crew increases, the pressurized vessel volume increases proportionally (according to NASA-STD-3000, Man-Systems Integration Standards). Based on conduction heat transfer, the increased surface area permits increased heat loss to the exterior of the vehicle. Since the heater power load is ~5% of the total heat load, the increased mass associated with a larger heating unit would only slightly increase the total mass of the vehicle (on the order of less than 1%).
A radiation protection layer is added to the vehicle exterior and as such, its mass is directly proportional to the exterior surface area of the vehicle. The exterior surface area was calculated for the base case (4 crew XTV).
According to Figure 75, the exterior was modeled as the following,
Since many of the mass contributions are a function of the vehicle surface area it was critical to calculate the approximate exterior surface area. Since docking hatches and other exterior attachments are located on either end of the vehicle it was assumed that only the main cylindrical volume and the lower conical volume contributed to the exterior surface area. Therefore,
.
The mass of the crew of four XTV, and the mass of the radiation protection material was scaled per external surface area. Based on the new vehicle geometries required for a larger number of crew the new radiation protection material mass could be predicted.
Figure 75: XTV scaling model
A surface area relationship was also assumed for the
pressure vessel. However, for the
pressure vessel, the surface area of the spherical vessel was used and scaled
based on the habitable volume of the vehicle.
The materials proposed for the Hybrid Propellant Module (HPM) were described by OASIS (2002) and are summarized below. The materials for this module are similar to the materials proposed for the Crew Transfer Vehicle (CTV). The HPM structural system meets the requirements of NASA Standard 5001, Structural Design and Test Factors of Safety for Space flight Hardware. This module can withstand the launch loads from a Shuttle-class RLV or an augmented Delta IV-Heavy ELV.
To protect this module and its contents from impacts due to micrometeoroids and orbital debris a Micrometeoroid and Orbital Debris Protection (MMOD) exterior shield was proposed. The designed shield is capable of withstanding an impact with no penetration from a 4mm diameter aluminum projectile with an impact velocity of 7km/s (ISS orbital velocity).
The longerons (axial members) are made from a magnesium metal matrix with long fiber carbon strands. This composite has a structural I beam cross section (S20cm-15cm), which facilitates the attachment of the MMOD. Both the upper and lower sections contain these types of members. For the upper section, the skin surrounding the shield is made of five layers of Kevlar fabric and epoxy composite. This serves as a stiffener to the main longeron and ring base structure. Space is left for layers of radiation protection of insulation for thermal protection. This section includes three layers of Nextel ceramic cloth that provides thermal protection (see Figure 76). The 30cm exterior thickness contains alternating layers of low density, open cell foam. This foam is a carbon-based graphite material with excellent thermal properties.
(Courtesy of OASIS, 2002)
Figure 76: HPM upper section material
For the HPM lower section, the exterior is slightly different. A Whipple type shield was chosen for the lower section MMOD shielding and thermal protection (see Figure 77). The shield is made of syntactic aluminum metal foam to minimize the material density, while maintaining sufficient strength.
(Courtesy of OASIS, 2002)
Figure 77: HPM lower section material
Tapered longerons were used for the internal support structure of the CEV. This was done to ensure the 4-G load (HPM thrusting) could be sustained. The upper half was untapered to maximize volume in the unpressurized area. It was also expected that the maximum loading during docking would be light.
The CTV
MMOD shield design is conceptually similar to the HPM MMOD shield. This section incorporates an expandable
multi-shock design, which is deployed on-orbit.
Non-expandable syntactic aluminum foam is used on the upper section to
avoid potential complications with shield deployment around externally mounted
systems (including solar arrays and radiators).
The number of longerons supporting the XTV structure was 8 (crew of 4). Therefore, it was assumed that the number of members was directly proportional to the exterior surface area. To a first order approximation, buckling scales well with surface area, provided additional ring supports could be added where necessary. This provides only a minor mass penalty and since the structural load per circumferential length will remain the same, this is a good assumption. It was assumed that the ring structure attached to each member also scales with the surface area. This is a reasonable assumption, as the support members constitute a greater portion of the total structural mass (assuming the same material is used for vertical members and ring base structure).
This exterior protective structure was assumed to scale with the exterior vehicle surface area in a similar manner as the Radiation Protection layer.
Similar to Radiation Protection and the MMOD, the Secondary Structure was assumed to scale with the exterior vehicle surface area.
Assumed to scale with the exterior vehicle surface area.
At this stage of the analysis it was assumed that the Thermal Control mass scaled with the volume of the pressurized vessel. This seems reasonable as the pressurized volume scales with the number of crew (NASA-STD-3000 90).
Each of the components was scaled according to the parameters specified above. The original vehicle proportions were maintained when the size was increased. Details of the vehicle proportions can be found in the Section 9.1.1.
A similar analysis was performed for the Apollo Command Module (CM) as was performed for the OASIS CEV.
The command module was approximated by a conical structure. The pressurized inner shell was fabricated from aluminum honeycomb panels and separated from the heat-resistant outer shell by a micro-quartz fiber insulator.
The external surface area (see Figure 78) was calculated as,
.
Figure 78: Apollo CM schematic
The mass breakdown of the Apollo CM is shown in Table 22.
Table 22: Apollo CM mass breakdown (http://www.astronautix.com/craft/apolocsm.htm)
·
A5G:
Ariane 5 “Generique”. It is the reference European commercial launcher.
·
ET:
External Tank. It is the external tank where de LH2 and LO2 are stored in the
shuttle.
·
Isp:
Specific impulse. It is a measure of the performance of an engine- propellant
combination. For a given system, it increases with altitude and is maximum at
vacuum. It is measured in seconds. If you have a kg of propellant and you burn
it to produce 1 kg of thrust, the Isp is the number of seconds it lasts.
·
SRB:
Solid Rocket Booster. Has a lot of thrust but low Isp, good for
lifting off.
·
J2:
high energy LH2 LO2 upper stage from the Saturn 5 third stage.
·
SL: Sea
level
·
P/L:
Payload
·
LEO:
Low Earth Orbit, an orbit between 150 to 1000 km over the Earth surface. Unless
otherwise stated, it is assumed to be of 280 km.
·
·
SSME:
Space Shuttle Main Engine, it uses liquid H2 and liquid O2. It has a high Isp
but it is expensive and complicated.
·
RS68:
Main engine of the Delta IV common core, it uses liquid H2 and liquid O2, it
has a lower Isp than SSME and it is not human-rated, but it is
cheaper and simpler.
·
STS:
Space Transportation System. It is the whole system commonly known as shuttle.
·
Pod:
Canister where the payload is stored in an STS-derived heavy launch vehicle.
·
Isp
of the SRB is constant and average between SL and vacuum.
·
Pod
weights 10000kg.
·
Isp
of SSME and RS68 is equal to the SL value while the SRBs burn and equal to the
vacuum value afterwards.
·
Massflow
is constant for each engine.
·
SSME
can be throttled to 109% of the nominal thrust at lift off.
·
The Isp
during parallel burn is the weighted average using the massflow rate as the
weight.
·
Inline
and piggyback configurations of STS-derived heavy launch vehicles are
considered equivalent.
·
The
gravity, turning, and aerodynamic losses are the same for all STS-derived
vehicles.
·
The
payload mass to LEO assumes a total delta-v (including all losses) of 9,400
m/s, the mass to ISS assumes total delta-v of 9,600 m/s, and the mass to escape
assumes delta-v of 11,600 m/s
No high-energy
second stage would be developed or revived.
· Parallel burn of the SSMEs (or RS68s) and SRBs for 123 seconds
· SRB separation.
· SSMEs (or RS68) continue to burn until they ran out of propellant.
· SSMEs (or RS68) and ET separate
· Payload separates from the Pod.
Figure 79: Shuttle-C elements (Source:
NASA)
The curves show total mission delta-V versus P/L mass for various
launchers.
Figure 80: Performance curves
A high-energy
second stage of the J2-class, similar to the third stage of the Saturn five
would be developed or revived.
·
Parallel burn of the SSMEs (or RS68s) and
SRBs for 123 seconds
·
SRB
separation.
·
SSMEs
(or RS68) continue to burn until they ran out of propellant.
·
SSMEs
(or RS68) and ET separate
·
J-2
class stage burn.
·
Payload
separates from the Pod and J-2 class upper stage.
The curves show
total mission delta-V versus P/L mass for various launchers.
Figure 81: Performance curves
Table
23: Mass requirements in
LEO (ISU SSP Report 99’)
Project
- |
Payload Mass Demand (tons) |
90 day study -
Moon mission |
60-100 tons (A.
