CranSpace - Mars Societymembers.marssociety.org/inspiration-mars/semifinalists... · 2019-09-06 · The following report proposes a design for a two-person Mars ﬂy-by mission to

CranSpace

Idriss SISAÏD Edward ANASTASSACOS

Anees AHMAD Laia RAMIÓ

Kaitlyn MORRIS Enrique GARCIA BOURNE

Daniel PASTOR Mairéad BEVAN

Thomas AURIEL Jack LONGLEY

International Student Design Competition

INSPIRATION MARS

Cranspace.com

Inspiration Mars Student Design Competition Section 0

Acknowledgments

We would like to thank Dr. Jennifer Kingston (Cranfield University) for her support through-

out the project, particularly for the advice she gave us regarding cost estimation; Professor

Michael J. Rycroft (University of Cambridge) for his keen interest in the project and con-

tributions especially with regard to radiation shielding components. We also extend our

thanks to Dr. Donald Rapp (former research professor, University of Southern California)

for his professional insight and in-depth knowledge specifically regarding radiation related to

Mars travel. Furthermore, Dr. Don Boroson (MIT Lincoln Laboratory) for sharing his ex-

pertise on laser communication systems and Dr. Arturo Casadevall (Albert Einstein College

of Yeshiva University) for his contribution to our understanding of the shielding proper-

ties of Melanin Nano-Shells. Finally, we would like to thank Dr. Joseph Hamaker (Senior

Cost Analyst - SAIC / The Millennium Group International; former Director of Cost Anal-

ysis Division, NASA) for his invaluable contribution to the development of our cost analysis..

We would also like to acknowledge the contributions made to the project by Andrew McMil-

lan and Daphne De Jong.

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Contents

1 Report Overview 1

2 Inflatable Space Habitat 2

2.1 General Assumptions and Considerations of Sundancer . . . . . . . . . . . . 2

2.1.1 Specifications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

2.2 Advantages and Disadvantages . . . . . . . . . . . . . . . . . . . . . . . . . . 4

3 Radiation Shielding 5

3.1 Radiation and Water . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

3.2 Waste Management System . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

3.2.1 Waste Collection and Delivery . . . . . . . . . . . . . . . . . . . . . . 11

3.3 Application of Melanin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

3.4 SPE Events . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

3.5 Magnetic Field Probe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

3.6 SPE Shelter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

4 Internal Housing and ECLSS 14

4.1 Bubble Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

4.2 Food Management . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

4.2.1 Food Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

4.2.2 Food Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

4.3 Waste Management Configuration . . . . . . . . . . . . . . . . . . . . . . . . 17

5 Communications 18

5.1 Traditional Communications System . . . . . . . . . . . . . . . . . . . . . . 18

5.2 Remote Communications System . . . . . . . . . . . . . . . . . . . . . . . . 19

5.3 Communication Satellite Design . . . . . . . . . . . . . . . . . . . . . . . . . 20

5.3.1 Design Drivers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

5.3.2 Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

5.3.3 Added Payloads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

5.3.4 Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

5.3.5 Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

6 Science 24

6.1 Extended Boom and General Sensors . . . . . . . . . . . . . . . . . . . . . . 24

6.2 Analysis of the e↵ectiveness of Melanin Nano-shells . . . . . . . . . . . . . . 24

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6.3 The E↵ects of Deep Space Radiation on the Growth of Plants . . . . . . . . 24

6.4 Crew Activity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

6.4.1 Psychological evaluation . . . . . . . . . . . . . . . . . . . . . . . . . 25

6.5 Laser-based Communication System Demonstration and Mars Reconaissance 25

6.5.1 Technology Demonstration . . . . . . . . . . . . . . . . . . . . . . . . 25

6.5.2 Scientific Observations . . . . . . . . . . . . . . . . . . . . . . . . . . 25

6.5.3 Earth Re-entry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

7 Launch Vehicles 26

7.1 Launcher Comparison . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

7.2 Maximum Number of Launches and LEO Launch Mass . . . . . . . . . . . . 27

7.3 In-Orbit Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

7.3.1 Delta-V Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

7.3.2 In-Space Propulsion System . . . . . . . . . . . . . . . . . . . . . . . 28

8 Reentry 29

8.1 Reentry Vehicle Configurations . . . . . . . . . . . . . . . . . . . . . . . . . 29

8.2 Reentry Trajectory . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

8.3 Heat Flux . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

8.4 Heat Load . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

8.5 TPS material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

9 Electrical Power Systems 33

9.1 Dragon . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

9.1.1 Power Generation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

9.1.2 Energy Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

9.2 Sundancer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

9.2.1 Power Consumption . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

9.2.2 Life Support . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

9.2.3 Waste Management . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

9.2.4 Communications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

9.2.5 Crew Activities / Science . . . . . . . . . . . . . . . . . . . . . . . . . 35

9.3 Power Generation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

9.3.1 Case A . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

9.3.2 Case B . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

9.4 Energy Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

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10 Cost Evaluation 37

10.1 First approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

10.2 Detailed Cost Study - QuickCost 5.0 Model . . . . . . . . . . . . . . . . . . 37

10.2.1 Main Spacecraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

10.2.2 Communications Satellite . . . . . . . . . . . . . . . . . . . . . . . . 38

10.3 QuickCost Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

10.4 Cost Scheduling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

10.5 Validation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

10.5.1 Communications satellite . . . . . . . . . . . . . . . . . . . . . . . . . 39

10.6 Main spacecraft - NAFCOM Weight-Based Model . . . . . . . . . . . . . . . 40

11 Mission Phases and Scheduling 42

11.1 Development Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

11.2 Launch and Docking Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . 43

11.2.1 Launch and mass distribution . . . . . . . . . . . . . . . . . . . . . . 43

11.2.2 Launch sequence . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

11.2.3 Docking and in-orbit assembly . . . . . . . . . . . . . . . . . . . . . . 44

11.3 Interplanetary Events . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

11.4 Risk mitigation and 2021 launch window . . . . . . . . . . . . . . . . . . . . 47

12 Conclusion 48

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1 Report Overview

The following report proposes a design for a two-person Mars fly-by mission to be launched

in 2018. Building on the foundation laid out by Inspiration Mars, Cranspace has sought

to resolve areas of weakness and propose the designs outlined in this report. The key ar-

eas addressed include the provision of a suitable environment for the crew, ensuring safety

from radiation, the identification of science experiments that can be carried on board, and

an analysis of the re-entry phase, whilst ensuring that the cost of the mission is as low as

possible. Other areas such as launch vehicles and propulsion systems are discussed as they

need to accommodate these systems.

As minimising the cost of the mission is one of the main overall objective, o↵-the-shelf

technologies and structures have been incorporated wherever possible. One such structure is

the Sundancer Inflatable Habitat by Bigelow Aerospace. As Bigelow has ceased development

of the Sundancer in favour of developing the larger BA 330 module, the cost analysis in this

report has allocated funding for the completion of its development []. However, this cost

will be greatly reduced as Bigelow has used similar technology to develop the BA 330. The

Sundancer will contain six bubble rooms that will accomodate the astronauts. The purpose

of these bubbles is primarily radiation protection. They will be surrounded by water, which

will be transferred between one bubble and the next as the crew moves between them. This

system will aim to minimise the amount of water needed for shielding on the spacecraft. For

further shielding, the waste management system will treat solid waste and store within tubes

around the spacecraft. The radiation shielding and waste management systems are areas for

which development costs will be higher.

The mission will also carry a small satellite carrying a laser communication payload that

will enable a high data transmission rate during the interplanetary trajectory. This satellite

will be left in Martian orbit as a future communications relay from the Red Planet.

SpaceXs Dragon capsule was chosen as a re-entry vehicle. Minor modifications will be

required to resist the high re-entry velocity. A literary survey of materials used in re-entry

has been made.

The mission will employ four Falcon Heavy launches, and will cost an estimated 2.1 to

2.85 billion FY14 U.S. dollars. There are a number of constraints that have, so far, limited

man’s reach in the solar system and a great many considerations that need to be made in

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order to ensure the well-being of those that take the risk to travel out there. This mission

is intended to act as a technological stepping stone, demonstrating the necessary techniques

and providing the technologies as the next phase for the human race to finally fulfil the

universal dream of establishing an extraterrestrial human colony on Mars.

This report considers how to overcome the constraints of interplanetary travel such as galac-

tic cosmic rays, solar particle events, waste management problems and the associated cost

with a mission of this magnitude. Configurations of primary subsystems are proposed to

suit the requirements of the astronauts. Food management is investigated and scientific

experiments are suggested to take advantage of the spacecraft’s unique position. Following

a comprehensive literary survey of Launch Vehicles, this report outlines the requirements

and parameters of the launch configurations, engine specifications, capabilities and payload

constraints.

For additional information, simulations and videos, please visit our website: http://www.cranspace.com

2 Inflatable Space Habitat

CranSpace design proposes the implementation of the Sundancer inflatable module [1] de-

veloped at Bigelow Aerospace [2]. Bigelow Aerospace are currently preparing for the first

rendezvous of an inflatable spacecraft with the International Space Station [3] providing a

system demonstration of the technology to be implemented on-board the Bigelow BA 330 [4]

and the Sundancer [1]. The Sundancer has been designed to accommodate up to 3 people

in Earth Orbit. This section considers the advantages and disadvantages of implementing

the Bigelow Sundancer module and possible modifications to ensure that it is suitable for

long-term deep space flight.

2.1 General Assumptions and Considerations of Sundancer

The assumptions made of the Sundancer’s capabilities are presented below and are based on

the specifications outlined by Bigelow Aerospace [2] (see table 1).

2.1.1 Specifications

The assumptions made of the Sundancers capabilities are presented below and are based on

the specifications outlined by Bigelow Aerospace (see table 1).

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Occupancy Up to three on a long-term basis

Volume 180 m3

Radiation Protection Bigelow Aerospaces shielding is equivalent to or better than the Inter-

national Space Station and substantially reduces the dangerous impact

of secondary radiation.

Ballistic Protection The Sundancer utilises an innovative Micrometeorite and Orbital De-

bris Shield. Hyper-velocity tests conducted by Bigelow Aerospace have

demonstrated that this shielding structure provides protection superior

to that of the traditional aluminium can designs.

Propulsion The Sundancer utilises two propulsion systems on the fore and aft of

the spacecraft. The aft propulsion system can be refuelled and reused.

Electric Power Every Sundancer habitat will include an independent power system

comprised of solar arrays and batteries.

Avionics Each module will contain an independent avionics system to support

navigation, re-boost, docking, and other manoeuvring activities.

Environment Control and

Life Support System

Each Sundancer will contain its own independent ECLS system includ-

ing lavatory and hygiene facilities.

Windows Sundancer will boast four large windows coated with a film for UV

protection, providing an unparalleled opportunity for both celestial and

terrestrial viewing.

Table 1: Specification of the Sundancer module [1]

Radiation Protection The radiation shielding specifications provided by Bigelow Aerospace

are assumed to be limited as there is no publicly available documentation regarding its capa-

bilities. Whilst it is true that testing on board the ISS will occur, radiation levels within LEO

are considerably lower than those experienced beyond the natural shielding of the Earth’s

magnetic field [5]. It is therefore necessary to consider additional passive and possibly active

radiation shielding in order to provide a degree of protection in the case of Solar Particle

Events (SPE) (see section 3).

