Illini Mars Mission for the Opportunity

to Revitalize The American Legacy

Faculty Advisor: Steven J. D’Urso, M.S.

Team Leads: Braven Leung and Christopher Lorenz

Mohammed Alvi ● Alexander Case ● Andrew Clarkson ● Logan Damiani ● Shoham Das

John Fuller ● Thomas Gordon ● Pranika Gupta ● Andrew Holm ● Guangting Lee ● Brandon Leung

Scott Neuhoff ● Anthony Park ● Jeffrey Pekosh ● Sri Krishna Potukuchi ● Kelsey White

1

Table of Contents

I. Abstract ................................................................................................................................................. 3

II. Concept of Operations .......................................................................................................................... 3

III. Launch Vehicles ................................................................................................................................ 5

IV. Orbital Mechanics ............................................................................................................................. 6

V. Propulsion ............................................................................................................................................. 7

VI. Habitat Design ................................................................................................................................ 13

VII. Re-entry Technologies .................................................................................................................... 16

VIII. Power .............................................................................................................................................. 19

IX. Communications ............................................................................................................................. 20

X. Attitude Control and Navigation ......................................................................................................... 21

XI. Environmental Control and Life Support System ........................................................................... 22

XII. Human Factors ................................................................................................................................ 24

XIII. Radiation Protection ........................................................................................................................ 27

XIV. Scientific Return ......................................................................................................................... 29

XV. Cost ................................................................................................................................................. 30

XVI. Risk ............................................................................................................................................. 32

XVII. References ................................................................................................................................... 36

XVIII. Appendix A: Mass Budget .......................................................................................................... 40

2

List of Tables

Table III-1: Launch Vehicle Trade Study ..................................................................................................... 5

Table V-1: Pratt & Whitney RL-10B-2 Engine Specifications [7] ............................................................... 8

Table V-2: LOX Boil-Off Rate for Two Centaur Tank Designs ................................................................ 11

Table V-3: LH2 and LO2 Total Propellant Boil-Off Rates for Two Centaur Tank Designs ....................... 11

Table V-4: Propellant loss summary with standard fuel management ....................................................... 12

Table V-5: Projections of LOX and LH2 Boil-Off Rates with VDMLI Implemented................................ 13

Table V-6: Propellant Loss Summary with Standard Fuel Management System + VDMLI...................... 13

Table VI-1: Habitat Module Trade Study ................................................................................................... 14

Table VII-1: Re-entry Capsule Selection Matrix ........................................................................................ 16

Table IX-1: Communication Trade Study [34] ........................................................................................... 20

Table X-1: Trajectory Correction Maneuvers ............................................................................................. 22

Table XIII-1 Organ Specific Exposure Limits [57] .................................................................................... 29

Table XIII-2: Career Exposure Limits by Age and Gender [56] ................................................................ 29

Table XV-1: Cost Analysis (All values in $MM USD) .............................................................................. 31

Table XV-2: Cost Summary ....................................................................................................................... 31

Table XVII-1: Probability-Impact Scale ..................................................................................................... 32

Table XVII-2: Risk Matrix ......................................................................................................................... 33

Table XVII-3: Risk Analysis for Launch/Deployment Systems ................................................................. 33

Table XVII-4: Risk Analysis for Power/Thermal Systems......................................................................... 34

Table XVII-5: Risk Analysis for ECLSS/Human Factors Systems ............................................................ 34

Table XVII-6: Risk Analysis for Avionics, Controls, and Navigation Systems ......................................... 35

Table XVII-7: Risk Analysis for Spacecraft Structure ............................................................................... 35

List of Figures

Figure II-1: Concept of operations diagram showing the integration of all components. ............................ 4

Figure III-2: Diagrams showing the location of all components within their payload fairings. ................... 6

Figure IV-1: STK Model displaying the orbital track of the flyby mission. ................................................. 7

Figure V-1: Diagram of the Delta Cryogenic Second Stage [4]. [Courtesy: ULA] ...................................... 8

Figure V-2: Subsystem interfaces for a typical Cryogenic Fuel Management System............................... 10

Figure V-3: Cross section of VDMLI demonstrating spacing gradient between layers. ............................ 12

Figure VII-1: Graph describing optimal habitat volume [11]. [Courtesy: NASA MSFC] .......................... 14

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I. Abstract The University of Illinois’ Illini Mars Mission for the Opportunity to Revitalize the American

Legacy (IMMORTAL) is a practical proposal for the chance of achieving a once in a generation

opportunity. The alignment of the planets in January of 2018 offers a unique chance for America

to take the next bold step in mankind’s continuing endeavors to reach farther into the space: the

opportunity to send a man and a woman to fly past Mars and return to Earth quickly and safely.