Cohen, 1989) |
90 day study -
Mars mission |
140 tons |
Synthesis group
on |
150 - 250 tons
(T.P. Stafford, 1991) |
Alternative
infrastructure study from General Dynamics |
98 - 150 tons
(General Dynamics, 1993) |
NASA Mars
reference human mission. Version 1.0 |
240 tons (S.
Hoffman, 1997) |
NASA Mars
reference human mission. Version 3.0 |
80 tons |
Table 24 summarizes the calculations made on the
various STS derived options.
Table 24: Various STS-derived
options
|
Launchers with a newly developed high
energy J2-class upper stage |
Launchers with existing components |
|||
|
3SSME + 2SRB+ J2 |
3RS68+2SRB+J2 |
3SSME+ 2SRB |
3SSME + 3SRB |
3RS68 + 2SRB |
LEO (407km) |
110694 |
100424 |
86124 |
100813 |
49118 |
P/L to ISS |
103915 |
79151 |
80176 |
93299 |
39225 |
P/L to Escape |
50425 |
30694 |
Unpractical range |
||
TRL |
5 |
4 |
6 |
4 |
5 |
The most attractive
combination for heavy cargo launch is 3RS68, 2SRB and a J2 class upper stage.
·
It has
1 engine out capability.
·
It uses
RS68 which are not human rated and therefore cheap.
·
It
matches nicely with the NASA Mars reference human mission v. 3.0
The Evolved
Expendable Launch Vehicle program (EELV) is the backbone of the
In the framework of
the new human exploration initiative there are two potential uses of EELV
launchers, one is cargo and the other one is human transportation.
Concerning cargo
there are two approaches that can be based solely on EELV technology. One is to
use the Delta IV Heavy to launch the heavy exploration payloads in pieces of
about 20000 kg. This approach has been studied in Section 6.4.2.2.2. The other
approach is to develop a new heavy launcher based on Delta IV technology. One
such concept could be a Delta-IV with five common booster cores (instead of the
three used on the Delta IV Heavy) and a J2 class upper stage. We will refer to
it as Delta –IV Super. The Delta IV is assembled horizontally, that would not
be possible for the Delta-IV Super, and substantial changes in the way that the
launcher is assembled would be required. According to our calculations this
option promises payload capabilities that are slightly inferior to those with
an STS based architecture. A Delta-IV super would require new infrastructure
and more development work than an STS based launcher to arrive at a lower
performance. From our simplified analysis it seems that it is a less mature and
technically inferior option than to use the current EELV fleet or an STS
derived.
Concerning human
transportation it should be first noted that the EELVs were not conceived to
send humans, but classified payloads and commercial telecommunications
satellites instead. The fact that the vehicles are not man rated does not mean
that it cannot be done very effectively; as an example: the most family of
vehicles that has ever been used to launch humans, the Soyuz (originally the R7
missile), was conceived to launch thermonuclear warheads.
The launch pads
would have to be modified to allow for the access and escape of astronauts.
This would however be less costly than designing new launch pads from scratch.
Table 25 shows some characteristics that have been
calculated for EELV vehicles.
Table 25: Various STS-derived
options
|
Delta-IV Medium |
Atlas-V (552) |
Delta-IV Heavy |
Delta-IV Super |
G.L.O.W. (no payload) |
241,160 |
533,749 |
677,220 |
1,209,493 |
Payload to LEO (407km) |
10,081 |
11,446 |
18,531 |
51,599 |
Payload to ISS |
9,308 |
10,392 |
17,000 |
47,724 |
Payload to Escape |
3,873 |
3,566 |
7,162 |
20,508 |
Tech. Readiness Level |
9 |
9 |
8 |
6 |
Reliability |
94% |
87% |
94% |
90% |
•
All
masses in kilograms
•
Payload
mass calculated assumes no fairing and no escape system
•
Reliability
is estimated from current flight history and the flight history of the vehicle
generation immediately prior to EELV (e.g. Delta-II and Atlas-II Centaur)
It has been argued that a single Solid Rocket Booster (SRB)
is a viable alternative to launch humans. In this section we will evaluate the
attractiveness of using SRBs to launch crews.
The concept has some advantages. First, the vehicle
components are already human rated since they are used on the STS. Second there has been only one SRB failure of
in 113 flights. That is a 99.5% reliability, which is similar to that of the
Soyuz. Another advantage is that it supports the business of ATK Thiokol, a
critical supplier of the strategic nuclear forces of the
It should be noted though that the reliability of an SRB as a part of an STS stack cannot be simply assumed to be preserved in an SRB based launch system.
This concept would require new launch pads and ground infrastructure,
A concern that is often raised up is the environmental effect of the SRB plumes. Due to the low flight rate (in any case less than 12 a year) that we expect for the program, we consider that effect to be minor.
A more serious concern is the safety of the use and storage of stages that are constantly fully loaded with explosive. An explosion of an SRB in the VAB would be very damaging and it is a scenario that, although unlikely would be catastrophic. Naturally this applies to the STS derived launcher because it uses SRBs too.
A study of the capabilities of various combinations of the SRB with different high energy cryogenic upper stages has been performed and is summarized in Table 26:
Table 26: Various combinations
All masses shown in the table are in kilograms. The payload mass has been calculated assuming no fairing and no escape system. It should be pointed out that SRB with S-II combination involves a rather unusual geometry and mass distribution, the diameter of the second stage being 2.75 (10.2/3.71) times than that of the SRB. The aerodynamic, structural and control problems that such a configuration would have are at first sight very large however its assessment is beyond the scope of this report.
An important concern using this vehicle is the peak acceleration. Since the SRB has a very high thrust to weight ratio a very heavy mass would have to be launched every time that an SRB based vehicle is used to launch humans. To keep the STS requirement of less than 3 g is not necessary since, in the context of the new exploration missions, it will not be claimed that almost anybody can use the system, as was the case on the shuttle. 5 g is a more reasonable threshold. The minimum payload in LEO compatible with an acceleration of 6g has been calculated to be 47000 kg for a 5 g limit, 74500 kg for a 4 g limit and 116,000 kg for a 3g limit. Comparing these minimum masses to the payload masses with the various upper stages show that the only option that would have an acceleration of less than 5 g would be the SRB plus S-II stage. As has already being commented that option has problems in its geometrical configuration. This problem would not be easily dealt with just by designing a new upper stage with a different geometry. Actually, the S-II is quite a slender body. The physical reason can be traced back to the different in density of the propellants used in the SRB and the upper stages. Due to the very low density of liquid hydrogen, cryogenic stages occupy a very large volume. If the diameter of an upper stage where reduced by a 30% the diameter of the SRB then, to preserve the volume, the length of the stage would have to be increased by a 57% making then the second stage longer than the SRB.
This
consideration renders the use of a single SRB to launch astronauts troublesome.
The effect on the
final P/L mass of 1kg of inert mass placed on a stage has been evaluated for a
typical LEO mission for both an A5G and an STS-derived that used SSME and a
J2-class upper stage.
Figure 82: Ariane V and STS-Derived
For missions
requiring on orbit assembly, such as large Moon missions or any Mars missions,
it may be useful to have an assembly platform with a robotic arm. Using STS
legacy hardware, this capability could be achieved rather easily. To be able to
launch such a platform the STS Based has to be side mounted. It should be
launched on a low inclination LEO. External Tank Corporation studied the
development of a space station using the external tank as living quarters and a
modified orbiter such as the one that would be needed for this assembly
platform. The cost was estimated to be $3 billion FY92.
Figure 83: STS derived assembly platform
The tool described
in this Appendix was developed in LabView, and provided a way to evaluate
launch capabilities that could be considered for crew or cargo launch. Two
types of analysis can be done: a combinatorial evaluation of the possible
architectures and an evaluation of a single architecture selected by the user.
Figure 84 shows a view of the graphical user
interface. The following paragraph describes how to use this simulation tool,
and the results it generates for the evaluation of a single user-defined
architecture.
-
Step 1:
The user enters the parameters available for each option:
o masses, TRL and rank for the EDL technology,
the habitable module and the launch escape system
mass capability of lift for the first stages (4th and 5th rows of
o Figure
84)
-
Step 2:
The user selects one of each option (knowing that all the combinations between
rows are not always possible) with the help of the rectangular select button on
the right.
Step 3: The user runs the simulation to obtain the results, which are displayed in the bottom - right corner of
-Figure
84.