Ballistic Protection It is assumed that the ballistic protection provided by the Sundancer

is capable of shielding the spacecraft from micro-meteoroids whilst in deep space. The

only additional ballistic protection methods that will be considered are collision avoidance

manoeuvres, rather than additional passive ballistic shielding.

Propulsion It is assumed that the in-built systems are su�cient for manoeuvres such as

collision avoidance, as well as emergency procedures that are initiated in the case of solar

flares and coronal mass ejections (see section 7).

Electrical Power The electric power is sourced from solar arrays and batteries. It has

been assumed that the solar arrays and batteries can be configured to satisfy the electrical

power requirements of the mission. A assessment into the required sizing of the solar arrays

and batteries is made and presented (see section 9).

Life Support While the life-support systems have limited available specifications, they

are undergoing initial human trials [6]. Life-support systems are therefore expected to be

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ready by 2018. Modifications are expected to occur in order to accommodate additional life

sustaining measures (see section 4).

Figure 1: Design outline of the Sundancer Inflatable module. Developed with Sketchup.

2.2 Advantages and Disadvantages

Volume The Sundancer inflatable module provides the astronauts with a dramatically

larger available volume that can be used for both storage and living space. This spatial

advantage provided by Inflatable Habitats over more traditional ”can-like” structures is an

important one, especially considering the long flight duration. Cramped conditions similar

to the ATV [7] or Orion [8] would dramatically reduce moral on the spacecraft and such

continuous close proximity contact between astronauts may induce conflict over time in

what will be an already stressful environment. Additionally, Bigelow Aerospace inflatable

spacecraft have been designed to be inflated in-orbit after being stored in low volume rocket

fairings.

Developed Sub-Systems By incorporating the Sundancer, a large number of the sub-

systems, including the critical life support system, are being independently developed by

Bigelow Aerospace, relieving development costs. Modifications are expected, however, and

coordination with Bigelow Aerospace is needed to tailor aspects of various system designs.

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Multilayer Structure The multilayer structure nature of the Sundancer’s exterior acts as

an external thermal blanket, helping to keep the spin-stabilised vehicle’s interior at a constant

temperature while shielding the spacecraft’s systems and inhabitants from micro-meteoroids

and space debris. Studies suggest that inflatable structures provide better shielding than is

available on-board the ISS and also provides significant protection against radiation [9]. A

survey of these studies indicate that the exterior structure is to be made from kevlar-like

materials, Vectran, Nexttel and polyethylene.

Development Status The Sundancer’s development has been frozen as Bigelow Aerospace

have focused their attention to larger vehicles. However, there have been continuous advances

of the sister BA 330 vehicle which is expected to be launched as early as 2015 [3]. The Sun-

dancer is expected to use near identical systems implemented by the BA 330 and so, if

Bigelow Aerospace could be convinced (financially) into resuming development procedures -

the design could be fast-tracked using technology demonstrated by the BA 330.

The overall layout of the internal structure of the Sundancer is expected to be modified

in order to address mission constraints such as space and shielding requirements. Given the

nature of the inflatable of the mission, much of the assembly of the internal components

need to be made in-orbit. This is explained in detail in the sections related to the individual

sub-systems.

3 Radiation Shielding

One of the most di�cult obstacles to overcome in interplanetary travel is the need to pro-

vide adequate shielding to protect crew and equipment from radiation. The main radiation

sources of radiation in deep space are heavy ions fully ionised atomic nuclei of galactic

cosmic rays (GCR) and energetic protons originating from SPE. This secondary radiation

e↵ects occuring after initial exposure. Without adequate protection, astronauts can su↵er

from acute radiation syndrome (ARS) or radiation sickness, damage to the central nervous

system, cancer and death. In practice, such shielding often requires a large mass and high

development costs. NASA regulations state that astronauts can be exposed to no more than

3% of added Risk of Exposure-Induced Death (REID) [10].

5


Figure 2: Image illustrating approximate values of radiation levels with increasing distance fromthe Earth’s surface. Units are displayed as a log scale of annual levels of rems. Image taken fromScientific America.

The spacecraft will traverse a number of di↵erent radiation environments through the du-

ration of the mission, most notably the Van Allen Radiation Belts (see Figure 2). These

consist of a large number of highly energetic charged particles [11]. In the Apollo missions,

astronauts who traversed the belt regions were largely unprotected from this radiation. De-

spite their limited exposure time, they received the majority of their total mission dose while

passing through the inner belts. This dose varied through a range of 1.6 to 11.4 mGy less

than the limit set by the United States Atomic Energy Commission for people working with

radioactive materials [12]. A similar approach is used in this mission design, whereby the

spacecraft performs its delta-v before passing through the belts, so as to minimise the time

spent within them.

3.1 Radiation and Water

The level of radiation protection provided by a material is measured by the extent to which

it is penetrated by energetic particles [10]. Other important factors that must be taken into

account include the chemical properties of the material and the way in which the human

inhabitants must manage and interact with it. Aluminium is often considered as the material

of choice, but has a high mass and still produces secondary neutrons following exposure to

radiation. Polyethylene, a hydrogen rich material, has been successfully tested on board

the ISS [13] and is therefore a candidate for radiation protection. Water is also capable of

absorbing a high amount of energy while emitting a low number of neutrons, although it

has not been used in previous missions due to the large amount that would be required. It

was also deemed unsuitable for exterior mounting on a spacecraft due to its structural and

thermal properties [14]. There are, however, a number of advantages to using water to shield

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against both GCRs and SPEs. Figure 4 compares waters shielding properties to those of

other materials.

Figure 3: Dose equivalent as a function of wa-ter shield thickness resulting from galactic cos-mic rays [15].

Figure 4: Point estimates of 5 cm depth dosefor GCR at Solar Minimum as a function ofareal density for various materials [16]

.

Water has the added advantage of being in a liquid state at room temperature and atmo-

spheric pressure. It can therefore easily be transported around the spacecraft in order to

provide continuous and varying levels of shielding thickness in di↵erent parts of the habitat,

depending on the local radiation levels. For future designs, this concept could be used as

an adaptive shielding mechanism that could provide a thin level of shielding for regions of

space with low incident flux, and thicker shielding for regions with higher flux. In order to

reduce complexity, a simple approach has been used within this report. It is worth noting

that whilst moderate amounts of shielding reduce the dose of GCRs due to the removal of

lower energy components, the e↵ectiveness of shielding approaches diminisihing returns be-

yond 25 cm2 (see Figures 3 and 4). This design uses water as a shielding medium within the

Sundancer - essentially acting as a secondary shielding layer for the entire habitat module

beneath its external structure. It has been previously mentioned that the radiation shielding

on board the inflatable modules provided by Bigelow Aerospace is considered as incapable

of su�ciently shielding astronauts in deep space. Whilst the ballistic and radiation shielding

properties of the materials used are to be demonstrated on the ISS by 2015 [3], it will not be

subjected to the same levels of radiation as experienced in deep space. Even so, information

regarding the structural properties of the Sundancer is limited, correlations have been made

with previous studies and reports concerning inflatable structures in space. The European

Space Agency’s (ESA) Aurora Exploration Program has provided a considerable amount of

information regarding the comparison between inflatable and rigid vehicle structures, and

7


states that the material structures of inflatable vehicles is expected to be based on polyethy-

lene, polyacrylate, ethylene vinyl alcohol (EVOH), Kevlar and Nomex [9]. The thickness

of the Sundancer’s external layer is expected to be around 40-50 cm when fully inflated.

According to both the Hadronic-Bertini and the Hadronic-Binary physics models, the level

of shielding provided by this structure would be comparable to a similar thickness of water,

as shown in Figure 4. Accounting for the various considerations listed above, an internal

radiation-resistant housing system has been designed, as shown in Figure 5.

Figure 5: Simplistic design of the encased Bubble Structure within the Sundancer Inflatable Module.Image developed using Sketchup.

Radiation can approach the spacecraft from any direction. The system’s additional shielding

comes in the form of dome-shaped ”bubble” rooms that are surrounded by water to protect

the astronauts from incoming GCRs. As the astronauts move into neighbouring rooms, the

water is flushed into the relevant bubble. This method will provide the astronauts with

adequate shielding in all directions at all times, while still taking advantage of the space

available inside the Sundancer. The general description of this system can be separated into

two sections: the cylindrical core structure and the bubble modules.

I. Cylindrical Core This is a component of the Sundancer’s internal structure which

must be modified prior to launch and adapted in order to accommodate mission critical

components such as the life-support and central computing systems. The modified core acts

to support the network of tubes which provide a connection between the bubble modules,

the waste management system and the water systems. The core itself is segmented into 3

sections separated by a hatch. Each section is connected in parallel to the life support and

central computer systems so that all bubble structures and core sections have access to the

on board data handling system as well as a continuous supply of oxygen. At each end of

the core, any docked structures (e.g. the Dragon capsule) can be accessed via a pressurised

hatch. A comparison between the original and modified core structure is presented in Figure

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6. The modified core is structured so that the original four separate pillars are combined

to provide two larger support pillars in order to allow for easy deployment of the bubble

modules (see Figure 7). The gaps in the core allow the astronauts to enter and exit the

’bubbles’ while providing the same, continuous level of protection.

Figure 6: A simplified design of the original core (left) compared with the modified core structure(right). This figure illustrates the gain in space available to the astronauts when transitioningbetween bubble structures.

II. Bubble Modules: General Configuration Each pair of bubbles is assembled so

that they are oriented on opposite sides of the central core. Two larger sets of bubble mod-

ules will have external and internal diameters of 2.9 m and 2.4 m respectively, providing a

shielding thickness of 25 cm (or 25 g/cm2) and a smaller bubble set will have an external and

internal diameter of 2.56 m and 2 m respectively, providing a shielding thickness of nearly 30

cm. The overall configuration of the bubble pairs is described in Section 4.1. The bubbles

have been designed with the following considerations in mind: robustness, ease of repair,

ease of construction, ease of stowing, ease of water transportation, simplicity and cost. The

design is based on the truncated icosahedron structure - essentially two tiled buckyballs, the

larger one making up the exterior surface and the smaller one providing the interior surface.

The final configuration is depicted in Figure 7. The surface tiles are composed of 3mm thick

composite panels. The opening of the structure (the bottom face in Figure 7) is directly

connected to the core of the Sundancer. The tiles of adjacent bubble pairs are connected to

allow water to flow between them. The arrangement of the tiles is shown in 8. The edges of

the tiles are lined with solid poles, which will provide the structure with a level of rigidity

once mounted. To ensure adequate ventilation, small gaps in the surface are provided, with

fans to allow for the circulation of air.

This design allows the structure to be folded so that it is compact. These bubbles will

then be assembled onto the core before launch. The dimensions of the plates have been

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designed so that the overall width of the system is less than 40 cm when stowed. Each group

of bubbles is then folded in order to minimise the deformation of the flexible pipes between

the blocks. The assembly procedure is as follows: following the docking of the Dragon to

the Sundancer, and inflation of the latter, the astronauts will enter. They will then connect

each set of bubbles to the core. Their edges will then be connected as described above.

Figure 7: Illustration of the bubble’s filled and empty constituent blocks.

The walls of each bubble can be compressed in order to expel its water contents. This

controlled compression will be performed in a predefined sequence, with the furthest away

from the core being the first. This will ensure that no trapped water remains in the structure.

Once the water has been flushed and all the structures emptied, the valves will be shut.

Figure 8: The illustration on on the left provides the mechanims for directing water within atile group. The arrows represent a force ontho the surface compressing the water and forcing adisplacement out of the tile group. With the valve open, water is confined to flow in the directionshown by the arrow. These groups are laid out as shown on the right.