By taking advantage of an orbital alignment that will not reappear until 2031, it is possible to send

human beings beyond the Moon for the first time since the Apollo program.

Because this unique orbital alignment requires a launch date close to three and a half years

from the present day, the IMMORTAL mission is built around an accelerated timeline.

Consequently, the only way to achieve the mission directive of a manned fly-by mission around

Mars, is to construct the mission architecture around innovated use of existing technologies.

Through the use of heavy launch vehicles, chemical propulsive units, modern heat shielding,

commercial deep space capsules, retrofitted living habitat, and solar power, human beings will sail

around the red planet for the first time.

II. Concept of Operations The IMMORTAL mission architecture requires two launches. The first launch will utilize

the SpaceX Falcon Heavy rocket and will consist of the following

Dragon capsule that will carry the crew into space

Cygnus habitat module where the crew will spend most of the 501 day journey

Service module that holds life support systems for the crew

This launch carries significantly less payload than the estimated lift capacity of the Falcon

Heavy to Low Earth Orbit (LEO). As a result, the Falcon Heavy Upper Stage (FHUS) will have a

significant amount of propellant leftover at burnout. This stage will be retained to help perform

part of the trans-Mars injection (TMI) burn.

Shortly following this launch, a 4 m Delta Cryogenic Second Stage (DCSS) propulsion

module will be launched using a Delta IV Heavy. Subsequently, the DCSS will dock with the

Dragon-Cygnus assembly in LEO. After a series of checkouts, the FHUS will ignite and transfer

the assembly into a highly elliptical orbit around Earth. The spacecraft will then discard this stage

and reorient to perform the second half of the burn using the DCSS. The DCSS will burn at perigee

of the next orbit, setting the craft on its free return trajectory which will take the assembly to Mars.

After the DCSS stage is discarded, the Dragon capsule will detach from the top of the Cygnus

module and perform a maneuver similar to that required by the Apollo missions. This maneuver

will involve the Dragon capsule performing a 180 degree spin to move from its launch position to

the orientation for docking with the Cygnus habitat. The crew will then transfer over to the habitat

for the remainder of their journey. After 224 days, the Dragon-Cygnus vehicle containing the crew

will reach Mars and perform a flyby at an altitude of 100 km. Following this momentous

accomplishment, the crew will spend an additional 271 days in the habitat heading back to Earth

before transferring over to the Dragon capsule for Earth re-entry. This mission places a mass of

15,875 kg on this trajectory with an 11% margin for contingencies. The total budget amounts to

$1,493M USD, according to NASA’s Project Cost Estimating Capability (PCEC) framework.

4

Although higher than the mission proposed by the Inspiration Mars Foundation, this project is still

relatively inexpensive and significantly more feasible within the 2018 timeframe.

Figure II-1: Concept of operations diagram showing the integration of all components.

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III. Launch Vehicles To satisfy the mission requirements, a total of two launches will be made. To confine the

mission to one launch, the Space Launch System (SLS) would be the only option for a payload

consisting of all mission components. The use of SLS Block I would present significant risk with

the first launch planned for December of 2017, well within the vicinity of the IMMORTAL launch

date [1]. The availability of this launch vehicle is uncertain due to potential delays or cancellations.

Similarly, the SLS Block IB is not expected to launch until 2021, which is outside of the mission

timeframe [2]. Alternatively, increasing the launch count to three total launches would introduce

unnecessarily complex orbital docking operations.

The choice of the SpaceX Dragon as the re-entry capsule necessitates the use of the Falcon

Heavy launch vehicle, which has a sufficient lift capacity to transport both the Dragon capsule and

the Cygnus habitat/service modules through a single launch with spare propellant for the transfer.

The large payload capacity and low launch cost of the Falcon Heavy make it a cost-effective choice

for this mission [3].

The second launch includes the 4-meter DCSS, which requires liquid hydrogen fueling.

The Falcon launch pad supports only RP-1 fueling, eliminating the Falcon Heavy as an option.