The program outputs
include,
·
The average
TRL for the selected architecture as well as the lowest TRL among the selected
options of this architecture;
·
The
mass margins for launch capabilities for ISS and 28.8 deg destinations, as well
as the mass margin for the escape system. The mass margin is defined as the
extra mass that the launcher could launch in addition of the crew module. Note
that in this part of the study, no care was taken of separating human-rated and
cargo launches;
·
The
rank of the crew module architecture;
·
The
reliability of the launch system in percentage.
Figure 84: GUI interface for the LabView combination
tool
The advantage of using LabView was also in that it enabled evaluation of all the combinations for each option of CEV (+ EDL + crew escape system) and launcher.
Figure 85 shows the result of such a simulation.
LabView evaluated, for 999 different architectures, the mass margin to ISS in
kg. The mass margin is defined as the extra mass that the launcher could launch
in addition of the crew module. For some architecture, the mass margin is
negative, which means that the selected launcher couldn’t lift the selected
crew module. On this graph, too, some gaps are presented (red arrows) and the
corresponding launch technologies, which enabled launching the mass
(3RS68+2SRB, Delta IV super, 3SSME+2SRB).
For the final
decision of a launch system, not only the masses (given by this model), but
also cost and feasibility issues should be looked at, these are not included in
this model.
Figure 85: Mass margin to ISS for 999 options of
launch + CEV configurations
Shown in Table 27 are the detailed requirements for the Mars and Moon missions. These requirements are discussed in greater detail in their respective Report chapters.
Table 27: Form/Function matrix
A long duration mission, such as one to Mars and back poses many new challenges that have not been the focus of earlier human exploration initiatives like Apollo. The size and composition of the crew is an extremely important factor based on psychological and sociological aspects of such a mission. Important factors to consider are summarized from an earlier MIT study in 16.851 Satellite Engineering (2003). Large crews tend to have lower levels of deviance and conflict and this tends to decline with increasing mission duration. Also, heterogeneous crews have lower rates of deviance and conflict (Dudley-Rowley, 2002). This same investigation indicated that the least dysfunction of any crew studies was a crew of nine people. Since a short duration mission will have the crew anticipating their return in the short term, long-term group dynamics are less of an issue.
Gender, ethnic and cultural make-up is an important factor for long mission durations. More heterogeneous crews begin a mission with some deviance, conflict and dysfunction, but this tends to decrease as the mission progresses. The opposite is observed for a homogeneous crew, whereby deviance, conflict and dysfunction tend to be initially less than a heterogeneous crew, but increase as the mission progresses (Dudley-Rowley, 2002).
As was discussed in an earlier section, the interior “free” space for a crew is important and should be sufficiently large for long-term mission durations. This results in increased performance and mental health.
Equipment necessary to keep the crew alive must be extremely reliable and robust to external perturbation. This equipment involves a galley & food system, waste collection system, personal hygiene, clothing, recreational equipment, housekeeping, operational supplies, maintenance, sleep provisions, and health care.
The pressurized spacecraft must have the appropriate temperature, pressure, and atmosphere, as well as control over disturbances from living organisms. According to Larson (1999), open loop life support systems (carry food, water and oxygen on board) are reliable, but are limited by mission duration (cost and volume occupation). An open loop system processes waste products and recovers useful resources. However, the disadvantage are; development costs, power requirements, reliability and maintainability (Larson, 1999).
The functions of a life support system involve managing atmosphere (pressure, temperature, humidity, removal of contaminants, composition and ventilation), water (provide, monitor, process and store for hygiene and drinking), waste (collect, process, store) and food (provide, store and prepare) (Larson, 1999).
Deciding which technology to select depends largely on the characteristics of the crew size, mission duration and mission location. For long-duration missions, hygiene water will likely dominate design decisions on sizing the ECLSS (Larson, 1999). An atmosphere management system suggested by HSMAD consists of 4BMS (4-bed molecular sieve), TCCS, and Sabatier P/C atmosphere management system (see Table 28).
Table 28: ECLSS atmosphere management
A flow chart of the processes used to manage the atmosphere is shown in Figure 86.
Figure 86: Atmospheric control and supply (Wieland, 1999)
A triple redundant system was selected, which consisted of the three systems listed earlier. The resultant mass was multiplied by a factor of three for redundancy (Wieland, 1999). This results in an overall mass and volume per crewmember (CM) of 255 kg/CM and 1 m3, respectively.
A P/C water management system of vapor compression distillation (VCD) was presented in the Mars mission design example in HSMAD. A flow chart of the water recovery and management is shown in Figure 87.
Figure 87: Water recovery and management (Wieland, 1999)
Two water management systems were selected, bringing the total mass and volume to 50 kg/CM and a volume of 0.2m3/CM (Wieland, 1999). The ECLSS atmosphere and water management systems were predicted for various crew sizes. The method of scaling was similar to the method described earlier in the Report, in which the mass of the ECLSS was scaled as being directly proportional to the number of crew (see Figure 88).
Figure 88: Mass and volume of ECLSS atmosphere and water management systems
For missions more distant than geosynchronous Earth orbit, the Earth's magnetic field does not provide protection and radiation from the Sun, especially during solar storms (during a Mars mission), and galactic cosmic radiation (Wieland, 1999). Since a crew will leave the Earth’s atmosphere on a mission from Earth to Mars, radiation protection from high-energy sun particles is required. Background space radiation, such as galactic cosmic rays (GCR) may also influence crew health during the mission to Mars.
The radiation dose limit is used to predict the hull thickness. Typically, spacecraft radiation wall thickness is determined by the wall thickness that does not permit the radiation dose limit. As such, a radiation dose of 1 Gy requires an aluminum hull thickness of 1.5cm (Larson, 1999).
Distinctions were presented between the design of the crew habitat for the interplanetary space travel and ones for the Mars surface (Cohen, 1996). This is known as the “Being There Versus Getting There” philosophy that argues that interplanetary and surface capabilities are fundamentally so different that it is not possible to optimize them within a single set of habitation elements. Cohen indicates that “radiation shielding is the most overlooked feature of proposed interplanetary vehicles” and that “NASA and space industry mission planners consistently underestimate the radiation hazards on a trip to Mars, particularly from GCRs and thus minimize the shielding to protect against this exposure.” Cohen indicates that the shielding requirements form radiation hazards in interplanetary space indicates the need for substantial omni directional shielding on the order of 30g/cm2.
It may be possible to extract shielding for the Mars surface habitat from the Martian surface, which indicates that the mass penalty of carrying sufficient shielding for the Martian surface habitat from Earth is unnecessary (Cohen, 1996). By having a small “solar storm shelter” in the crew transport vehicle, the overall radiation protection mass would be reduced, however this comes at the expense of close crew quarters, which as discussed earlier, affects group dynamics. From multiple standpoints, providing radiation protection for the entire crew vehicle is critical to the success of the mission.
Since a sphere has a minimum ratio of surface area to volume for any solid, as closely approximating this shape reduces the overall radiation shielding mass. According to Cohen, the shielding can either be solid, as in the form of solid aluminum gore panel or liquid, as in water to pump into interior perimeter tanks. Cohen also indicates that “whatever the shielding, it makes no sense to waste the effort, cost and energy that put it in LEO by landing it on Mars as part of a multi-purpose habitat.”
In ISS orbit, roughly half the radiation dosage comes from trapped protons and half from Galactic Cosmic Rays (GCR). As well, the flux of low-energy GCRs is inversely proportional to solar activity and
Solar Particle Events (SPEs) are mostly low-energy protons (30-120 MeV) and are more common at solar maximum. Since SPEs vary so much in size, it’s hard to design a spacecraft that will be totally immune to their effects (Larson, 1999).
This information was summarized from Larson (1999).
A spacecraft needs continuous housekeeping power to operate and support guidance, navigation, and attitude control, telemetry downlink, active thermal control, the computer system, and crew needs – circulating atmosphere, lighting, cabin heat, etc. Consider the following housekeeping power requirements:
· Apollo Command Module – less than 2 kW
· Skylab space station – about 4 kW
· International Space Station – about 36 kW
· For a Mars mission, the baseline (or continuous) power needed is estimated 30 kW-50kW for transfer vehicles.
The power system must withstand thermal cycling because of the short 35-minute orbits. Solar power systems should be able to store energy and account for radiation in low-Earth orbits.
PV arrays consist of solar cells with transparent covers to protect the cells from radiation. Arrays may mount directly on the spacecraft’s body without deployment or pointing. This option limits the array area and may compromise thermal control. The power produced from a given area of solar array from the power-conversion efficiency is:
where, ΦSun = [1368/d2] is the flux of sunlight at a distance d (in astronomical units) from the Sun, η the conversion efficiency that is dependent on the solar cell type chosen (for silicon cells used on International Space Station about η=.145), Fp the fraction of the array actually covered by solar cells (85-90%), and θ the angle of the array normal to the Sun.