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3.2 Waste Management System

Human waste produced during the journey must be stored and managed carefully. Current

methods used on board the ISS involve disposing of solid waste via a freighter spacecraft.

Paragon Space Development Corporation has proposed a system that uses ionomer mem-

brane ”bags” [17] to extract water from human waste through reverse osmosis. These bags

can be used to provide a source of radiation shielding by lining the spacecraft walls with

them. However, the need for astronauts to manually interact with faecal matter raises a

number of health issues. The following solution builds on this idea while employing a simple

yet innovative system that provides both radiation shielding in an automated and e�cient

manner to manage waste.

3.2.1 Waste Collection and Delivery

The flow mechanism of the delivery system relies on a peristaltic motion to push waste

through flexible tube structures. This method was inspired by mammalian intestines. The

peristaltic motion is generated by the activation of concentric muscles, and drives solid waste

through the digestive system. The same principle will be applied to the waste delivery system.

This system utilises the same material that makes up the ”bags” developed by Paragon.

However, rather than a bag-like container, the material is extended so that it forms a tube-

like structure with an approximate length of 1.5 km and a diameter of 2 cm (see table 2).

The ionomer membrane slowly extracts water from the system through reverse osmosis. This

water is then collected for filtration and reuse. The tube is contained in an external vessel

designed to insulate the waste-containing inner tube. The peristaltic motion of the waste

management system is induced by inflatable rings positioned at intervals along the tube.

Positioned at intervals along the tube the synchronised contractions and relaxations along

the conduit induce the forward motion of the solid waste (see Fig. 9). The total e�ciency

of the water extraction procedure increases as the waste travels further along the tube, as

the contact surface increases allowing for more water to be absorbed.

Inner pipe diameter 2 cm

Outer pipe diameter 5 cm

Dry mass 570 g/m

Loading mass 2500 g/m

System Length 1.5 km

Waste volume 0.5 m3

Table 2: Dimensions of the waste management delivery system minus integration with space toilet.

11


Figure 9: Illustration of the artificial intestine. Schematic representation of the waste managementsystem On this schematic, only the inner pipe, toric ring and bracing to maintain in position ringsare display. Pipe which delivers the pressure, micro valves and external pipe are not displayed.

Solid waste delivery requires water input into the system in order to reduce static friction

along the surface of the inner tubes. The water supply for waste management is isolated

from all other systems.

An additional development cost needs to be taken into account in order to test the integrity

of the ionomer membrane within our proposed system. However, the design strengths fulfil

criteria regarding hygeine and storage, whilst additionally contributing to radiation shielding.

3.3 Application of Melanin

Melanin is an organic molecule capable of absorbing radiation (UV to gamma-rays). Dr

Dadachova at the Albert Einstein College of Medicine, University of Yeshiva, in New York,

demonstrated that it is possible to use Melanin for radiation shielding [18]. The outermost

tube will be coated with a layer approximately 3.5mm thick. This will provide additional

protection to the crew. This total 7mm thickness is intended to reduce penetration by inci-

dent gamma rays by up to 50%.

The Linear Attenuation Coe�cient of Melanin µ, which quantifies the quality of a radiation

shield, is 1.01 cm�1. From this, the Half Value Layer thickness is given by the following

equation:

HLV =1

µln2 = 0.68cm (1)

12


Name µ (cm-1) HLV (mm) Density (g. cm-3) Weight (g. cm-2)

Melanin 1.01 7 1.33 0.931

Lead 27.1 2.5 11.34 2.835

Aluminum 0.384 18 2.7 4.87

Table 3: A comparisson of HVL values between various materials.

3.4 SPE Events

Solar Proton Events produce extremely high radiation levels, and can result in instant radi-

ation induced death. The mission will take place during a solar minimum (year 2018), and

hence very few SPEs are expected. However, precautions should be taken to prevent poten-

tially catastrophic results. In the case of an SPE, the crew would be alerted by the ground

station, who would continue to provide observations and analyses. This would provide the

crew with approximately a day to prepare the spacecraft’s emergency orientation (see 3.5)

and begin preparations of the SPE shelter (see 3.6).

3.5 Magnetic Field Probe

A magnetic field probe has been designed to be deployed on a telescopic boom of approxi-

mately 8 m during the beginning of the interplanetary travel phases to monitor the environ-

ment’s interplanetary magnetic field (IMF). This would cheaply provide detailed information

of the magnetic field lines and strength. In the event of an SPE, the spacecraft would then be

able to properly align the SPE shelter (see Section 4.7) to counter the incident flux. Data col-

lected throughout the mission would also provide useful scientific data to potentially develop

or contribute to existing IMF models.

3.6 SPE Shelter

Small modifications are required to convert the cylindrical core into a SPE shelter. The

procedure in the event of an SPE is as follows: the onboard water used for radiation shielding

is to be evacuated from the bubbles to fill an enclosed segment of the core (see section 3.1).

13


Figure 10: Flushing water from the bubble to the core to provide a water shielding thickness ofapproximately 3 m in length oriented in the direction of the incident radiation. The crew are keptinside a hatch which itself adds material protection in all directions. A full design for this hatchwould need to be optimised in a way to minimise radiation dose.

4 Internal Housing and ECLSS

The housing structure has been designed to keep the inhabitants continuously protected

from GCRs. As such, the overall housing which is encased within the Bigelow Sundancer

module is centered around the two radiation shielding components described in sections 3.1

and 3.2. This section describes the general configuration of the internal housing system and

discusses how it is suited for long-duration deep space flight.

4.1 Bubble Configuration

The configurations of the bubble structures are dependent on the requirements of the inhab-

itants. The bubbles provide what are essentially ”rooms” and each can be customised to

suit the room requirements, for example the sleeping space requirements are di↵erent from

the exercise space requirements. The total dimensions of the bubbles must fit within the

Sundancer with some margin allowed for other sub-systems, such as the waste management

system. The final configuration (see Figure 11) consists of three sets of rooms: two large

sets and one slightly smaller. The smaller rooms are allocated to the the sleeping quarters

and lounge or leisure areas as they do not require as much space and benefit from thicker

shielding. Each room pair is separated by a hatch that allows access to the next room as

water is flushed between them.

14


Figure 11: Illustrates the configuration of the bubble housing structure together with the specificationfor the two di↵erent types of bubbles

.

Rooms The following room allocation is provided to accommodate the various activities of

the astronauts during the trip to Mars. The largest choice for the bubble dimensions are 2.5

m. The design also makes use of a smaller bubble set at just 2 m long. This is to be used as

a dormitory which does not require as much volume while allowing thicker shielding. With

the chosen bubble sizing and accommodation requirements, the next step is to distribute

the space for all the activities that will be conducted during the mission. The main areas

include:

1. Hygiene and WC: This includes a shower approximately 2x1.5x1 m. Toilet dimensions

are 0.5x0.5x1 m, with extra space to provide integration with the water and waste

management system.

2. Medical and Exercising: It is important that astronauts do enough exercise to main-

tain acceptable bone mass. In addition to exercise equipment, medical equipment is

provided on board, not only for possible contingencies but also to take regular tests

and provide health monitoring.

3. Kitchen and Dining: This area are used for food preparation and dining. It would also

provide space for food storage, e↵ectively acting as a pantry(see section 4.2.2).

4. Science: Carrying out scientific experiments remains a goals of the mission. A dedicated

area has been assigned to accommodate computer systems and a small space laboratory

for technology development and scientific research (see section 6).

5. Sleeping quarters: Contains the beds and most of the personal belongings.

6. Lounge/Communications: Provides an area of comfort to the astronauts and a place to

relax, where entertainment systems could be included. Additionally, this room acts as

15


a communications post from which astronauts would transmit back to ground stations

on Earth. This post can be accessed remotely from all parts of the spacecraft.

4.2 Food Management

4.2.1 Food Requirements

For a long duration mission to Mars, managing the astronauts food is essential to their

survival and the success of the mission. It is important to consider the nutritional content,

the variety of food and ease of preparation in zero gravity. In space, the average persons

intake of solid foods is 0.62 kg per day; hence, a minimum of 621 kg of solid food must be

taken with the two astronauts [19]. The total mass estimate of the entire food management

system, including 650 kg of dehydrated food, packaging, and preparation equipment, is 1458

kg [20]. This food will consist of a large variety of foods to maintain morale throughout the

mission. It must also ensure that the astronauts get the required nutrients, minerals, and

other compounds to keep them strong and healthy. The nutritional value required daily for

a male astronaut of 98.5 kg and a female astronaut of 61.7 kg is shown in Table 4 [21].

The body masses of the astronauts were assumed using the 95th percentile male and fe-

male from [21], from which the daily intake values were calculated assuming a requirement

of nutrients and energy on a per body mass basis. The 95th percentile mass values were

used to ensure that su�cient food was on board to account for any delays in the mission or

di↵erences in astronaut metabolism. Using these values, large bags of dehydrated food can

be arranged to consist of the weekly amount of food required for the astronauts. The food

will be carefully planned to provide the required weekly minerals and energy.

Daily Male Req. Daily Female Req. Total Weekly Req.

Energy (kJ) 12950 9600 22550

Protein (g) 78.8 49.36 897.12

Fat (g) 128.05 80.21 1457.82

Carbohydrates (g) 472.8 296.16 5382.72

Phosphorus (g) 0.8 0.8 11.2

Sodium (g) 3.5 3.5 49

Iron (g) 0.018 0.018 0.252

Calcium (g) 0.8 0.8 11.2

Magnesium (g) 0.35 0.35 4.9

Potassium (g) 2.7 2.7 37.8

Table 4:The recommended daily nutrient intake for a male and female astronaut [19], as well as the totalweekly nutrient intake of both astronauts.

16


4.2.2 Food Storage

Food storage was a closely assessed issue as food is one of the few resources that cannot be

recycled. The total dehydrated food volume for the mission’s 501 days, plus contingencies, is

calculated to be 14 m3. This 14 m3 is approximately the volume provided by a single small

bubble structure and would take up a considerable amount of useful space. Instead, approx-

imately two thirds of the food is kept in the Dragon capsule with the remaining third kept in

the kitchen bubble. This would mean access to the Dragon capsule would only be required

two or three times throughout the mission’s interplanetary phase and poses minimal risk of

damaging the capsule. An analysis of the food volume for 8 weeks worth of dehydrated food

showed that less than 35% of the kitchen bubble would be filled while requiring only 2 added

entries into the dragon. Despite the consumed space of the bubble, Figure 12 illustrates that

there is su�cient space within the module to still cook and eat.

The option of storing food in the surrounding volume of the bubble structure to be used as

an additional layer of shielding has been discarded. This would require the astronauts to

rely on mechanical procedures in order to reach the food. Should something occur to the

retrieval mechanisms, exposure to harmful radiation levels would occur as the astronauts

would be required to exit the internal housing structure.

Figure 12: Kitchen Illustration. Despite the food storage taking considerable space within the mod-ule, this illustration shows that there is su�cient space to accomodate both individuals at the sametime.

4.3 Waste Management Configuration

The waste management system was outline in Section 3. The tube system is shipped within

the Sundancers structure. Post inflation, the tubes must be position manually on a face of

the Sundancer. This would e�ciently be handled by providing a layer of velcro along the wall

and simply strapping in the system of tubes in order to provide a completely covered walled

17


that would provide additional protection for CGRs (shielding from one direction only) and

SPEs (see Figure 10 - left).