Consequently, the ULA’s Delta IV Heavy was selected for its payload capacity as well as its ease

of integration with the DCSS, which is a stage of the Delta IV series launch vehicles [4].

Table III-1: Launch Vehicle Trade Study

Launch

Vehicle

Launch

Site

Cost per Launch

(USD in millions)

Payload

Mass (kg) to

LEO

Inclination

28.5°

Cost per kilogram to

200km Inclination 28.5°

($/kg)

Estimated

Availability

Delta IV

Heavy

CCAFS,

VAFB

290 28,790 10,070 In Service

Falcon

Heavy

KSC,

VAFB

135 53,000 2,500 2014

Falcon 9 CCAFS,

VAFB

56.5 (as of 2013) 13,150 4,300 In Service

Atlas V

552

CCAFS,

VAFB

250 20,520 12,180 In Service

SLS Block

I

KSC 500 70,000 7,100 2017

SLS Block

IB

KSC ??? 118,000

2023

6

Figure III-1: Diagrams showing the location of all components within their payload fairings.

IV. Orbital Mechanics On January 1st 2018, SpaceX’s Falcon Heavy will carry the crewed Dragon capsule along

with the Cygnus habitat and service module to a circular Low Earth Orbit with an altitude around

200 km and a period of 90 minutes. Shortly after, a Delta IV Heavy launch will carry the DCSS to

LEO in order to dock with the habitat. In order to mitigate propellant boil-off and to minimize

astronaut downtime, the launches will need to occur in quick succession. This docking maneuver

will rely on highly developed technologies for docking that have been perfected throughout the

life of the International Space Station (ISS) as well as during other manned missions.

After assembly, leftover fuel in the FHUS will be burned first, imparting about 1.70 km/s

delta-v before being discarded. This will place the assembly into a highly elliptical orbit around

Earth with period 4 hours. It will reorient itself and upon reaching perigee the DCSS will perform

the TMI maneuver. An analysis of the capabilities of the Falcon Heavy and the Delta IV Heavy

shows that they are capable of performing such maneuvers [3] [4]. After the FHUS burn, the DCSS

will be required to impart 3.1 km/s into the habitat to perform the flyby. Calculations show that

the DCSS has 3.3 km/s delta-v available for the burn. The TMI maneuver was targeted using STK’s

Astrogator module in order to optimize C3. The method is based on similar methods used in the

IEEE Conference Feasibility Analysis [5]. Using these methods, the DCSS would need to leave

the Earth with a C3 of 39.0 (km/s)2.

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Facilities on the ground will continuously assess the trajectory of the habitat to ensure it

was put onto the proper orbital trajectory, and that the habitat will be put onto the correct free-

return trajectory after the flyby, calculating any possible course corrections that would need to be

made. Over the course of the mission, there are multiple opportunities to perform Trajectory

Correction Maneuvers (TCM’s) if need be. Such maneuvers would be handled by attitude control

system aboard the service module, and would require as little as .5 m/s of delta-v. A TCM at

periaerion of at most 30 m/s would be the upper bound of possible course corrections. The habitat

would be in transit to Mars for 227 days, coming near the planet in Mid-August of 2018 [6]. The

flyby would bring the habitat within 100 miles of the Martian surface before Mars’ gravity would

swing the habitat on its return trajectory. The IMMORTAL habitat will spend a short time near

Mars during which detailed observations of the planet and its moons can be made. 274 days later

in Mid-May of 2019, the habitat will return to Earth, arriving with a velocity 14 km/s relative to

the Earth before entering Earth’s atmosphere. The total delta-v required from LEO to re-entry for

the mission, excepting any extraneous course corrections is 4.8 km/s.

Figure IV-1: STK Model displaying the orbital track of the flyby mission.

V. Propulsion There are two potential Delta-Cryogenic Second Stage (DCSS) configurations that are

worth considering. One is the 4-m version which is used on rockets such as the Delta IV M and

Delta IV M+. There is also the extended 5-m version which is used typically for larger rockets

such as the Delta IV H, or for when more propellant is needed on the Delta IV M or Delta IV M+.