The solar array blanket
mass is the mass of the solar cells, including cover glass and
interconnects, and the substrate on which the cells are mounted but does not
include the structural mass, array orientation motors, or the deployment and
packaging mass. Because of the
difference between the “standard” efficiency and actual delivered to the user,
the load power output is about 33% of the array’s output in nominal sunlit
conditions. Typical blanket mass per
unit area is 1.7 kg/m2 for silicon.
The blanket mass is typically
about 55% of the array’s total mass marray. The drive mechanism’s estimated mass mdrive is a function of the
array’s mass marray (in
kilograms) from
mdrive =
|-.014marray + 20.6|*marray/100.
The launch volume depends on
the technology chosen and the packaging.
The rough estimate for the total packaged volume for a solar array is 0.05m3 per m2 of the
array’s total area.
Solar thermal systems (known as solar dynamic, or SD) are predicted to have lower specific mass and cost for high power (>100 kWe) applications. Nuclear systems may be preferable for long eclipse, high-power missions such as large lunar base; and missions requiring low, long-term continuous power such as missions with low sunlight intensity. All thermal-conversion (SD and nuclear) require radiators to reject the waste heat, which account for much of the system’s mass and area.
The baseline SP-100 reactor can produce 100 kWe of electrical power with a specific power of 30 We/kg and a 7 year life-span. The shielding requirements to protect the crew and payload from the radiation produced by a nuclear power system can be a large fraction of the power system’s mass. A reactor system is typically not radioactive until energized. The reactor should be launched from Earth un-powered to allow handling on the pad.
One primary option is the hydrogen-oxygen fuel cell, which reacts hydrogen with oxygen to produce electricity and water. Fuel cells produce high energy per unit reaction mass, but are less compact and more complex (requiring storage of hydrogen and oxygen) than other batteries.
The cell’s mass is proportional to the power level required, and the mass of the reactants is directly proportional to the power required times the mission length. For example, the Shuttle’s system has three primary fuel cells that provide a total of 14 kWe. Each fuel cell has a mass of 91 kg. Together, the three cells use about 150 kg of hydrogen and oxygen per day.
This information was summarized from Larson (1999).
The thermal control system must maintain comfortable and a uniform temperature distribution for the crew, other systems and equipment. As well, the thermal protection system provides the first line of defense by shielding from extreme heat sources and sinks (sun and deep space, respectively)
Active thermal control implies movement of mass, information or energy. This is usually some type of loop to circulate fluids. This allows convective heat transfer to augment conduction and radiation. Passive thermal control involves conduction through the spacecraft and radiation from its surface to dissipate heat and keep temperatures relatively low. Geometric design and layout, insulation, heaters, heat pipes, and louvers are common passive techniques that aid in maintaining all parts of the spacecraft at acceptable temperatures.
Depending on the mission, the mission phases each contribute unique thermal requirements. After considering a particular mission, examine the phases (Table 16-2, Larson, 1999) and complete the following design iteration (Table 16-1, Larson, 1999).
Various mission phases are described in HSMAD (Table 16-2, HSMAD).
Typical heat rejection values of space radiators are given in Table 16-4 (Larson, 1999). Note that some of the rejection systems are not suitable for long-term missions because significant degradation will take place over time.
Thermal control fundamentals include convection, conduction and radiation. These three heat transfer processes are used in some form to control the temperature of the spacecraft interior. Table 16-7 (Larson, 1999) discusses the masses, power and volume of the components of a thermal control system.
The Moon temperatures range from 100K to 400K because there is no moderating atmosphere. The radiators on the Moon can be as low as 3K (facing deep space) and as high as 325K (vertical at the equator at noon). The interplanetary thermal environment is generally cold, dominated by solar radiation and deep space. In this environment, simple radiators, facing away from the Sun, are very effective.
Radiant heat exchange dominates the Martian thermal environment, but a thin CO2 environment helps transfer heat through conduction and convection. High velocity winds stir up huge dust storms that block radiation. Surface temperature of Mars can range from 130K to 300K.
Radiative and absorptive systems are the two basic external types used to dissipate entry heating. Radiative typically dissipates 80-90% of the heating. Absorptive systems absorb the heat through heat sinks, ablation, or transpiration. Ablation is a self-regulating transfer of heat and mass in which material absorbs entry heating then degrades. Transpiration means injecting a fluid through the skin of the vehicle into the boundary layer to provide cooling.
Radiative thermal protection systems are limited to about 1370ºC. This is a passive system and usually does not involve mass loss or shape change. Absorptive systems absorb heat through a phase change, chemical change, temperature rise, or convective or transpiration cooling. These systems are typically more complex and weigh more than radiative systems. If the material burns up, we can use the system only once. However, absorptive systems can handle the high heating rates from velocities needed for planetary travel or missile entry.
To solve the surface temperature of the thermal protection concept for a range of entry velocities between 6 and 14 km/s,
. (1)
Tw = wall temperature (K),
qs = surface heat flux (W/m2)
σ = Stefan-Boltzmann constant = 5.67 × 10-8 W/m2K4, and
ε = surface emissivity at wavelength mix corresponding to temperature, Tw.
Re-entry heat gain can be obtained by assuming a re-entry surface temperature and an insulation conductance,
. (2)
q = re-entry heat gain (kW/m2),
kA = insulation conductance (W/m2K), and
ΔT = temperature difference between interior and exterior surfaces (K).
For short-duration missions, heat sinks can be used as an active method of thermal control. Since fluids like water or ammonia absorb a tremendous amount of heat during phase change, and the space environment is conducive too this (near vacuum), the approximate mass of expendable heat sink required for this is,
. (3)
Me = heat sink mass (kg),
Q = required heat rejection rate (kW)
D = mission duration (s), and
hfg = latent heat of vaporization (kJ/kg).
Table 16-3 (Larson, 1999) should be consulted for general guidelines for using expendable heat sinks instead of radiators, and Table 16-10 (Larson, 1999) for the latent heats of vaporization for commonly used fluids.
Thermal capacitors are short duration heat sinks. Using the same principle of evaporative cooling, elements can be “cold-soaked”, which when heated, will cool the element. This technique incurs only a very small mass penalty (only the weight of the fluid soaking the element), however a complex analysis is required to model this transient heat transfer.
This thermal control system gathers all of the heat loads from within the pressurized volume (cabin or module), including those collected directly from equipment through heat exchangers and cold plates.
The internal thermal control system must cool the cabin air heat exchanger to control humidity and temperature (see ECLSS Report). A coolant capable of removing the heat from the cabin must be selected. An important consideration is the crew’s safety should the coolant leak. Water is often selected, however it is critical that temperatures do not fall below freezing. It may also be necessary to insulate plumbing, cold plates, and other surfaces that operate below the air’s dew point temperature.
Additional considerations for human space flight can be found in Larson (1999).
ADCS stabilizes the vehicle and orients it in desired directions during the mission despite the disturbance torques acting on it. ADCS determines the vehicle’s attitude using sensors and controls it using actuators. ADCS is a major spacecraft system, and its requirements often drive the overall S/C design. Components are cumbersome, massive and power consuming. Table 29 shows a summary of the process to design ADCS system on a spacecraft. This table was inspired from Larson (1999).
Table 29: Design process of ADCS
STEP |
INPUTS |
OUTPUTS |
1. Define control modes and derive the
corresponding requirements |
Type of insertion for launch vehicle |
List of different control modes during
mission Requirements and constraints |
2. Select type of spacecraft control by
attitude-control mode |
Orbit, pointing direction Disturbance environment |
How to stabilize and control: -
three-axis -
spinning -
gravity gradient |
3. Quantify disturbance environment |
Spacecraft geometry Orbit Solar/magnetic models |
Values for forces from -
gravity gradient -
magnetic, aerodynamic and solar pressure -
internal disturbances Effects of powered flight on control (center
of mass, cm offsets, slosh) |
4. Select and size hardware for the ADCS |
Spacecraft geometry, pointing accuracy and
direction, orbit conditions, mission requirements, life time, slew rates |
-
Sensor suite: Earth, sun, inertial or other sensing devices -
Control actuators (reaction wheels, thrusters or magnetic torquers) -
Data-processing electronics |
5. Define algorithms for determination and
control |
|
|
Note: In Table 29, steps 1 and 3 can supposedly be found in the literature for each type of mission. But for each chunk to be designed, work has to be done on steps 2 and 4 (shown in gray). Especially the output of step 4 determines the hardware and corresponding masses that should be taken onboard the spacecraft.