5 Communications

5.1 Traditional Communications System

The spacecraft is equipped with standard X and Ka Band transceivers for TT&C and low

bandwidth video respectively. A preliminary link budget is shown below, with the prop-

agation path length taken to be 1.5AU, approximately the maximum distance from the

spacecraft to Earth. Low gain antennae will also be included for redundancy. Note that if

deemed necessary, the required data rate could be dropped to allow for lower transmitter

power or antenna size. A link budget for the laser system can be found in [22].

X Band Link Frequency (GHz) 8

Transmitter Power (W) 550

Transmitter Power (dBW) 27.40

Transmitter Gain (dB) 42.71

Spacecraft Antenna Diameter (m) 2.2

Propagation Path Length (km) 2.24 ⇥ 108

Free Space Loss (dB) �2.78 ⇥ 102

System Losses (dB) -5

Receiver Gain (dB) 72.77

System Noise Temperature (K) 200

Minimum Data Rate (Mbps) 1

Eb/N0 (dB) 5.95

Required Eb/N0 (dB) 3

Link Margin (dB) 2.95

Table 5: Communications budget for tradi-tional communications systems.

Ka Band Link Frequency (GHz) 30

Transmitter Power (W) 300

Transmitter Power (dBW) 24.77

Transmitter Gain (dB) 54.20

Spacecraft Antenna Diameter (m) 2.2

Propagation Path Length (km) 2.24 ⇥ 108

Free Space Loss (dB) �2.89 ⇥ 102

System Losses (dB) -5

Receiver Gain (dB) 84.25

System Noise Temperature (K) 200

Minimum Data Rate (Mbps) 7.5

Eb/N0 (dB) 6.05

Required Eb/N0 (dB) 3

Link Margin (dB) 3.05

Table 6: Equipment Power Consumption

Constraints of Traditional Communications System There are a number of con-

straints in using conventional RF antennas for communication over such long distances. RF

systems su↵er from increasing propagation losses with increasing separation between the

transmitter and receiver. While RF systems can benefit from high gain antennas, the power

requirements are still considerable, especially when transmitting across the distances con-

cerned here. With conventional systems, high data rates for audio and video transmission

(e.g. video feed of Mars) is extremely di�cult. However, free-space optical communications

using infra-red lasers have been identified by NASA as a technology that would provide

significantly higher performance data return rates [23]. A proposed laser communication

system is provided in Section 5.2.

18


5.2 Remote Communications System

It can be di�cult to justify the costs, financial and otherwise, associated with such a high

profile mission when the only objectives are to perform a flyby of Mars and return to Earth.

Without spending any significant period of time in Martian orbit, the scientific opportunities

o↵ered by such a mission are restricted to the study of human health and biology in space.

Although this is certainly a worthwhile area of research, the value of the mission could be

significantly increased by supplementing it with a smaller mission that would take advantage

of the unique opportunities on o↵er. It would be additionally beneficial if this secondary

mission could be carried out in conjunction with other organisations - this might allow both

the costs and benefits of such a mission to be shared. One such organisation is Mars One [24]

who have announced plans to place a telecommunications satellite in areostationary orbit by

2018 [25], therefore, a successful mission would certainly be of interest to them.

In recent years, a number of missions with the aim of demonstrating free-space laser com-

munications have been proposed by NASA - most notably the Lunar Laser Communications

Demonstration (LLCD), which flew on board the LADEE spacecraft in 2013 and successfully

demonstrated a downlink data rate of 622 Mbps from lunar orbit [26]. In 2009 the design of a

spacecraft that would demonstrate laser-based communications between Earth and Mars [25]

was cancelled due to budget constraints. Its payload, known as the Mars Laser Communi-

cation Demonstration system (MLCD) was expected to achieve a downlink data rate of up

to 30 Mbps [27]. Subsequently, a modified design - the Deep space Optical Terminal (DOT)

was designed to include the laser communications system along with a stabilisation platform

and was designed to be adaptable and fit onto generic Mars orbit vehicles [22]. While the

design has been shown to be feasible, there has not yet been a mission proposed that would

make use of the system.

Laser Communications Satellite The design proposes building a small satellite incor-

porating the DOT system that would be deployed in space to fly alongside the primary

spacecraft operating as a high performance data relay between the spacecraft and Earth

when conditions are favourable. This tandem flight was selected due to the high pointing re-

quirements incurred by the DOT system, which would be extremely di�cult to achieve with

the main spacecraft. During the mission’s flyby stage, the satellite will manoeuvre itself into

a 500 km altitude orbit about Mars to carry out a Mars observational mission while contin-

uously demonstrating DOT technology. Additionally, the spacecraft would also be equipped

with a standard Electra telecommunications package [28], allowing it to function as a data

relay for existing and future Mars orbiters and landers. Section 5.3 presents a baseline design

19


for this secondary spacecraft, detailing the main requirements and constraints that influence

the spacecraft design. This is followed by a brief overview of the di↵erent subsystems and

mission phases. The technical details of the laser communication payload are provided by

Biswas [22]. The overall mass budget is included in Section 10.6.

5.3 Communication Satellite Design

5.3.1 Design Drivers

Pointing Accuracy Constraint The primary constraint of using a laser system with

a transmitted-receiver separation of this scale is the necessary Attitude and Orbit Control

Systems (AOCS) to keep the spacecraft pointing at the Earth. A laser communications

link between Mars and Earth would require transmitter pointing accuracies on the sub-

microradian level [29]. This combined with an additional need for a milliradian level point-

ahead angle to compensate for the relative movement of Mars and Earth, as well as the

di�culty of implementing closed loop control due to the round-trip time delay places a

demand for extremely high pointing accuracy and stability on the entire system. While the

DOT system includes a stabilisation platform that shoulders much of this responsibility, the

spacecraft must still be able to point the laser terminal towards Earth with around 3 mrad

accuracy [22].

Orbit and Lifetime The mission lifetime chosen is 2.5 years, which should ensure a high

volume of scientific data return and justify development costs. The selected circular orbit has

an altitude of 500 km with an arbitrary inclination. While areostationary orbits are preferred

for communication satellites around Mars, it was decided that an orbit that saw more of the

planets surface would provide more useful data return, particularly from the cameras on-

board. 500 km was chosen because it was low enough to allow for high-resolution imaging

from the cameras without requiring large amounts of propellant to maintain altitude.

5.3.2 Design

As with the main mission, cost and technology readiness level are key factors. For this

reason, existing technology has been used throughout the design. As mentioned earlier, the

laser system is based on the design of the DOT, which consists almost entirely of components

already used in terrestrial optical communications (see [22] and [30] for information on the

system and stabilisation platform respectively). The rest of the spacecraft design uses many

of the components previously used on the Mars and Lunar Reconnaissance Orbiters, since

they are recently built spacecraft of a similar mass and performance level to what is required

20


of this system. The spacecraft is to be stored in the unpressurised trunk of the Dragon

capsule before deployment.

5.3.3 Added Payloads

As mentioned previously, the payload consists of the DOT system, along with two narrow

angle cameras and a wide angle camera based on those of the Lunar Reconnaissance Orbiter

[31], and the Electra package [29]. Although the DOT design places the laser terminal

mounted to the body of the spacecraft, the terminal can also be mounted on a two- axis

gimbal for improved accuracy. This explains the slightly higher mass and power estimates

in the budgets below, compared to those given in the DOT design.

Figure 13: Schematic of Laser Flight Terminal from DOT Design. [32]

Cameras The two narrow-angle cameras on the LRO have been able to capture extremely

high quality (sub-metre resolution [31]) images of the Moon, and the high pointing accuracy

of this spacecraft will allow for unprecedented imaging opportunities of the surface of Mars.

The cameras will likely require some additional development in order to be used for this

mission, which has been accounted for in the cost analysis. The cameras are mounted to the

exterior of the satellite, pointing in the nadir direction. A wide-angle camera similar to the

one carried on the LRO mission is also to be adapted for this mission at little extra cost.

This is mounted in the same way as the narrow angle cameras.

Electra The Electra Proximity Link Payload has flight heritage, having already been car-

ried on the Mars Reconnaissance Orbiter and MAVEN spacecraft to date. Electra allows

for UHF communications links with data rates of up to 1 Mbps to be established between

spacecraft around Mars, enabling data relay and ranging capabilities throughout the network.

A detailed description of the package can be found in [28] The main physical components

21


of the system are a transceiver unit with dimensions of 17.2 x 21.9 x 14 cm and a small

nadir-pointing UHF antenna.

5.3.4 Structure

As with the MRO, the structure will consist mainly of carbon composite and aluminium

honeycombed plates. The structural mass was estimated based on the total spacecraft mass

using techniques given in [33].

5.3.5 Power

This spacecraft will use Gallium Arsenide solar cells and Lithium-ion batteries. A preliminary

power analysis has been provided in Figure 7 and it has been determined that the maximum

solar array area required is 13 m2. This value is extremely large, but this is primarily because

it has been assumed that all equipment is constantly running at full power- that is, duty

cycling has not been taken into account. Another reason for this high value is the value

for solar flux in the calculation is that experienced when Mars is at aphelion. The power

subsystem requirements could be dramatically lowered by reducing usage of high power

consumption equipment at such distances.”

Maximum night-time load (W) 662.5

Maximum Battery Discharge (Wh) 477.28

Acceptable Depth of Discharge (%) 40

Required Capacity (Wh) 1193.19

Power needed to charge batteries (W) 501.06

Maximum daytime power requirement (W) 736.1111

Power needed from solar arrays (W) 1237.18

Solar flux at maximum Mars-Sun distance (W/m2) 470

End-of-Life solar array e�ciency (%) 25.76

Required solar array area (m2) 13.48

Battery mass (kg) 9.55

Solar cell mass (kg) 37.06

Table 7: Power Sub-system Sizing

AOCS This subsystem was based on that of the MRO, which is of a similar size and can

achieve the pointing accuracy and stability required for this spacecraft. It uses 4 Honeywell

Constellation Series HR16 reaction wheels for attitude control, along with thrusters for un-

loading. Attitude determination is performed by the Galileo Avionica A-STR Star Tracker,

along with SST US 2-axis Sun Sensors and a Honeywell Miniature Inertial Measurement

Unit for inertial navigation. If this combination is deemed to be inadequate for the precise

attitude knowledge required, the SSBV Earth Horizon Sensor has also been identified as

an additional attitude determination sensor. Extra units of each of these sensors will be

carried for redundancy. Technical specifications for each of these components can be found

22


in [34] [35].

For a more detailed analysis of the AOCS and propulsion systems, please see www.cranspace.com.

Ground Stations Although there are no currently operational ground stations that are

capable of supporting a data link for this mission, it is extremely unlikely that a ground

station would have to be constructed from scratch. The LLCD uses a ground terminal which

includes a transmitter, consisting of an array of four 15 cm telescopes, and a receiver, which

consists of four 40 cm telescopes as pictured below. Although it is based primarily in White

Sands, New Mexico, the terminal is designed to be transportable. Although this terminal

does not have the capability of supporting a link from Mars (the DOT report specifies

a minimum receiver aperture area of 3.8 m2 [22], building an upgraded version with the

required capabilities would not be likely to require huge research and development costs. If

this secondary mission is supported by another organisation such as NASA or Mars One, the

cost of doing so would not fall entirely on the Inspiration Mars Foundation. However, due to

the extreme sensitivity of a laser communications link to factors such as cloud cover and other

atmospheric conditions, it is likely that the system would require multiple ground stations

in carefully chosen locations around the world, to ensure that at least one ground station

is reachable when links are to be established. The DOT report explored the possibility of

retrofitting existing ground-based astronomical telescopes to operate as ground stations [22]

[32].