The engine core for both configurations is the Pratt & Whitney RL10B-2 engine. A diagram of the

upper stage is listed below (Figure V-1) along with specifications for the engine (Table V-1). The

extended design is based on the smaller 4-m design. The major differences include a larger LOX

tank, specifically 0.5 m longer in length; as well as a larger LH2 tank, which has an enlarged tank

diameter of 5 m. The propellant load in this larger design has a capacity of 27,200 kg and an

8

estimated burn time of over 1,125 seconds. The propellant management system takes advantage

of hydrogen boil-off by incorporating aft-facing thrusters which provide settling thrust and

controlled propulsion for the attitude-control system. The LH2 tank is re-pressurized using

hydrogen bleed from the system and the LOX tank is re-pressurized using stored helium.

Additional tanks of helium can be implemented if several engine restarts are required for the

mission. Typical loiter duration is around 2.3 hours on average, but loiter times as long as 7 hours

or more could be possible with additional hydrazine tanks. State-of-the-art engine gimbal, attitude-

control, and collision avoidance systems ensure high reliability for the mission [4]. The 5 meter

variant of the DCSS in-space propulsion system would be too massive for the Delta IV heavy, or

any other current liquid hydrogen fueled launch system to lift into orbit. Therefore, The

IMMORTAL mission will make use of the 4 meter variant of the DCSS stage covered in variable

density multi-layer insulation (VDMLI) to reduce propellant boil-off as the second in-space

propulsion stage. The propellant tank of the DCSS will also have a docking port attached to the

top that follows the International Docking System Standard that will allow Dragon to dock to it

during the loitering periods and for the TMI burn.

Figure V-1: Diagram of the Delta Cryogenic Second Stage [4]. [Courtesy: ULA]

As an upgrade to the original RL-10 engine, the RL-10B-2 uses an extendable exit cone to

increase specific impulse and to decrease storage space of the exit cone, and electromechanical

actuators replace hydraulic systems in the engine. The upgraded cryogenic second-stage Pratt &

Whitney RL-10B-2 engine is based on the 30-year heritage of the reliable RL-10 engine. The basic

engine and turbo pump are unchanged relative to the RL-10.

Table V-1: Pratt & Whitney RL-10B-2 Engine Specifications [7]

Thrust 110.00 kN

Specific Impulse 465 s

Burn Time 700 s

Expansion Ratio 250

Weight 277 kg

Nominal Flow Rate 24.1 kg/s

9

Mixture Ratio 5.85 to 1

Dimensions (h x r) 4.14 m x 2.21 m

Propellants LOX & LH2

Status In production

Boil-off can cause a significant loss in propellant if the mission loiters for too long before

TMI and must therefore be considered heavily in mission design. Typical magnitudes of boil-off

rates for LH2 and LOX tend to be at a minimum of several 10-1 kg/hr and a maximum magnitude

of 101 kg/hr. These rates are largely affected by the external heat flux penetrating the propellant

and less to do with the total mass of the propellant. Therefore it is of utmost of importance to

minimize the effects from solar radiation and from the heat leaked of internal components in order

to reduce propellant boil-off [8].

When coasting times are low, boil-off is typically dealt with by implementing various

control techniques. Currently the system in place is the typical Cryogenic Fuel Management

System, which essentially maintains a low temperature in the propellant tank to prevent the

diffusion and destratification of liquid phase propellant by various means. The fuel management

system is typically designed with primary objectives including effective propellant storage with

minimal loss, vapor-free propellant distribution including the inlets and outlets, and a robust

control system which minimizes propellant settling.

Such a system will have various implications on the design, not only limited to the

propellant boil-off rates. A Cryogenic system can greatly reduce the propellant launch mass and

the on-orbit margins. Such a design even tackles the complex problem of propellant settling and

allows major system benefits by simplifying and optimizing the system architecture. There is

already a pre-existing design for the Cryogenic Management system, developed by Lockheed

Martin called the ICES (Integrated Common Evolved Storage) developed for the Centaur Upper

Stage, which revolves around these goals. Centaur uses the same RL-10 engine as the DCSS in-

space propulsion stage being used on this mission, so integration of the ICES or similar fuel

management systems should be fairly compatible with the DCSS systems. With minor

modifications such systems can even be implemented for long duration missions, such modified

systems are generally for significantly larger missions to the outer solar system [9]. Although the

propellant will be expended on a much sooner into the mission than these long duration mission

modifications would be intended for, aspects of the long duration solutions can certainly be

adopted in order to preserve fuel boil-off in general.