Figure 89: Attitude control modes, from Larson (1999)
Figure 89 shows typical attitude control mode in which spacecraft have to be in. The method to control these different modes depends on 1/ the mode and 2/ the type and amount of disturbance.
Disturbance torques can be cyclic (which average to zero over an orbit) or secular (which do not average to zero over an orbit). These disturbances can be controlled passively, i.e. without moving parts (by taking advantage of vehicle’s inertia or favorable disturbances), or actively. For active control, the S/C senses the attitude motion and applies control torques to counter it.
The hardware to account for ADCS includes actuators, sensors and computers (+electronic wiring). Table 30 gives a description of control methods and the hardware associated. It was inspired by de Weck (2001) and Larson (1999).
Table 30: Description of actuators, inspired by de Weck (2001) and Larson (1999)
|
Type |
Pointing options |
Attitude maneuverability |
Lifetime limits |
Additional hardware required |
Passive attitude control |
Gravity gradient |
Earth’s local vertical only |
Very limited |
None |
Libration damper: eddy current, hysteresis rods No torquers |
Passive magnetic |
North/south only |
Very limited |
None |
? |
|
Pure spin stabilization |
Inertially fixed any direction Re-point with precession maneuvers |
High propellant usage to move stiff momentum vector |
Thruster propellant |
Nutation damper Torquers to control precession (spin axis
drift) magnetically or with jets |
|
Dual spin stabilization |
Limited only by articulation on de-spun platform |
Momentum vector same as above De-spun platform constrained by its own geometry |
Thruster propellant De-spin bearings |
|
|
Active attitude control |
Reaction wheels RW (most common actuator) |
No attitude constraint |
Rates limited by available momentum and low torques |
Propellant (if applies) Bearing life, motors |
External torque required for momentum dumping |
Control moment gyros |
No attitude constraints |
Rates limited by available momentum Double gimbal CMG has limited torques |
Propellant (if applies) Bearing life, motors |
|
|
Magnetic torquers (to de-saturate RW) |
Depends on the Earth magnetic field |
Harmful influence on star trackers |
|
|
|
Thrusters (to de-saturate RW) |
No attitude constraints |
|
Propellant |
|
Table 31
shows the masses of reaction control system for some crew vehicle, these were
obtained from mass breakdown from Braeunig
(2001) and other NASA document for the XTV.
Table 31: ADCS masses for some crew vehicles
|
Mercury |
Gemini |
Apollo |
XTV (avionics…) |
RCS mass |
40 |
133 |
400 |
200 |
Total mass |
1118 |
1982 |
5806 |
5760 |
Percentage |
3.6 % |
6.7 % |
6.9 % |
3.4 % |
For a LEO communication satellite, according to Springmann (2003), typically the ACDS mass is 7% of the dry mass of the satellite, as shown on
Table 32.
Table 32: ADCS mass of communications satellite, from Springmann (2003)
According to Gavin (2003),
Table 33 shows the masses for the LM configurations (in kg):
Table 33: Apollo lander ADCS
|
Ascent stage |
Descent stage |
Total |
Reaction control system |
110 |
0 |
110 |
Dry mass |
2154 |
2783 |
4937 |
Percentage of dry mass |
5.1 % |
0 % |
2.2 % |
Propellant for RCS |
275 |
0 |
275 |
Figure 90: Apollo lander mass breakdown, from Gavin (2003)
As a first approximation, the percentage of the ADCS of the total dry mass is 3 to 7 %.
Analytical solutions to the equations of motion provide estimates for preliminary mission and vehicle design for atmospheric entry. The peak aerodynamic loads and heat rate can be used to estimate crew’s acceleration exposure and the required thermal protection system. For example, consider the equations of motion for ballistic re-entry. The peak aerodynamic load (Gmax), ge’s, at v is 0.607(ve) is:
(3)
where ve is the atmospheric entry speed (km/s), γe is the flight-path angle (radians), e is 2.71828, ge is the gravitational constant (9.81 m/s2), and Hs is the planet atmosphere’s density scale height (km). This quantity can be determined from a table of atmospheric density (ρ) and altitude (H) by forming an exponential fit between the altitudes of interest: Hs = -(H2 – H1) / ln (ρ2/ ρ1).
The keystone to a conventional EDL system is its heat shield. The heat shield is usually built from a structure of aluminum honeycomb and CFRP skins covered with an ablative material that absorbs the heat of the entry and keeps the payload at an appropriate temperature. The ablator most used on the Shuttle External Tank, Mars Pathfinder, and the Mars Exploration Rovers programs was SLA561V (Allouis, 2003). This material, based on a mixture of cork wood, binder and tiny silica glass spheres, has a density, ρm, of 264 kg/m3 and an effective heat of ablation, Qm, of 5.41 × 107 J/kg.
The maximum heating rate for an entry profile is evaluated at the stagnation point from a Sutton-Grave correlation:
(4)
where rn is the nose radius (m), ve is the entry velocity (m/s), ρ∞ is the atmospheric density, and k is 2.84 × 10-5. From Equation 2, it is apparent that the bigger the nose radius, the lower the heat rates. The maximum heat rate for a lifting body is given by:
(5)
where β is the ballistic coefficient (m / CDS), kg/m2, where m is the mass of the vehicle, CD is the hypersonic drag coefficient, and (L / D) is the hypersonic lift-to-drag ratio. Notice that higher ballistic coefficient values result in higher heat and deceleration loads.
The total heat load, Q, of the mission is derived from the integration of the heat rates (Equation 3) over the heat peak during entry. The ablator on the heat shield must be thick enough to keep the back face of the heat shield below a threshold temperature even after the ablation process that takes place during entry. The radiative equilibrium temperature is given by qs = εσ(Ts)4 where ε is the black-body emissivity and σ is the Stefan-Boltzmann constant (5.67 × 10-8 W/m2K4). The minimum thickness of ablator can be estimated from the heat transfer formula at constant wall temperature:
(6)
where α the diffusivity of the ablator is defined by
the thermal conductivity (n), the
density (ρ), and the specific heat (Cp):
(7)
The thickness ablated during entry is given by:
(8)
where q is the heat per unit area on the heat shield, ρm is the density of the material, and Qm the effective heat of ablation of the material. The minimum thickness for the heat shield is therefore δ+Δδ, but a safety coefficient is of 1.5 is usually applied to δ.
The Earth’s thick atmosphere allows a spacecraft to perform an un-powered descent and landing after an aerodynamically controlled entry. The design of parachute and parafoil systems is typically an iterative process. Still, we can use the main design driver – the terminal velocity of the probe – to determine some estimates. Assuming the parachute has a circular cross-section, the parachute’s diameter for a given Lander mass (mL) can be calculated from:
Drag = ρ v2 / 2 × CD × A = mLg (9)
where ρ is the atmospheric density at parachute release, v is the desired terminal velocity, CD is the drag coefficient (parachute and vehicle), mL is the mass of Lander at engine ignition, g is the planet’s acceleration due to gravity. Both the Apollo and Russian Soyuz capsules used parachutes to get terminal velocities of about 9 m/s and 7 m/s, respectively.
The X-38 lifting body used a steerable parafoil that slowed vertical landing speeds (about 5 m/s), increased cross-range capability, and allowed for a gliding touchdown on land. Parafoils allow for improved hypersonic range and cross-range capability, as well as decreased entry acceleration and thermal loads without the complexity of horizontal landings. An Apollo-class vehicle, like the Modern Command Module (MCM), designed with this technology would no longer require a water landing and recovery. Furthermore, by using a flare of the parafoil system, we can reduce touchdown loads without the use of decelerating retrorockets. The following equation estimates a vehicle’s straight-line range while on a parafoil, assuming the flight path has a small radial acceleration and rate of change:
where R0 is the planet’s radius near the landing zone, (L/D) is the parafoil’s lift-to-drag ratio, and Hi and Hf are the vehicle’s initial and final altitude above the ground.
As suggested by the equations of motion, a spacecraft must be aerodynamic to perform EDL. The external shape is critical for interactions with the significant atmospheres of Earth and Mars. The lift-to-drag ratio (L/D) is a convenient way to express a vehicle’s ability to maneuver in the atmosphere. In addition, the vehicle’s mass distribution and location of the vehicle’s center-of-mass is fundamental to its controllability.