Figure 14: The Lunar LaserGround Terminal [22]

Figure 15: Concept of Operations for the DOT [32]

23


6 Science

The variety of space environments encountered, as well as the numerous challenges faced by

the mission and its crew, will provide a unique opportunity for scientific research. As no

spacecraft of this size and mass has ever been sent to Mars, the Inspiration Mars mission

has the chance to carry a multitude of science payloads.

6.1 Extended Boom and General Sensors

A number of magnetometers and spectrometers externally mounted on the spacecraft will

provide valuable science data for the mission. It was stated in Section 3.5 that a magnetic

field probe would be placed at the end of a telescopic boom approximately 8 m long when

fully extended. A record of the magnetic field variations in deep space may be of scientific

interest. Additional sensors may also be included to perform brief experiments when passing

through the Van Allen belts.

6.2 Analysis of the e↵ectiveness of Melanin Nano-shells

The Melanin on board the spacecraft (coating the waste management system) will provide

the crew with high levels of electromagnetic radiation protection. A sensor to determine

the X-ray and gamma-ray flux will be located outside of the spacecraft in order to allow

for comparison with values measured within the bubble structures. A successful demonstra-

tion of the technology could considerably help the development of highly radiation resistant

structures.

6.3 The E↵ects of Deep Space Radiation on the Growth of Plants

For future long term space exploration missions, it is envisioned that plants grown in space

could play a major role in the biological life support system (BLSS) providing both clean

air and food [36]. Furthermore, the presence of plants has been noted to have a positive

e↵ect on the well-being of crew members [37]. However, all such experiments have been

performed in LEO under significantly di↵erent radiation conditions than those experienced

on a deep space mission. The Inspiration Mars mission is therefore an excellent opportunity

to examine the growth of plants and to assess whether plants grown in such conditions will

be suitable for human consumption.

[36] identifies the selection of the perfect plant material for BLSS by considering the rate of

oxygen and biomass production and the radiation resistance of various plant species as an

24


important step in the development of bio-regenerative life support systems. This is therefore

an important theme that can be studied on board. There are several types of equipment

currently being used on the ISS for the growth of plants in space, therefore no further de-

velopment costs will be incurred.

6.4 Crew Activity

The monitoring of crew activity on board can be achieved using widely used devices such

as the Fitbit [38] and Google Glass [39]). These wearable electronics, when integrated

with specifically developed applications could provide benefits such as Augmented-Reality

capabilities, remote voice control, imaging, audio and video correspondence with Earth and

entertainment. The development of applications for space for Google Glass hardware is

not expected to require considerable investment. The Fitbit could provide physical activity

monitoring and it could be of scientific interest to observe how this varies throughout the

duration of the mission.

6.4.1 Psychological evaluation

On board cameras with facial recognition software [40] to monitor the mental state of the

crew could also be utilised. This software evaluates the astronauts concentration, which aids

in avoiding potential human errors.

6.5 Laser-based Communication System Demonstration and Mars

Reconaissance

6.5.1 Technology Demonstration

In Section 5.3, an outline for the small laser-based communication relay satellite was pro-

vided. The satellite will demonstrate laser-based communications for deep space applications.

While a number of missions making use of optical communications have been proposed, the

application of the technology is not yet widespread. By demonstrating the technology over

distances far greater than previously tested, broad implementation in the space industry

could be achieved.

6.5.2 Scientific Observations

In addition to the demonstration, the satellite can be adapted to give additional benefits

at little extra cost. Outfitting it with cameras similar to those used on the LRO [31] could

25


allow for extremely high quality images to be returned from Mars. As was done by the

Reconnaissance Satellites, stereo images will be returned and will allow the calculation of

topological data [41].

6.5.3 Earth Re-entry

With this being the first time a manned mission has attempted re-entry to Earth at such a

velocity, data on temperature and pressure distributions experienced by the capsule during

re-entry could provide valuable information that will aid future re-entry missions. This data

can be collected in a similar way as for the Mars Science Laboratory, by using atmospheric

data sensors such as MELDI and MEADS integrated into the re-entry capsule heat shield

[42].

7 Launch Vehicles

7.1 Launcher Comparison

Due to the substantial payload mass, the evaluation of launch vehicles was confined to only

the heaviest launchers. As Inspiration Mars has been put forward as an American mission,

all Russian and Chinese launchers were excluded as possibilities. Table 8 provides a list

of the considered launch vehicles, as well as their payloads to LEO and launch cost. The

Ariane 5, Delta IV and Atlas V all have high per-mass costs to deliver a payload into orbit,

when compared with either the Falcon Heavy or the Falcon 9 launchers. While the Space

Launcher System (SLS) Block 2 would allow for the largest payload to be lifted to LEO, its

price per kilogram in orbit remains higher than either of the Falcon launchers. Additionally,

the SLS is only expected to perform its first test flights in 2017 and is therefore unlikely to

be ready for use in time by 2018. In comparison, the Falcon Heavy launcher should perform

its first flight later in 2014 [43]. The Falcon Heavy launcher provides the most cost-e�cient

means of access to LEO, as well as a high payload capacity of 53 metric tonnes. It can also

be fitted with SpaceX’s own Dragon capsule, which is to be modified and demonstrated for

human occupation.

Launcher Ariane 5 Ariane 5 ME Atlas V 551 Altas V 541 Delta IV-H SLS Block II Falcon 9 v1.1 Falcon Heavy

Payload to LEO (t) 21 25.2 18.5 17.1 22.95 105 13.15 53

Cost of launch (M$) 120 150 190 223 380 750 56.5 135

Cost per kg (k$/kg) 5.714 5.952 10.270 13.041 16.558 7.143 4.297 2.547

Table 8: Total payload to LEO and associated costs for various launch vehicles.

26


7.2 Maximum Number of Launches and LEO Launch Mass

Due to schedule, planning and in-orbit docking procedures, as well as cost considerations,

the number of launches available for the mission was limited to an absolute maximum of four

launches. It was decided that the final launch sequence would consist of four Falcon Heavy

launches. This would allow a maximum payload of 172 metric tonnes in a 200km altitude

circular orbit. By extrapolating the loss of payload for higher orbits using SpaceXs Falcon

9 user manual [44], Table 9 was produced showing the total payload to 700 km per Falcon

Heavy would be reduced from 53 metric tonnes to 47.2 tonnes. This corresponds to a total

mass in orbit of 188.8 tonnes using four Falcon Heavy launches. This total would have to

include the dry mass of the craft, the propellant needed in order to inject it into its orbit,

as well as any additional fuel tanks.

Circular Orbital Altitude

200 km 300 km 400 km 500 km 600 km 700 km 800 km 900 km 1000 km

Falcon 9, v1.0 10450 10250 9950 9750 9550 9350 9250 9150 8950

Falcon 9, v1.1 Payload (kg) 13000 12751 12378 12129 11880 11632 11507 11383 11134

Falcon Heavy 53000 51986 50464 49450 48435 47421 46914 46407 45392

Table 9: Table shows the payload mass limit for a range of altitudes for the Falcon 9 and Heavylaunch vehicles.

7.3 In-Orbit Propulsion

7.3.1 Delta-V Budget

Earth Escape The velocities at infinity required for the desired trajectory to Mars were

calculated and detailed in the Inspiration Mars feasibility study [45], and can be found in

Table 10.

Departure V Infinity (km/s) Arrival V Infinity (km/s)

Mars-bound 6.232 5.417

Earth-bound 5.417 8.837

Table 10: Velocity at infinity of Earth’s departure and the return trip’s arrival.

These velocities correspond to an orbit optimised for free return. Using the velocity at

infinity upon departure from Earth, the following equation was used in order to calculate

the velocity required at the perigee of the escape hyperbola.

�v = vp

� vc

=

vuut2

v212

+µE

rp

!

�sµE

rp

where µE

is the gravitational parameter of Earth, vp

is the velocity at the perigee of the

hyperbolic trajectory, vc

is the velocity in the circular parking orbit and rp

is the radius of

27


the circular parking orbit.

As can be seen from the equation, the necessary velocity is dependent on the radius of

the parking orbit. This design implements a circular parking orbit at an altitude of 800

km in order to avoid manoeuvre complications with LEO orbiting spacecraft (ISS) while

ensuring the altitude is great enough to minimise atmospheric drag. This leads to a required

delta-v of 4800 m/s.

Gravity Loss Gravity losses are also dependent on the initial acceleration provided by the

engine and have been assumed to be approximately 200 m/s as a contingency, hence bringing

the total delta-v to provided 5000 m/s. The validity of this contingency is discussed and

approved below, using the selected orbital engine.

7.3.2 In-Space Propulsion System

Merlin 1D Engine The simplest launch option involves using the Falcon upper stage

engine, currently the Merlin 1D Kerosene/LOx engine. This would mean increasing the size

of the propellant tanks carried by the upper stage, or carrying separate tanks as payload.

This would be advantageous, as the only cost would be that of the new propellant tanks,

but would not represent a large investment in terms of development costs, nor the purchase

of other expensive engines. However, with an Isp quoted at 342s [43], the Merlin 1D would

require a large quantity of fuel in order to reach the required delta-v, as shown in Table 11

below. Other chemical propulsion systems were also considered. While the Merlin 1D engine

uses a combination of kerosene and liquid oxygen as propellants (responsible for the low Isp),

LOx/LH2 engines, like the RL-10B-2 [46] engine employed by the Delta IV launchers upper

stage, as well as the Space Shuttle Main Engine generally produce higher Isps. However, they

come with the disadvantage of requiring heavier and more complex systems for propellant

storage. Furthermore, these engines are more complex and more costly than the Merlin 1D.

It is clear that, in already using SpaceX’s Falcon Heavy launch vehicle, there is the distinct

cost-benefit and simplicity in using SpaceX’s Merlin 1D engine.

Raptor Engine SpaceX is currently developing a new orbital engine - the Raptor engine,

fuelled with a mixture of methane and liquid oxygen. These new engines should equip

SpaceXs vehicles in the future, but as of now remain in development. These new engines

are included in Table 11. As there is only limited data available for methane-fuelled engines,

two values of specific impulse were employed in this table: the pessimistic Isp value of 350s

28


represents a small increase from the Merlin 1D engine, whereas the optimistic value of 380s

is reportedly the target that SpaceX aims to achieve.

Engine Isp (s)Burn 1 Burn 2

Init. Mass (t) Delta-V (m/s) F. Mass (t) Init. Mass (t) Delta-V (m/s) F. mass (t)

Merlin 1D 342 196.7 825 154 147.5 1152 105

Raptor 350 196.7 862 153 147.5 1206 104

Raptor 380 196.7 936 153 147.5 1309 104

Engine Isp (s)Burn 3 Burn 4 Total S/C

Init. Mass (t) Delta-V (m/s) F. Mass (t) Init. Mass (t) Delta-V (m/s) F. Mass (t) Dry Mass (t)

Merlin 1D 342 98.3 1340 66 59.7 1682 36 31.8

Raptor (p.) 350 98.3 1413 65 59.7 1519 38 36.6

Raptor (o.) 380 98.3 1534 65 59.7 1221 43 41.7

Table 11: Comparison between the Merlin 1D values and the Raptor estimated values.