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Figure V-2: Subsystem interfaces for a typical Cryogenic Fuel Management System.

The fuel management system outlined above has limitations, in that the system is not

designed for long duration missions. Because the mission’s main propulsive maneuvers will be

done early on in the mission duration, this is not a problem. However, as mentioned before, the

fact that the LOX/LH2 tank will loiter in space for up to 24 hours is important to take into

consideration, as well as the time the propellant tank will take to reach the initial payload in the

subsequent launch; especially when considering the higher boil-off rate of LH2. Design

modifications can very easily improve the existing design to the mission objective particularly by

implementing improved passive storage.

Much can be learned about the boil-off behavior of the DCSS by looking at a very similar

upper stage, the Centaur. The Centaur tank is a very good example to compare the DCSS; not only

are the engines the same but the tanks are approximately the same size, and as mentioned before,

the primary contributing factor to boil-off is the thermal influence of the surrounding environment,

which is assumed to be very similar for the Centaur and DCSS in orbit. The boil-off rates for the

Centaur propulsion system has been examined extensively on several occasion by ULA [10]. They

have found that the Centaur has typical boil-off rates of 1-4.8% per day. The variation in the boil-

off rate here was based on different degrees of heat flux entering the controlled propellant tank

and the type of insulation used. This can become problematic if the loiter time becomes too high

while waiting for TMI. Minimizing boil-off is critical, since launching excess propellant which

will ultimately diffuse is wasteful of the limited payload mass, and when considering the length of

such an interplanetary mission, payload mass is extremely critical. Furthermore, since the mission

involves launching two separate payloads a system to decrease propellant loss in the first launch

which will have very large loiter times is necessary.

Lockheed Martin, in association with ULA has presented various boil-off rates for different

operation conditions for LH2 and LOX tanks in the Centaur Upper stage. These values were

reconstructed from post-flight measurements obtained from the tank heating system.

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Table V-2: LOX Boil-Off Rate for Two Centaur Tank Designs

Oxygen Boil-Off Hydrogen Boil-Off

Tank Design TC-15 TC-11 TC-15 TC-11

Tank

Insulation

3 Layer Radiation

Shield

3 Layer Radiation

Shield

3 Layer Radiation

Shield

3 Layer Radiation

Shield

Total Heat

Flux

615 Watts 381 Watts 733 Watts 909 Watts

Boil-off Per

Day

1.5% 1.0% 4.1% 5.1%

Table V-3: LH2 and LO2 Total Propellant Boil-Off Rates for Two Centaur Tank Designs

TC-15 TC-11

Propellant LOX LH2 LOX LH2

Single Prop. Boil-Off Rate 1.5% 4.1% 1.0% 5.1%

Total System Boil-Off Rate 2% - 1.6% -

The boil-off rates from the above study agree with the previous estimations of boil-off

rates. The Heat Flux values of approximately 733 Watts is a reasonable value for the design

parameter. As mentioned before the Centaur boil-off rates will be sufficient to do preliminary

calculations of propellant losses for the mission. Still, a conservative estimate would be preferred

for calculating the worst case scenario, and the largest reasonable boil-off rate values will be

chosen to proceed with the estimation of propellant loss. As such it will be assumed that the total

system boil-off rate will be 2.0% and LOX tank boil-off rate to be 1.5%. The total system boil-off

rate is the total mass of LOX and LH2 lost as a fraction of the total launch propellant mass per day.

The propellant loss will be a function of the boil-off rate, the amount of time spent loitering,

and the total mass of propellant during launch. In terms of the time spent loitering in orbit the

difference between the initial launch which uses RP-1 and LOX, and the second launch which uses

LH2 and LOX launch will be approximately one day in the worst case scenario. The SpaceX launch

will come from pad 39A at Cape Canaveral and ULA will launch the Delta IV Heavy. Assuming

a launch to orbital docking time of 4 hours for the second payload the total time spent loitering by

the initial oxygen tank can be estimated to be 28 hours and the time for the second LH2/LOX tank

to be 4 hours. Note the RP-1 boil-off is not considered since it non-cryogenic and will have

insignificant boil-off when compared to LH2 and LOX. The initial RP-1/LOX will contain 30,300

kg of propellant mass once it is in orbit and the secondary launch will contain 20,410 kg of launch

propellant LH2/LOX. The Oxidizer/Fuel Ratio of RP-1/LOX is 2.77 and thus it can be determined

that the LOX mass in the initial launch is 22,263 kg.