DV Table for Lunar Missions Using Lunar
Orbit |
|
|
|
|
DV (m/s) |
To
Moon Transit |
3100 |
To
Moon Orbit (100km) |
800 |
To
Moon Surface |
1870 |
To
Moon Orbit (100km) |
1870 |
To
Earth Transit |
800 |
Table 34: DV table for lunar missions using lunar orbit
DV Table for Lunar Missions Using EM-L1 |
|
|
|
|
DV (m/s) |
Transit
to EM-L1 |
3100 |
To
EM-L1 Orbit |
600 |
To
Moon Transit |
150 |
To
Moon Orbit (100km) |
600 |
To
Moon Surface |
1870 |
To
Moon Orbit (100km) |
1870 |
Transit
to EM-L1 |
600 |
To
EM-L1 Orbit |
150 |
Transit
to Earth |
600 |
Table 35: DV table for lunar missions using EM-L1
Payload Masses |
|
|
|
Module |
Dry Mass (kg) |
Crew
Operations Vehicle (COV) |
5700 |
Modern
Command Module (MCM) |
5200 |
Lunar
Lander (LL) |
10000 |
One
Octahedron of Habitation Module (HM) |
9167 |
Surface
Habitation Module (SHM) |
38150 |
Table 36: Lunar payload masses
Other Parameters |
|
|
|
Propulsion |
Isp (s) |
Cryogenic
Chemical Propulsion |
425 |
Electric
Propulsion |
3200 |
Structural
Factor |
0.1 |
Boil-off |
0 |
Table 37: Other lunar mission parameters
In order to compare different architectures and different trajectories, to select a baseline mission design, estimates of the initial mass in LEO were presented. In this section, a verification of the payload masses is given, followed by the relevant equations and a detailed calculation. All assumptions are restated for convenient reference.
A manned mission to
Mars will require a significant mass to be launched from Earth in order to
provide the required delta V (DV)
capacity and life support necessary for the transit and surface stay. The mass estimates derived in this report are
compared to those a paper by Walberg (1993).
Walberg’s paper
reviews four mission classes (opposition, conjunction, conjunction with a fast
transfer, and split-sprint mission) with three different scenarios (all
propulsive, aerobraking, and nuclear propulsion) and gives IMLEO estimates for
each. After describing each mission
class and scenario, I will compare the mass estimates for validation against
the reference values for each mission scenario.
The following is a short description of each mission class, as presented in Walberg (1993).
The first mission type
investigated is an opposition class mission, with a Venus swing-by, as
displayed in Figure
91. An opposition
class mission is one where the alignment for one leg of the transfer is not
optimal, but allows for relatively short planetary stays, between 30 and 60
days. Performing a swing-by at Venus
allows for a major reduction in the required DV for a
relatively small increase in time of flight.
Comparing the DV’s given for missions not employing a Venus swing-by
Larson (1999), to those using a swing-by Walberg (1993), and assuming that a
direct entry is performed at Earth for each, the average DV savings is on the order of 8.3 km/s, reducing the average mission DV from between 16-23 km/s to between 8-12 km/s. The total increase in flight time required to
perform a Venus swing-by is between 50 to 100 days (depending on a given
year). Other considerations, such as a
close pass the sun must be taken into account when selecting a mission design,
but for the purposes of analyzing initial mass, these considerations are
neglected.
Figure 91: Diagram of opposition class mission with a
Venus fly-by (NASA DRM website)
The second type of mission
is a conjunction class mission where a trade in surface time on Mars is made in
favor of a more optimum flight trajectory, as displayed in Figure 92. Thus, this
class of mission requires stays on Mars on the order of 300 to 500 days. However the DV required is
significantly reduced (between 5.2 km/s and 6.9 km/s), assuming a direct Earth
entry. In addition, the decrease of DV required to enter both an Earth orbit and a Mars orbit are reduced
which better facilitates aerobraking.
Figure 92:
Diagram of conjunction class mission (NASA DRM website)
The third type of mission
is a fast-transfer conjunction-class mission, as displayed in Figure 93. This type of
mission increases the required DV to between 8 and 10 km/s,
assuming direct entry at Earth, but decreases the transit times to recorded
zero-g levels (between 100 to 200 days per direction). The stay times at Mars are increased slightly
(approximately 50 days).
Figure 93: Diagram of fast-transfer conjunction class mission (NASA DRM website)
The fourth type of
mission is a split-sprint mission. The
basic idea is to pre-position the cargo, including return supplies and return
propellant using a fuel-efficient conjunction class mission. The crew is then transported via an
opposition-class (outbound) and fast-transfer (inbound) mission, such that the
overall mission duration is reduced to around 440 days, with a 30-day surface
stay. Although the DV’s are on the order presented above for
opposition-class and fast-transfer missions, the amount of payload on such a
mission is significantly reduced.
Table
38 :
|
Transfer TOF (days) |
Surface Stay (days) |
Total DV (km/s) |
Opposition w/ Venus Swing-by |
470-750 |
60 |
8-12 |
Conjunction |
400-700 |
300-500 |
5-7 |
Fast-transfer |
200-400 |
500-650 |
8-10 |
Split-sprint |
410 |
30 |
12-18 |
Having overviewed
the four missions classes described above, Walberg detailed the required IMLEO
for each mission using three scenarios.
The first scenario is an all-propulsive maneuver including propulsive
orbit insertion at both Earth and Mars.
The second scenario uses aerobraking at both Earth and Mars to reduce
propellant requirements. The third is
the use of nuclear propulsion. For
chemical propulsion, the specific impulse is 480 sec, and a structure to
propellant ratio of 0.1 is assumed. For
aerobraking, the structure mass is assumed to be 15% of the payload mass. For nuclear propulsion, a specific impulse of
960 sec. is assumed. In addition, a 5%
gravity loss is assumed for all propulsive maneuvers.
The general
trajectory as outlined by Walberg includes five maneuvers: a trans-Mars injection from a 500 km circular
Earth orbit, a mid-course correction, insertion into a 1-sol Mars orbit, a
trans-Earth injection, and an insertion into a 500 km circular Earth
orbit. Thus the mission architecture is
similar to that of Apollo, using a Mars orbit rendezvous. All propulsive maneuvers to the surface of
Mars and return to Mars orbit are included in the payload mass for the Lander. Walberg lists payload masses for the
habitation module, landing module, and Earth return capsule. The habitation module is the crew living
quarters during transit to and from Mars.
The landing module includes crew living quarters for the surface stay as
well as the propulsion for these maneuvers.
The Earth return capsule is similar to that of an Apollo style mission
and houses the crew during Earth orbit insertion and return to Earth.
The payload masses
defined by Walberg are taken from Freeman et al. (Freeman, 1990) and are
assumed to be in agreement with the NASA 90-Day study (NASA, 1989). However, the assumption of crew size is not
stated in either Walberg (1993) or Freeman (1990), and although, the NASA
90-Day study assumes a crew of four for a Mars mission, it does not outline any
solid numbers for payload masses. In
addition, Walberg lists two different masses for the habitation module. A larger mass is given for opposition and
conjunction class missions, due to the increased transit times. However, there is no difference in the Lander
mass despite different surface stay times and no verification of the reduced
mass estimates is provided.
The payload masses
used in this paper are derived from a number of sources, but loosely reference
equations and estimates provided in Larson (1999). These estimates assume a crew of 6, do not
account for mission duration and assume a surface habitat is provided for all
missions, such that the Lander mass is independent of surface stay
requirements. Table
39: Comparison of opposition class mass estimates with Walberg through Table
41 list the payload masses for each mission
class defined by Walberg, for presumably a crew of 4, and the current mass
estimates used for a crew of 6. The
Lander wet masses for the current estimates were calculated by applying the
rocket equation, using DV’s obtained from
Larson (1999), and the vehicle assumptions stated above.
Table 39: Comparison of opposition class mass estimates with Walberg
A |
Walberg mass (t) |
Current mass estimates
(t) |
Crew size |
4 |
6 |
Habitation |
61 |
60.7 |
Lander (dry mass) |
Unknown |
30* |
Lander (wet mass) |
76 |
53.5 |
Earth return vehicle |
7.8 |
9.2 |
* Assumes surface
life-support and habitation requirements are pre-positioned
Table 40: Comparison of conjunction class mass estimates with Walberg
Module |
Walberg mass (t) |
Current mass estimates
(t) |
Crew size |
4 |
6 |
Habitation |
61 |
60.7 |
Lander (dry mass) |
Unknown |
30* |
Lander (wet mass) |
76 |
53.5 |
Earth return vehicle |
7.8 |
9.2 |
Table 41: Comparison of fast-transfer mass estimates with Walberg
Module |
Walberg mass (t) |
Current mass estimates
(t) |
Crew size |
4 |
6 |
Habitation |
46 |
60.7 |
Lander (dry mass) |
Unknown |
30* |
Lander (wet mass) |
76 |
53.5 |
Earth return vehicle |
7.8 |
9.2 |
Using the payload
masses listed in Table 39 through Table
41, and DV’s
listed in Walberg, his results are verified with the rocket equation
calculations used. Under the same
mission assumptions as those provided in Walberg, we can compute the IMLEO for
the current estimates and these results are listed in Table
42.