As the future of the Raptor is yet very uncertain, it was decided that the Merlin 1D engine

would be used in this design. If the Raptor engine were to become available by the onset of

the mission, the potential mass gains shown in Table 11 could be applied as an enhancement

to this design with the additional fuel savings potential being used to reduce reentry velocity

or to reduce the number of total required launches.

8 Reentry

The purpose of the thermal protection system (TPS) is to slow down the spacecraft, thus

providing a human tolerant temperature inside the spacecraft and avoiding high acceleration

loads. This is achieved by using drag to counteract excess speed. However, the interaction

between the air and spacecraft produces a large amount of heat, which the material must be

capable of withstanding. The main requirements driving the design of the heat shield mate-

rial are its maximum load factor, the maximum heat flux that the material can withstand,

the total mass of the TPS and re-usability if aerocapture is the method used.

8.1 Reentry Vehicle Configurations

Develop a new vehicle is very expensive and time consuming. Due to the available time, a

available spacecraft will be selected. It will focus on the TPS subsystem on each vehicle.

NASA proposal NASA’s Orion MPCV (Multi-Purpose Crew Vehicle) is being designed

for travelling beyond low Earth orbit (LEO). However, the capsule was not considered for

this mission due to scheduling uncertainties- it is unlikely that the capsule would be ready

for use by 2018. The extra mass of the capsule, compared to that of the Dragon, is another

reason why the Dragon is preferred.

29


SpaceX Proposal SpaceX’s Dragon capsule is designed to deliver both cargo and eventu-

ally humans to destinations in LEO. It is the only spacecraft currently in production that is

capable of returning significant amounts of cargo to Earth. Under an agreement with NASA,

SpaceX is currently adapting the capsule to enable it to transport humans [47].

8.2 Reentry Trajectory

The main problem encountered when designing a spacecraft to re-enter from Mars is the

high approach velocity on returning to Earth, which is around 14km/s. To date, no manned

spacecraft has achieved re-entry at such a velocity- the highest that has been achieved is

by the Apollo missions, which at 11km/s [48]. The highest successful re-entry velocity for

an unmanned spacecraft was experienced by the Stardust probe, which returned at 12km/s.

The maximum acceleration load that is permissible for astronauts returning from long du-

ration missions to experience has been limited by NASA to 5gs.

The re-entry path angle can be altered to limit these loads- as this angle increases, the

acceleration also increases. In order to be captured into Earth orbit, the speed must be re-

duced to around 11km/s. The lifting body configuration of the Apollo or the Soyuz capsules

is also used to to reduce acceleration in the case of a purely ballistic re- entry. If the path

angle is too low, the spacecraft is unable to re-enter and returns to space. This e↵ect can be

used to reduce the acceleration load and heat flux experienced, although it comes with the

cost of increasing the mission duration by a few days. After one missed re-entry attempt,

the capsule can reduce its speed to a value low enough to allow for re-entry. In this case, the

heat shield must be designed so that it can be used for multiple entries into the atmosphere.

The aerocapture method also imposes restrictions on other elements of the mission design-

since the service module and inflatable habitat must be jettisoned before the first re-entry,

the subsystems on board the Dragon capsule must be capable of supporting the astronauts

throughout the duration of the re-entry period. The electrical power subsystem is of partic-

ular importance- see section 7.1.2 for a sizing analysis of the hydrogen fuel cells.

The model used for the Apollo missions [48] has been implemented using the properties

of the Dragon capsule. Fig. 16 shows the trajectory for the first re-entry from Mars. Once

the altitude is high enough the new orbit is calculated. A second re-entry is used to decrease

the speed to 8.1 km/s resulting in a 1.5 hour orbit. The third and final orbit eliminates the

remaining speed in order to return the astronauts safely to Earth.

30


0 50 100 150 200 250 3000

0.5

1

1.5

2

2.5

3

3.5

4

4.5

5

time, s

Acce

lera

tion,

g

Reentry Acceleration

0 50 100 150 200 2500

0.5

1

1.5

2

2.5

3

x 105

time, s

Altit

ude,

m

Reentry altitude

0 50 100 150 200 250

2000

4000

6000

8000

10000

12000

14000

time, s

Velo

city

, m/s

Reentry Speed

Figure 16: First re-entry. The aerocapture reduces the speed from the original 14.2 km/ to a suitable10.95 km/s that has 3.5 days as orbit period. The initial fligh path is 6.1 deg

0 100 200 300 400 500 600 700

0.5

1

1.5

2

2.5

time, s

Acce

lera

tion,

g

Reentry acceleration

0 100 200 300 400 500 600 700

4

5

6

7

8

9

10

11

12x 104

time, s

Altit

ude,

mReentry altitude

0 100 200 300 400 500 600

1000

2000

3000

4000

5000

6000

7000

8000

time, s

Velo

city

, m/s

Reentry velocity

Figure 17: Simulations of the last re-entry.

8.3 Heat Flux

This is the key parameter in the TPS selection. Fig. 19 shows the maximum heat flux as

a function of the re-entry speed [49]. In this case, the heat flux is expected to be around

3300 W/cm2. This graph also shows results for two di↵erent mass loadings. Fig. 18 shows

this heat flux against time for each trajectory. The peak heat flux is experienced at the

beginning of the trajectory.

Figure 18: Heat flux throughout the reentry.For the expected reentry at 14 km/s the max-imum heat flux is around 3300 W/cm2

Figure 19: Heat Flux for di↵erent configura-tions

31


8.4 Heat Load

The heat load defines the required thickness of the TPS. For example, the Dragon capsule

has 3 inches of PICA-X and the Apollo mission had 7 inches of Avco 5026-39G [50] With

lifting bodies the heat flux peak reduces but the heat load increases. For the Mars re-entry

this parameter is not the key parameter to select in the TPS selection.

Figure 20: Heat load for several configuration, from [49].

8.5 TPS material

Considering all possible materials on the market, there are few possibilities to use for the

TPS:

AVRCOAT Used in the Apollo missions [50]. It was the standard TPS material until the

development of PICA.

PICA Developed by NASA. It has been used in the Stardust probe and the MSL.

PICA-X Recently, Space-X has developed a new material called PICA-X based on the

original PICA. It is ten times more cost e↵ect and more reusable [51]. They have tested it

and it has been shown to withstand temperatures of up to 1850 degrees Celsius. However, it

cannot withstand the 3500 W/cm2 that are expected [49]. According to the CEO of Space-

X, Elon Musk, the spacecraft can re-enter from Mars [52] but no further references has been

found.

High Density Carbonic Phenolic A material developed by NASA. The document [49]

shows that it is valid for the conditions of the re-entry from Mars. It has been used in the

Phobos Grunt [53]. This mission was designed to travel to one of the Mars moons, Phobos,

and return with samples. The spacecraft had problems in low Earth orbit and it re-entered

32


without leaving the Earth. The material has also been proposed for the Mars Sample Return

mission. [54]

Hence, only the High Density Carbonic Phenolic can perform a re-entry at 14 km/s, so

this material will be the one used for the TPS. For testing the material, new facilities will be

necessary. This material will be adapted to be used on the Dragon body instead of PICA-X.

9 Electrical Power Systems

This design only considers traditional electricity generation methods and storage systems -

solar arrays and batteries. Bigelow’s Sundancer module and SpaceX’s Dragon capsule are

expected to be equipped with solar arrays large enough to sustain full functionality of their

original equipment systems at a distance of 1 AU from the Sun, but not at the 1.39 AU

distance of Mars.

General Assumptions The Dragon will be equipped with solar arrays and batteries as

standard. A regenerative hydrogen fuel cell will be added inside the crew compartment. The

Sundancer will be modified considerably to accommodate the radiation shielding systems.

Both of these are expected to have significant additional power requirements. Additional

solar panels and batteries will therefore be required.

9.1 Dragon

9.1.1 Power Generation

During the interplanetary phase of the mission the Dragon will not be used as a primary

habitat and therefore will require minimal power. Even at the distance of Mars, this will

be provided by the Dragon’s own solar arrays. SpaceX states the mean power provided by

the solar arrays in LEO is 2kW [55]. At the Mars flyby this will be reduced but will remain

greater than 1kW [47].

9.1.2 Energy Storage

The batteries of the Dragon are capable of powering the capsule through LEO eclipse and

normal re-entry operations. The two-pass re-entry method used will require additional energy

storage. Assuming a continuous power drain of 2 kW for a maximum of 3 days (72 hours),

144 kWh of energy will be needed. If Li-ion batteries with an energy density of 100 Wh/kg

and a charging e�ciency of 0.96 were used, the battery mass would be of 1500kg. Due to

33


this large mass, these batteries will be replaced by hydrogen fuel cells. These have extensive

space heritage from the Apollo and Space Shuttle programs. The typical energy e�ciency for

a proton exchange membrane fuel cell is 50%. This leads to approximately 20kg of reactants

(hydrogen and oxygen) for a total mass of 90kg including the fuel cell and tankage.

Re-entry electrical energy required 144 kWh

Chemical energy required 259.2 MJ

Mass of hydrogen [56] 2.16 kg

Mass of hydrogen tanks [57] 40 kh

Mass of Oxygen 17.28 kg

Mass of Oxygen Tankage 24.1 kg [58]

Fuel cell dry mass at (275 W/kg) 7.27 kg [33]

Table 12: Fuel cell requirements

9.2 Sundancer

9.2.1 Power Consumption

During the interplanetary phase the major power consumption will come from life support,

waste management, communications and crew activities/science systems. The total load of

these systems will be 8188W, as detailed below.

9.2.2 Life Support

Life support systems are nominally provided by the Sundancer module, but their power

consumption must be estimated in order to modify the spacecraft’s solar arrays. Assuming a

four bed molecular sieve is in use on board the ISS for CO2 removal, and a Sabatier process

O2 generation system were assumed, air management would total 600W of power. For water

management 140W were assumed based on filtration and distillation needs. This led to a

total system power requirement of 740W.

9.2.3 Waste Management

The peristaltic action of the waste management system is provided by pumping water. The

power required to do this was calculated by Bernoullis equation.

P =M ⇥�P

✏

where P is the power; M is the mass flow rate; �P is the drop in pressure; ✏ is the e�ciency

of the whole system.

A typical centrifugal pump [59] is used and a mass flow rate of 50 g/s is assumed. The

34


pressure drop is considered to be 5 bar. This would give a power requirement of approxi-

mately 333 W.

9.2.4 Communications

In addition to the ComSat’s laser communications system, a conventional radio system will

provide redundancy and primary communications on the return to Earth. An X/Ka-band

system was sized using the link budget equation to estimate the power required. The X-

band link will provide telemetry, data transfer and some communication capability. The

Ka-band link will provide video and image transfers on the return travel. A 2.2 m antenna

and power of 550 W for the X-band was determined to allow a data rate of 2Mbps. The

Ka-band required the same antenna and 300 W of power for a 7.5Mbps data rate. Both

systems allowed for a 3 dB margin. All assumptions made in this calculation were the same

as for the ComSat. Maximum power for simultaneous communications is therefore 850 W.

This also assumes that the radio links to the sub satellite will only require a small amount

of power.

9.2.5 Crew Activities / Science

The power required for the crew activities and equipment and the mission science is shown

in Table 13.