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Table V-4: Propellant loss summary with standard fuel management

Propellant Mass Boil-Off Rate Loitering Time Propellant Lost

Falcon LOX Tank 22,263 kg 1.5% per day 28 hours 389 kg

DCSS LOX/LH2 20,410 kg 2.0% per day 4 hours 68 kg

These values with the above conventional cryogenic solutions lead to noticeably large

propellant loss for the mission as seen from the summary chart above. 389 kg is a significant

portion of the payload mass and thus a novel solution to greatly improve the boil-off loss is

required. The simplest solution will be in the form of passive-storage improvements, the most

promising of which is the newly developed VDMLI material.

As mentioned the greatest contributing factor to boil-off is the influence of the heat flux

from the surrounding environment. If better thermal insulation is implemented on the propellant

tanks, propellant loss can be significantly minimized, without adding excessive weight to the

spacecraft. The current standard Multiple Layer Insulation has a new design improvement referred

to as VDMLI. Variable density MLI is a type of insulation material that optimizes the radiation

insulation capability relative to standard MLI by having a gradient of spacing in between each of

the layers of insulation. The concept is to space the inner layers (closer to the cold tank wall)

further apart than the outer layers (closer to the warm radiation region) where the bulk of the

radiation heat transfer penetrates. The spaces are held fast by bumped ridges that keep the layers

spaced. Thus for a given insulation mass the design achieves an optimal insulation via this variable

spacing. The drawback with the technique is the slightly larger volume due to the larger spacing.

Figure V-3: Cross section of VDMLI demonstrating spacing gradient between layers.

Lockheed Martin has also conducted studies and projections of potential design

improvements for their Centaur upper stage using this VDMLI instead of the standard MLI used

currently [8]. The VDMLI can greatly reduce the amount of heat flux entering the tank and

therefore greatly reduce boil-off and save propellant mass that would otherwise have been

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launched and lost. Take into consideration the VDMLI heating values and boil-off values seen

below:

Table V-5: Projections of LOX and LH2 Boil-Off Rates with VDMLI Implemented

LOX LH2

Heat Flux 950 Btu/hr 1350 Btu/hr

Single Species Boil-Off per day 0.7 % 2.4 %

Total Prop. Boil-Off per day 1.0 %

As can be seen, the previously estimated LH2/LOX and LOX boil-off rates were almost

twice as large as the estimated rates with VDMLI implemented. Using improvements to lower

these boil-off rates in the mission design will therefore create significant improvements in limiting

propellant loss.

Table V-6: Propellant Loss Summary with Standard Fuel Management System + VDMLI

Propellant Mass Boil-Off Rate Loitering Time Propellant Lost

Falcon LOX Tank 22,263 kg 0.7% per day 28 hours 182 kg

DCSS LOX/LH2 20,410 kg 1.0 % per day 4 hours 34 kg

The typical weight of standard MLI is approximately 56 kg for the dimensions of an upper

stage tank. The weight for the VDMLI insulation of the same tank would be 57 kg. For such a

marginal cost in dry mass, the propellant loss can be greatly reduced and approximately 241 kg

less of propellant mass launched if VDMLI is implemented on both the RP-1/LOX tank and the

LH2/LOX tanks of both in-space propulsion systems utilized on the mission.

VI. Habitat Design This two person fly-by mission of Mars will require sufficient space for the astronauts to

live for just over 500 days. In a study published by the NASA Johnson Space Center Figure VII-

1, the total habitable volume per crewmember was calculated for tolerable, performance, and

optimal limits. Based on chart, the needed volume per person is 5m3 for tolerable, 10.5m3 for

performance, and 19m3 for optimal. Because of the length of this mission, the goal is to construct

a habitat between the performance and optimal range, that being 21m3 to 38m3.

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Figure VI-1: Graph describing optimal habitat volume [11]. [Courtesy: NASA MSFC]

Aside from the habitable volume constraint, the mass and availability of the structure will

determine the habitat chosen for this mission. Because this mission is focused on the revitalization

of the American legacy, only American made habitats were considered. Using a modified version

of the modules utilized by the ISS was considered. An example of this type of structure is the

Destiny Module constructed by Boeing for NASA. This module has a high pressurized volume of

106m3 which would be extremely desirable for two crew-members to function for 501 days;

however, this module has a very high dry mass of 14,515 kg [12]. This mass is simply too high to

be supported by the current mission architecture. Thus, the ISS modules are not valid options.