If we compare the
IMLEO for each mission scenario, we notice a few trends. The mass estimates are higher than Walberg,
but these estimates take into account a crew of 6, which would seem to validate
the results. If we compare each mission
scenario with and without aerobraking, it becomes obvious that aerobraking
yields a significant benefit for each mission.
Nuclear propulsion yields a significant mass savings for each mission
class, which would indicate that this is an area of technology that may be
worth developing.
Table 42: Comparison of IMLEO estimates with Walberg
|
Walberg IMLEO (t) |
Current 16.89 mass IMLEO
(t) |
Opposition |
1268 |
1505 |
Opposition with Aerobraking |
593 |
806 |
Opposition with Nuclear Propulsion |
409 |
493 |
Conjunction |
597 |
715 |
Conjunction with Aerobraking |
500 |
599 |
Conjunction with Nuclear Propulsion |
285 |
345 |
Fast Transfer |
1440 |
1888 |
Fast Transfer with Aerobraking |
599 |
806 |
Fast Transfer with Nuclear Propulsion |
392 |
527 |
The payload
estimates provided by Walberg can be reproduced using the rocket equation
calculations defined, which validates the model used in this paper. The current mass estimates seem to provide a
reasonable approximation of the IMLEO estimates since they are only slightly
higher than those provided by Walberg, but are meant to accommodate a crew of 6
instead of a crew of 4.
Having analyzed the
different IMLEO estimates for comparison with each other and with the different
mass scenarios, it has become obvious that aerobraking is required to achieve
reasonable IMLEO estimates for chemical propulsion. In addition, nuclear propulsion has been
shown to yield an additional benefit, and may warrant further development.
The initial estimates of
the mass in LEO derived by applying the rocket equation in succession to the
payload masses. The rocket equation for
n stages is simply
where ai is the
structure factor. By applying this
formula from the final payload mass delivered back to Earth, we can determine
the initial mass in LEO for a chemical burn.
If aerobraking is employed, the formula for calculating the initial mass
for that stage is simplified to
where g is the aerobraking factor, which is set to 0.15 for this
analysis. If electric propulsion is used
to determine the initial mass for a pre-positioned element, the equation used
is
where SP is the specific
power, g is the efficiency, and t is the time of flight.
Table 43 shows an example calculation. In this example, we list the payload masses
and the DV’s necessary to perform each maneuver. The module names are the payload masses for
different maneuvers, the functions are different transfer maneuvers required,
and the specific impulse is assumed to be 425 sec. In this example, we assume that the return
propulsion, Landers with fuel, and habitation module is pre-positioned. The total mass in LEO is the sum of all the
components.
Table 43: Example calculation
The knowledge
delivery infrastructure will consist of two parts the delivery of data in the
form of bits and the delivery of samples from the planets surface. This section will largely deal with the
delivery knowledge in the form of bits and thus will in this section be
referred to as communication delivery system.
For every mission
size the same communication radio frequency has been selected in order to
provide an easily extensible system. The
radio frequency that each of these missions will use is Ka-Band or 32
gigahertz. This frequency was selected
because it can support a high data rate with comparably lower power than all
lower frequency bands, and because the DSN ground infrastructure will support
it by the year 2007, while other higher frequency bands are not supported by
the DSN. There is some concern about
weather interference especially when communicating with Mars, however a Martian
sand storm would prevent a X-band communication as it would a Ka-band communication,
the differences would mainly lie in the moderate weather such as a cloudy day,
or light dust storm in which case the Ka-bands data rate would be decreased.
For the small sized
lunar missions a direct link can be set up between the Moon Lander and one of
the Earth’s DSN stations. This would
allow constant communication between Earth and the Moon throughout the entire
mission. The data rate required for this
mission would be 1 gigabit/day would require 0.01 Watts of power per
transmission with a transmission data rate of 0.07 megabits/sec. After the mission is completed the
communication equipment that was landed on the Moon will be left there for two
reasons, one if a future mission decides to use that spot as a landing or
settlement site then they won’t have to bring their own equipment and in the
unlikely case that another future mission communication equipment fails the
crew will have the option of traveling in a rover to the old site and using its
equipment.
For the medium
sized missions the infrastructure is essentially the same as the small mission
except that it will require a higher daily data rate and transmitting power.
Daily data
rate: 10 gigabits/day
Transmission data
rate: 0.7 megabits/sec
Power
required: 0.1 Watts
For the Large sized missions require the ability to communicate between the far side of the planet and Earth. The astronauts will communicate through one of four possible ways. For the first option an comm. relay satellite could be placed at the L4 point in the Earth Moon system, this is good because it allows for a constant communication stream between the Earth and the Moon, unfortunately this option would only allow for communication for the first 900km onto the Moon’s far side. The next option is to set up a relay satellite in a Low Lunar orbit that has the advantage of covering most of the Moon, but the disadvantage of a large time delay between far side communications. The third option is probably the best and involves setting up a satellite in orbit around the Earth-Moon L2 (EM-L2) point thus covering the far side of the Moon in it's entirety and can keep in almost constant communication with Earth, the drawback to this architecture is that though orbiting the L2 point is technically and theoretically feasible it is untried and less stable than placing the satellite at a Lagrangian point. The last option is more for emergencies sake than anything else, it is possible for the dark side of the Moon to use an orbiter around Mars or a Martian settlement as a relay to Earth, the main problem with this is that it would require a large amount of power for a very small data rate, and would only be feasible at certain windows when Mars is visible to the far side of the Moon. The daily data rate for a large sized mission would be 50 gigabits/day, the transmission data rate would be 3.5 megabits/sec and the transmission power required would be 0.5 W.
All of these missions will have the capacity in some manner
to point their antennas at Mars and send or receive communications from or to
future Mars missions at low data rate.
For the Mars missions only a small a large size mission will
be considered. For the small sized Mars missions a direct link can be set up between
the Mars Lander and one the Earth’s DSN stations. This would allow semi-frequent communication
between Earth and Mars throughout the entire mission. The data rate required for this mission would
be 1 gigabits/day and would require 8 Watts of power per transmission with a
transmission data rate of .035 megabits/sec.
After the mission is completed the communication equipment that was
landed on Mars will be left there for two reasons, one if a future mission
decides to use that spot as a landing or settlement site then they won’t have
to bring their own equipment and in the unlikely case that another future
mission communication equipment fails the crew will have the option of
traveling in a rover to the old site and using its equipment.
The Large sized missions require the ability to communicate with much greater data rate and thus it might be necessary to create a relay satellite around Mars. There are two realistic options for the location of this sat. The satellite could be placed in a Geostationary Martian orbit around the landing site, the advantage of a GMO satellite is that it increases the time that the astronauts can communicate with the Earth, the disadvantage is that it can only really be set up for one portion of the planet. The other option is to position a satellite at the Earth-Mars L1 point thus decreasing the power required to send large communication streams to the Earth, unfortunately this would not add any extra time that the mission could communicate with the ground. As with the Moon missions there is an option in the case of emergencies to communicate with the Moon and use it as a relay station. The daily data rate for a large sized mission would be 10 gigabits/day, the transmission data rate would be 0.35 megabits/sec and the transmission power required would be 8 W.
Figure 94: Communication Architecture
Link Budget:
P =
E/N+10*LOG(k)+10*LOG(T)+10*LOG(R )-Ll-Lh-Lit-Lw-Lp-Gt-Ls-La-Lo-Gr
Where:
P = power
E/N = signal to noise ratio
Derived from the required bit error rate (BER) and type of coding.
k = Boltzman’s
constant
T = antenna noise
temperature
Provided by the ground stations
R = data rate
Estimated for different sized missions
Ll = transition station line loss
Estimated
Lh = hot body noise loss
Estimated
Lit = Ionospheric & Tropospheric loss
Estimated
Lw = weather losses
Estimated from DSN 810-005
Lp = polarization mismatch loss
Estimated
Ls = space loss = 10*log((λ/(4πS))2)
Where S is the distance between the transmitter and the receiver and λ is the wavelength
La = receiving station line loss
Estimated
Lo = other losses
Estimated
Gt = transmitting antenna gain = 10*log(η((π*D)/ λ)2)
Where D is the diameter of the antenna, λ is the wavelength, and η is the antenna efficiency.