Activity Required Power (W)

Crew Electronics and OBDH 3000

Human Research 1000

Plant growth experiments 2 x 200 W 400

Experiment Freezer 500

Total 4900

Table 13: Power requirements for Crew Activity

9.3 Power Generation

Two limiting cases were identified on the Sundancer’s solar arrays’ power production:

• Case A: at Mars flyby, the solar flux will be at its lowest.

• Case B: when returning to Earth, radiation degradation of the solar cells will reduce

the power generated.

The lowest power raised from these two cases is used to size the array. Standard assump-

tions of inherent degradation factor of 0.77, electrical conversion e�ciency 0.9 and pointing

accuracy of 5o were made.

35


9.3.1 Case A

Mars will be close to perihelion during the flyby, at 1.397 AU from the sun. The solar flux

at the point of the Mars Flyby is calculated to be approximately 700 W/m2 [60]. Therefore

the electrical power output will be 131 W/m2, accounting for radiation degradation.

9.3.2 Case B

Assuming GaAs based solar cells, the power conversion e�ciency will be 28 % at beginning

of life. This is degraded by radiation exposure by a factor of about 2.75 %/yr in LEO [33].

As the spacecraft is travelling through a harsher radiation environments than LEO, this

factor was increased to 10 %/yr. Therefore, for a 501 day mission the e�ciency is reduced

to 26.1%, leading to an electrical power output of 245 W/m2 at the end of life.

Case A is therefore the limiting case for the power generation system. The arrays are sized

to this case with an area of 66m2, generating 8360W. This can be divided into two 11m x

3 m arrays deployed. These can be stored along the sides of the Sundancer module during

launch.

9.4 Energy Storage

As only 1 eclipse will be encountered outside LEO this does not drive the battery require-

ments of this module. Peak loads such as pumping water between radiation shield bubbles

may require more power than instantaneously generated. It is di�cult to estimate this

without further as the power depends on the pressure drop which would have to be found

by prototyping or specialist analytical methods. In this case the hydrogen fuel cell on the

Dragon should be employed. No additional batteries will be used on the Sundancer other

than those installed as standard for LEO.

Energy Considerations during Reentry During the aerobraking and re-entry manoeu-

vre the Sundancer and Dragon’s trunk will be jettisoned, hence eliminating the primary

source of power. The Dragon capsule will therefore rely on internal power between the

initial separation from the other components until the crew is recovered. As only one aer-

obraking orbit will be performed before the final re-entry, this length of time will be of 3

days.

36


10 Cost Evaluation

Most of the components used in this design are either o↵-the shelf components or modifica-

tions of existing technologies. These modifications will need some development investment,

but the cost of doing so will still be considerably lower than that of developing a new tech-

nology.

10.1 First approach

The paper ‘Manned Mars Mission Cost Estimate’ [61] was used as a reference to aid in

estimating the costs of the mission, although the values given were not directly applicable

since the estimate was based on values and technology from 1985. As the mission described

in the paper included a landing on Mars while this mission does not, the values produced

had to be modified in order to account for this. An adjusted comparison yielded an initial

estimated average mission cost of $1.91B. This is below the stated maximum of $2B. However,

the maximum estimate obtained is $4B, which largely exceeds the requirements. For this

reason, further studies were conducted.

10.2 Detailed Cost Study - QuickCost 5.0 Model

One of the authors of the previously mentioned paper, Dr Hamaker, was contacted for further

discussion about cost estimation for such a mission. He provided his QuickCost 5.0 model to

assist in this task. The model calculates Cost Estimate Relationships (CERs) for a number

of mission types based on a large database of information on previous missions. The models

that were used to perform the estimation for the mission were the Module & Transfer Vehicle

Model for the main spacecraft and the Satellites Model for the ComSat.

10.2.1 Main Spacecraft

The parameters used by the QuickCost 5.0 model included the spacecraft dry mass, design

life, whether or not the mission was manned, whether the vehicle was to be used as a transfer

vehicle, whether it required a propulsion system and its Initial Operating Capability. The

percentage of the spacecraft that consists of new designs is also an input parameter - the

nominal value for this is taken to be 50%. However, the dependence of cost on this parameter

was also studied. With these parameters it was possible to obtain the median estimated

cost for DDT&E (Design, Development, Test and Evaluation) and TFU (Theoretical Flight

Unit). The results obtained are shown in the figure below. The confidence level refers to the

confidence in meeting the expected budget.

37


Figure 21: Main spacecraft cost of the mission as the function of the desired confidence level inpercentile

The highlighted region in the graph was seen as the most relevant to this mission. With

more than 50% new components, and with a confidence level of 75% a total cost estimate of

$2.18B was obtained. If the development costs were reduced to a smaller percentage (30%)

of new components, this cost could be reduced to $1.63B, with a confidence level of 75%.

The following table shows the QuickCost inputs and results for the nominal case.

Total Dry Mass (kg) 31,800

New Design (Percentile) 50%

Design Life (months) 18

Human Rated? 1

Transfer Vehicle Capability (i.e. Propulsion) 1

Initial Operating Capability (IOC) 2018

Non-Market Economy (i.e. Russian) 0.0

Year Dollar of Output (20XX) 2014

Median Estimated Cost $1,200.3

Median Estimated Schedule (months) 54.7

Desired Confidence Level (percentile) 75%

Median Cost + Reserve to Achieve Desired Confi-

dence Level

$1,474.4

Median Schedule + Reserve to Achieve Desired Con-

fidence Level

67.2

Calculated MCPLX for PRICE Users 10.82

Calculated Peak Size of the Government Project Of-

fice

143.1

Ground System Cost $88.5

Launch Services Cost $474.4

Annual MO&DA Cost $73.7

Life Cycle Cost (Over Design Life) $2,184.6

Table 14: QuickCost analysis for main spacecraft. All cost values are given in Millions of Dollars.

10.2.2 Communications Satellite

A similar analysis has been performed for the laser communications satellite, using the Satel-

lite Model from QuickCost 5.0. Parameters include the total instrument complexity, assumed

to be 65% doe to high AOCS pointing and stability requirements, and the percentage of the

38


bus and instruments which consist of new designs, assumed to be 40% for both as existing

designs are being modified. With these parameters, the total life cycle cost of the ComSat,

including development, production, ground system, launch and operations is estimated to

be around $303M.

10.3 QuickCost Results

From the QuickCost model, the total development and production costs of the mission,

including the ComSat, through phases B/C/D will be around $1.45B, and the total life cycle

cost is $2.49B with the assumptions stated above.

10.4 Cost Scheduling

An analysis was performed to determine how costs will be spread over the mission lifetime.

A Beta function showed that 60% of the total funds were spent in the first half of the mission

life cycle. Annual MO&DA Costs were considered at an additional 5%/year on top of these.

This was then adjusted for inflation using the NASA New Start Inflation Index [62].

Figure 22: Time Cost Distribution

10.5 Validation

To validate these cost estimates di↵erent models were used and their results were compared.

10.5.1 Communications satellite

A second cost estimation for the ComSat was performed using the USCM8 method, outlined

in SMAD [33] and adjusted for inflation. Due to schedule constraints and the fact that

39


most of the components have considerable flight heritage, it was assumed that a protoflight

method will be used. For this reason, the non-recurring cost CERs were used to calculate

subsystem costs and a 30% margin was included to cover refurbishment costs. The results

of this analysis are shown below.

Table 15: USCMD8 Costing of the ComSat

Element Cost (FY14$K) Standard Error Min Cost (FY14$K) Max cost (FY14$K)

Structure/Thermal Control 18317.40 22% 14287.57 22347.23ADCS 37536.36 44% 21020.36 54052.36EPS 3857.52 41% 2275.93 5439.10Insertion Engine 5860.35 22% 4571.07 7149.63Orbit Acquisition Engine 3461.72 22% 2700.14 4223.30TT&C 31768.58 n/a 31768.58 31768.58Laser 21453.66Electra 3218.05Narrow Angle Cameras 11197.74Wide Angle Camera 2356.37Payload Total 38225.82 39% 23317.75 53133.89Integration,

26074.77 42% 15123.37 37026.18Assembly and TestingProgram Level 158306.96 37710.81 28% 27151.78Ground Equipment 31258.24 37% 19692.69 42823.79Reaction Control System 18627 2526.51 35% 1642.23

Total Cost 307577.49Minimum Cost 212616.92Maximum Cost 402538.06

As can be seen, the estimate given by this model ($288M) is similar to that given by the

QuickCost model ($303M). Note that since the USCM8 model only provides CERs for com-

munications payloads, the NASA Instrument Cost Model, also given in [33] was used to

estimate the costs of the Narrow Angle Cameras and Wide Angle Camera. The costs for the

Electra package and the laser subsystem were estimated based on the relative masses and

costs of similar missions carrying these payloads.

10.6 Main spacecraft - NAFCOM Weight-Based Model

The second estimation of the main spacecraft cost used the NAFCOM method which ex-

trapolates from other space projects, with mass as a scaling parameter. In 2011 a NAFCOM

study was carried out on the Falcon 9 launch vehicle with a total cost estimated between $4B

(traditional approach) and $1.7B (commercial approach). SpaceX had only spent $390 M,

highlighting the radical cost savings of their methodology. If our mission can be more cost

e�cient than traditional aerospace projects the NAFCOMmodel may be overly conservative.

NAFCOM looks at the di↵erence between manned and unmanned missions and to iden-

tifies the main characteristics which can be compared to the mission at hand. The manned

mission data was taken mainly from 60 NASA missions (Apollo, Gemini, Space Shuttle Or-

biter, Skylab and SpaceLab).

A cost approximation can then be made for this mission using the complexity factor, ↵,

40


which is an adjustment to the CER to compensate for the projects unique features which are

not accounted for in the historical data. It also takes into account the fact that this mission

is using a higher percentage of COTS components when compared to those missions.

By assuming a range of ↵ = 0.4 to ↵ = 0.5, and a total spacecraft mass of 30 metric

tonnes, a development cost ranging between $2031M and $2539M was obtained, whereas the

Flight Unit production cost went from $242M to $303M. Thefore, this model predicted a

total mission cost between $2273M and $2842M.

By varying the NAFCOM approximation parameter values for each individual subsystem,

a more accurate cost of the overall mission can be obtained. The following table shows the

results obtained when applying this method, with a confidence estimation level of 90%.

Using historical data points obtained from previous analogous space missions, the ”first

pound cost” approach from NAFCOM can be used. The same model was used for most sub-

systems but the fact that many COTS components were used was also taken into account.

The ensuing risk assessment analysis was performed to provide a cost estimate for di↵erent

confidence levels. The results can be seen in the table 16.

These results provide validation of the QuickCost model, as the total mission cost is esti-

mated as $2.49B. As mentioned above, this cost can be reduced by lowering the development

investment in the development of new technologies. The radiation shielding and waste man-

agement subsystems are the ones that require the highest amount of development, so reducing

the requirement on the level of protection would significantly lower the cost

Model Spacecraft Min Cost $B Nominal Cost $B Max Cost $B

QuickCost ComSat 0.228 0.303 0.490

USCM8 ComSat 0.213 0.308 0.402

QuickCost Main 1.87 2.19 3.03

NAFCOM (Spacecraft Level) Main 2.27 3.06 3.84

NAFCOM (Subsystem Level) Main 2.05 2.28 2.51

Table 16: QuickCost Minimum estimated at 50% confidence, Nominal at 75%, Maximum at 99%

41


Figure 23: NAFCOM Subsystem Analysis

11 Mission Phases and Scheduling

This section provides a general overview of the various mission phases and an outline of a

potential schedule.