Gr = receiving antenna gain = 10*log(η((π*D)/ λ)2)
Where D is the diameter of the antenna, λ is the wavelength, and η is the antenna efficiency.
The case for the use of optical communication in an extensible exploration program is not as strong as the current knowledge of Ka-band or X-band communication. Optical communication has several advantages over the more common radio communication channels, these being a much higher data rate for the same amount of mass, power and volume, all very important in the design of a space mission. Unfortunately there are also several devastating drawbacks to optical communication, in particular the serious losses due to atmospheric interferences and its high pointing requirements, making it next to impossible to use optical communication over any distance greater than from the Earth to the Moon. These drawbacks while serious can be overcome in some cases by the aforementioned advantages, however, in an extensible exploration program such as this one that is being proposed, it is far more important to have a common communication system for the entire program, thus when the need arises to extend to the next exploration site there is already a communication network in place to help relay transmissions. This is not to say that optical communication should not be used in all space missions, on the contrary there are many situations where the use of optical communication would benefit the mission substantially, however in this case it is wiser to use a more established form of radio communication.
One of the main interests of Mars is the search for water. Possible locations for water are the polar caps, subsurface ice, gullies, stream lined islands, rampart craters, outflow channels, and layered terrain (PSSS 2003). There are two scientific methods of determining water on Mars, geologically or by studying the climate and their objectives and approach are summarized below.
Some of main questions in Mars geology include understanding planetary origin and evolution by determining the core and mantle size and composition and mapping the current and past tectonic activity (PSSS 2003). Additional knowledge questions can also be found using Mars Field Geology, Biology, and Paleontology Workshop (Harvey 1998).
A great deal of research has been done on the instruments needed to gather scientific and resource related knowledge on the Moon and Mars. An important first step before sending robotic explorers is to understand what current information exists. An example of a database of Moon and Mars constituents and their locations can be seen in Table 1. Future robotic missions can add to the resolution of location and occurrence until it is necessary to send a human mission.
Table 44: Moon resources -
preliminary findings (
A method of determining science and resource related knowledge is through the use of a geophysical network. There are several methods of achieving this, outlined by the Lunar Exploration Science Working Group and the 2003 JPL Planetary Science Summer School (LExSWG 1995, PSSS 2003). Both studies included using penetrators and soft Landers. A summary of different deployment methods and their advantages and disadvantages are seen in Table 45. A challenge of creating geophysical nets is aligning the instrument and achieving a large global access for a long duration. Current planning tends to focus on penetrators and soft landings, which can be accomplished by robotic missions such as the Mars Net Landers. However human missions have two advantages over robotic geology missions. They are the experience knowledge gained by a human mission and a more optimal aligning and positioning of geologic instruments. The goals of the space transportation system are to eliminate the current disadvantages listed such as the high expense and global access.
Table 45: Methods of creating geophysical networks (LExSWG, 1995)
Past research on the capabilities needed to varying amount of knowledge returned can also be seen in “Geoscience” (1988). Outlined in Table 46 are three levels of knowledge, correlating with small, medium, and large, for a Moon mission and their respective instruments with focus on geology. Future Moon and Mars architectures should also have detailed instrumentation and the information gathered levels for climatology and resource related knowledge.
Table 46: Knowledge levels and instrumentation for a moon mission (Geoscience, 1988)
For this design study we do not need a thorough understanding of the geologic value of the Moon (or Mars), but we do need to know what landing sites will be sought by the scientists, so that we can design our missions accordingly. Unlike early Apollo missions (Figure 96), we should not be so constrained to the equator. Based in recommendations from some planetary geologists as well as landing site selection papers, we recommend the capability for orbits to at least +/- 30° latitude, and maintain the capability to land on the far side. This should satisfy most of the suggestions for landing sites.
The following is an elementary and somewhat oversimplified summary of some of the scientific motivations for going to the Moon, and associated landing sites.
Figure 97 and Figure 98 show the sites (sites [A] through [G], plus Apollo 11 site in Figure 96), and images of each site are also shown in the figures section.
Volcanism. When did volcanism end on the Moon? The Lichtenberg Basalts (68°W, 32°N, [A], suggested by Robinson, personal communication) are potentially the Moon’s youngest basalts; they may be between 1 and 2.5 billion years old. This area would give us a view of what basalts looked like from the beginning of the Moon’s volcanism (~3 billion years) to the formation of this area.
Lunar
volatiles. Aristarchus Plateau
(48°W, 24°N,
[B], suggested by Robinson and Taylor, personal communications) is a
complex pyroclastic center with a rich volcanic history. Samples are likely to be a diverse suite of
the magma source, and it would be possible to determine spectral
reflectance/composition from this previously unsampled material. Alphonsus (3°W,
13°S, [C])
and Sulpicius Gallus (20°N,
12°E, [D],
both suggested by Robinson, personal communication) are two other
regions with pyroclastic materials that would hint at the history of lunar
volatiles (which are necessary for volcanic explosions). The
KREEP basalts. Apollos 12, 14, and 15 found a strange basaltic material they nicknamed KREEP (potassium, Rare Earth Elements, and Phosphorus). Rare earth elements are extracted from liquid magma when other elements in the magma thermally differentiate and cool into a crystalline structure; the rare earth elements do not fit into this crystalline structure, so they form abnormally high concentrations, nicknamed KREEP. This finding introduced the concept that lunar maria did not form simultaneously, but over hundreds of millions of years. The Apennine Bench Formation (mountain range centered at 0°, 20°, [F], Robinson, personal communication) would allow us to sample this mysterious material and learn more about the thermal differentiation of the ancient magma and maria formation. Aristarchus Crater (48°W, 24°N, [B], Taylor and Schmitt, personal communications) is also likely to contain high concentrations of KREEP, and would also allow us study of cratering processes and crustal stratigraphy.
The
poles. The
Stratigraphy.
As mentioned above, Aristarchus Crater [B] and the South Pole
Aitken Basin [E] are useful for studying crustal processes. The crater Tsiolkosky (129°E,
21°S, [G],
suggested by Taylor and Schmitt, personal communications) has a central peak,
which may be a part of the original lunar crust. It may be a “great place to study the nature
of cumulate anorthosites” (
Seismology. The nature of the lunar interior is still somewhat ambiguous. Neal et al. (2003) suggest a lunar seismic network (see Figure 99), including a minimum of 8 seismometers deployed around the Moon. These could be deployed with unmanned or manned missions, but could certainly involve international cooperation. Seismometers would allow testing of the hypothesis that the Moon was formed from a magma ocean in its early stages.
Other questions (Ryder et al., 1989). The origin of the Moon may be better studied if early thermal differentiation were better understood (see KREEP discussion above). Lunar mare basalts should be better studied to understand not only the thermal history of the Moon, but also the depth of the ancient magma oceans, and their detailed compositions. Lunar stratigraphy may help us to understand the impact history of the Moon – when were the greatest periods of bombardment? (This has implications for the history of Earth, because if there was heavy bombardment on the Moon, there also was on Earth.) Lunar regolith (loose soil), which sits on the top layer of the Moon, contains the history of billions of years of solar wind and flares. Finally, usable lunar resources (such as water and Helium-3) need to be explored for future manufacturing plausibility. Landing sites for this purpose may include Mare Tranquilitatis (Apollo 11 site) and Aristarchus [B] (both suggested by Schmitt, personal communication).
Figure 96:
Apollo landing sites. Near side
of the Moon, center (0, 0). |
|
Figure 97: Near side of Moon. Landing sites are numbered
according to text of this report.
Center (0, 0). |
Figure 98: Far side of the Moon. Landing sites are numbered
according to the text of this report.
Center (0, 0). |
Site A.
Lichtenberg Basalts. |
Site B.
Aristarchus Plateau. |
|
Site B.
Aristarchus Crater. |
Site C.
Alphonsus. |
Site D.
Sulpicius Gallus. |
Site E.
|
Site F.
|
Site G.
Tsiolkoksy Crater (southwest edge shown; peak is on left side of
picture). |
Note: All images are from Schultz 1972,
except Site E, which is from Harland 1999.
Figure 99: Figure 1 from Neal et al. 2003. A lunar seismic network is proposed to study the Moon's interior.
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