11.1 Development Schedule

The concept presented has been designed so that most of the elements are COTS (Commer-

cial O↵ The Shelf) so as to reduce development costs. However, some will need time to be

tested and developed, as it is discussed below.

42


Waste management: While the reverse osmosis membrane material will be ready on

time, the design and verification of the whole pump system will need to be done. As this is

a versatile water treatment system, additional funds may be available to accelerate develop-

ment.

Bubble radiation shield: This innovative system will require significant testing on Earth

and potential verification aboard the ISS. Prototyping can begin very quickly with low cost

materials for rapid iteration.

Laser communications: The concept of laser communication from space has been demon-

strated on board the LADEE spacecraft. Moreover, the OPALS laser communications pay-

load is scheduled for launch to the ISS in March 2014. [23]

Modification of the Dragon and Sundancer: It is anticipated that the Dragon and

the Bigelow Expendable Activity Module (BEAM) will be flown with crew before 2018, but

additional development to take these spacecraft beyond LEO must be considered. Even if

both Bigelow Aerospace and SpaceX wish to make these alterations, the timescale for such

development is unknown.

The QuickCost 5.0 model 10 also provides a schedule estimate based on parameters from

previous missions, giving a development and production time of 54 months. While this is on

pare with the available time, it does not take into account either the accelerating factor of a

deadline or the high proportion of COTS hardware that this mission will have.

11.2 Launch and Docking Schedule

11.2.1 Launch and mass distribution

As previously stated 7, the total payload to be distributed amongst the four Falcon Heavy

launchers will be of 188.8 tonnes. It has been calculated that the use of Merlin 1D engines

for the injection into the desired trajectory would require a combined fuel and tank mass

of 157 tonnes. The total dry mass of 31.8 tonnes is to be distributed between the Dragon

capsule, which is specified as having a dry mass of 4.2 tonnes [47] and carries a cargo of 6

tonnes, and the Sundancer, with a mass of 8.6 tonnes [1], carrying 4.4 tonnes of cargo.

The two initial Falcon Heavy launches will be dedicated entirely to propellant and tanks,

with a total of 47.2 tonnes each. These will be followed by the Sundancer and its cargo,

43


as well as a total of 25.6 tonnes of additional propellant. The final Falcon Heavy will carry

the Dragon capsule, crew and additional cargo, as well as the remaining 37 tonnes of fuel

required.

11.2.2 Launch sequence

Within the next few years, SpaceX hopes to increase its rate of production and launch to

ten Falcon Heavy launchers and ten Falcon 9 launchers per year [43]. This corresponds to

a minimum delay of about five weeks between each launch. However, a contingency should

be applied to this, and therefore two months (about eight weeks) will be taken as the ap-

proximate delay between each launch. An additional constraint on the schedule comes from

the launch site. By assuming a time delay of two months this would cease being an issue, as

demonstrated by SpaceXs two Falcon 9 launches from Cape Canaveral in less than a month

(13 December 2013 and 6 January 2014).

This two-month delay will also allow for time to perform the successive in-orbit assem-

blies as well as the ensuring verifications. The Merlin 1D engine is fuelled by kerosene and

liquid oxygen so propellant boil-o↵ is not as much of an issue as it would be for a LOx/LH2

engine. Therefore, the propellant payloads can be safely stored in orbit for longer periods of

time. For this reason, the two Falcon Heavy launchers carrying propellant will be launched

first, followed by the combination of the Sundancer module with the added cargo and the

remaining propellant.

In order to limit the time the astronauts spend in space, which in turn will reduce the

dose of radiation inflicted, the food consumed and the physiological e↵ects of habitation in

space, the manned Dragon will be launched in the very last instance.

11.2.3 Docking and in-orbit assembly

All rendezvous manoeuvres and docking will be performed autonomously, and will be mon-

itored from the ground. Upon completion of the second docking manoeuvre, once the Sun-

dancer is attached to the two propulsion modules, the module will be automatically ex-

panded. A remote verification of the functionalities of the system will confirm the correct

deployment of the module.

Following the docking of the Dragon capsule to the Sundancer module, the crew will en-

ter the expandable module and assemble the bubbles. This operation should only be a few

44


hours. The assembly of the waste management system will ensue. As this operation may

take up to several days, it does not need to be complete before departure from Earth orbit.

Following a final verification, the green light will be given for the injection into Mars orbit,

with the propulsion burn being accomplished on the 5th of January 2018. The full launch,

docking and verification sequence is shown in Fig. 25.

Figure 24: Timeline of launches (orange), dockings (blue) and equipment verifications (light green)before injection into Mars trajectory.

11.3 Interplanetary Events

Following the docking manoeuvres and subsequent verifications, the authorisation will be

given to proceed, and the injection into Mars trajectory will be performed on the 5th of

January 2018. During this manoeuvre, each Falcon Heavy upper stage will consume the

fuel of its additional fuel tank. Then, each stage will be jettisoned, before the ignition of the

following stage. Details of the relevant delta v achieved and of the remaining mass can be

found in the Launch and Injection into Mars trajectory section, 7. The entire burn sequence

will last less than 30 minutes, and will be followed by a final verification.

On the 6th of January, once the spacecraft trajectory has been confirmed and verified,

the ComSat will be deployed through a spring-loaded mechanism. This will ensure that the

satellite and the main craft travel as two separate entities, although the di↵erence in velocity

between the two will be insignificant and will hence not a↵ect their respective trajectories.

A relative velocity of 25mm per hour, 6.94 ⇥ 10�6m/s, would prove optimal as it would be

negligible in orbital calculations, whilst still providing a su�cient separation between the

two spacecraft within a few days.

On the 9th of January, assuming a relative velocity of 25mm per hour, the separation

45


between both spacecraft will reach 2m. At this point, communication between the two

spacecraft will begin, and the ComSat will be used as a relay satellite for all primary com-

munications between the mission and Earth.

The fly-by of Mars will occur on the 20th of August 2018. By this time, assuming a

relative velocity of 25mm per hour, the distance between both spacecraft will be a safe

135.6m. Upon reaching the periapsis of the fly-by trajectory, the ComSat will perform a

burn in order to reduce its velocity and adopt an elliptical orbit about Mars, with a semi-

major axis of 28470 km and an eccentricity of 0.994. Subsequent orbital manoeuvres will

finally place the ComSat into a circular orbit about mars, at an altitude of 500km (3890km

radius). The combined Dragon and Sundancer spacecraft will remain on its unaltered course,

returning towards Earth.

The Earth approach will occur on the 21st of May 2019. The Sundancer module and the

Dragons trunk will be separated and jettisoned, as the crew will take place in the Dragon

capsule alongside food supplies for five days. A first pass through the Earths atmosphere,

as described in the Reentry section, will provide the aerocapture: the capsule will be in an

elliptical orbit about the Earth. An orbit around this ellipse will take three days. Follow-

ing this, on the 24th of May 2019, the final re-entry into the Earths atmosphere will occur.

Therefore, the interplanetary travel will last between the injection into Mars trajectory on

the 5th of January 2018, and the final Earth re-entry on the 24th of May 2019, or 504 days.

When including the scheduled launch date of the astronauts on the 31st of December 2017,

the entire manned section of the mission will last 509 days.The sequence of events during

the interplanetary trajectory can be seen in Fig. 26.

Figure 25: Interplanetary trajectory and events schedule

46


11.4 Risk mitigation and 2021 launch window

Many of the mission components, including the launch segment, the Dragon capsule and

the Sundancer module, are to be developed and produced by private companies. This not

only means that their development expenses are accounted for by these companies, but also

that their development time may prove shorter than that expected for an agency such as

NASA. Additionally, the Falcon Heavy will perform its first flight late in 2014, the unmanned

Dragon capsule is already in operation and progress is being made towards a manned version,

whereas most of the technologies used by the Sundancer will be demonstrated by BEAM

starting in 2015 [3].

Similarly, the satellite communication systems to be used on the ComSat are being widely

researched, through such missions as LADEE or OPALS, and therefore should then be a

proven technology by 2018. Table 1 below gives a risk mitigation matrix, in which delays

in each components development are met with their likely e↵ect on the mission. As this

table shows, the radiation shield, the Sundancer module, the Dragon capsule and the Falcon

Heavy launcher are critical to this design, as their unavailability would lead to abandon the

2018 launch opportunity in favour of that in 2021. However, it is expected that each of these

elements will be ready on time.

Conversely, this mission design will not require large modifications if the 2021 launch win-

dow were to be adopted. Indeed, despite the longer flight time (597 days from launch of the

crewed Falcon Heavy until touchdown, considering the same launch and docking sequence as

for the 2018 mission). Food supplies for the additional 88 days will have to be accounted for,

but despite the radiation levels being higher due to the solar minimum experienced in 2018

no longer being a benefit, the radiation shield designed for the original scheduled mission

will protect the crew su�ciently to ensure their health and well-being.

Mission component Marginal delay (not ready by 2018) Severe delay (not ready by 2021)

Waste management Replace with Isonomer membrane bags, de-

spite loss in functionality

Replace with Isonomer membrane bags, de-

spite loss in functionality

Laser communication Replace with high gain antenna, despite loss

of data rate

Replace with high gain antenna, despite loss

of data rate

Bubble radiation shield Postpone mission to 2021 window Attempt to replace with polyethylene layers

Sundancer module Postpone mission to 2021 window Attempt to use Cygnus- or ATV-type of

module

Manned Dragon capsule Postpone mission to 2021 window Attempt to use Orion capsule, despite cost

considerations

Falcon Heavy launcher Postpone mission to 2021 window Attempt to use SLS launcher, despite cost

considerations

Table 17: Risk and consequence matrix for the mission components which require development.Note that the four last elements are critical to the mission, as a delay will imply opting in favourof the 2021 launch window instead.

47


12 Conclusion

This report presented innovative solutions to the problems of interplanetary travel and elab-

orated a feasible architecture for Inspiration Mars 2018 fly-by mission. Using COTS hard-

ware and incorporating novel technologies and approaches enabled a low mission cost with

outstanding system performance. This design employs a Dragon capsule and a Bigelow Sun-

dancer module, both of which are close to completion, as well as four launches with the

Falcon Heavy, which will perform its first flight before the end of 2014.

The spacecraft’s radiation shield will employ water to provide protection, in addition to the

radiation protection provided by the structure of the Sundancer habitat. The waste man-

agement system will not only allow for automated disposal of waste and water recycling, but

also provide additional radiation protection to the spacecraft. The ComSat will implement

laser communications in order to provide unprecedented data return from Mars, providing

a defining moment for space exploration as well as expanding humanity’s interplanetary in-

frastructure. This satellite will remain in Martian orbit and constitute the cornerstone to

the planet’s subsequent exploration and potential colonisation.

Four Falcon Heavy launchers will be required, for a total payload into 800km altitude orbit

of 188.8 metric tonnes. The injection into Mars trajectory will be performed by reusing

the upper stages and using some additional fuel carried into orbit, and this will allow for

a final dry mass of 31.8 tonnes to be transported to Mars. The Falcon Heavy launch rate

is sustainable but ensures critical elements will have a minimum loiter time in LEO before

trans-Martian injection. The cost of the mission was calculated with QuickCost and verified

with several other models. This gave a total cost estimate of $2.49B with a 75% confidence

level, $2.10B with a 50% confidence level and $2.85B with a (90%)confidence level.

48


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