There are two main companies which produce these possible habitats for use in this mission:

Orbital Sciences Corporation and Bigelow Aerospace. The options which these two companies

provide are listed in Table VII-1.

Table VI-1: Habitat Module Trade Study

Manufacturer

Type Pressurized

Volume (cubic

meters)

Height

(meters)

Mass (dry)

(kilograms)

Status

Bigelow Genesis 11.5 2.54 1,360 Successful

Bigelow Sundancer 180 8.7 8,618 Cancelled

Bigelow BA

330

330 14 20,000 In-design

Cygnus Standard 18.9 3.66 1,500 Successful

Cygnus Enhanced 27 4.86 1,800 Launch

in early 2015

Cygnus Two standard

(proposed)

37.8 3.66 3,000 Proposed

15

Cygnus Super

enhanced

(proposed)

35 6.06 2,300 Final Design

selected

(proposed)

To date, Bigelow Aerospace has four concepts for in space habitats. All of Bigelow’s

habitats are modules that inflate upon reaching orbit, allowing for a lower structural mass as well

as a smaller launch package. This concept presents an issue of finding enough space in the pre-

inflated structure to store all of the supplies needed for this mission. In addition, the crew would

be required to setup their habitat upon reaching orbit, taking up valuable crew time. The Bigelow

Genesis I and II modules were the first two habitats constructed by Bigelow, and they are the only

ones developed by Bigelow to ever fly in space. Despite being lightweight, this habitat is simply

has too little volume (11.5m3) to sustain the two crew-members for 501 days [13]. The next concept

proposed by Bigelow was the Sundancer model. This design concept had a large pressurized

volume of 180m3 with a mass of 8618 kg. The module would be too heavy to be used as a part of

any realistic Inspiration Mars architecture and was also cancelled by Bigelow in July of 2011; thus,

it is not a viable option for this mission [14]. Bigelow’s next concept, the BA-330, has a pressurized

volume larger than ever would be needed for a two person, and, consequently, the proposed mass

for this structure heavily exceeds the limitations of the mission architecture [15]. In the end, all of

the Bigelow concepts do not meet this mission’s requirements.

Currently there are two models of Cygnus designed: Standard and Enhanced. Both modules

have a low mass which fits into the mission architecture. The standard has already had two

successful missions which involved delivering pressurized cargo to the ISS. The Enhanced model

will be tested in early 2015. The Enhanced module is effectively a Standard module extended from

its original size by increasing the height of the module by 1.2m. This increased the pressurized

volume from about 18.9 cubic meters to 27 cubic meters. To support these increased dimensions

the service module has been made lighter and more efficient and the propulsion has been boosted

[16]. The Standard module’s pressurized volume would lead to an incredibly small living quarters

for this two-person crew to survive for 501 days. The Enhanced module has a pressurized volume

larger than the minimum volume requirement; however, not all of the pressurized volume is usable

for habitable volume. Based on calculations NASA used for the ISS, the habitable volume of the

Enhanced module would drop below the minimum volume requirement of 21 m3 [17].

Consequently, the Enhanced section is not a feasible option for this mission.

In order to facilitate the habitable volume needs of the crew, a modified Cygnus capsule

architecture is necessary. The first proposal calls for using two Cygnus Standard modules which

would dock together in orbit via conjoining module which would be need to be researched and

developed. Despite doubling the available habitable volume of a Standard module, this proposal

would require complex orbital assembly to ensure proper docking of the two Standard modules

which is undesirable. Instead, a Cygnus Super-Enhanced is proposed. This module would be

Standard module stretched to 2.4m instead of the 1.2 m stretch applied to the basic Enhanced

Module. The proposed module would have a mass of 2,300 kg with a total of height 6.06 m. This

proposal avoids simply doubling the mass of the Standard module by only extending the center of

the Standard module, such that there are still only the two end caps, and utilizing only a single

service module to house all necessary support systems. The Super-Enhanced module would have

a PCM with the standard 3.07 m diameter and a total pressurized volume of 35 cubic meters [18].

Once again, based on the calculations used by NASA for ISS and once an estimated 5 m3 usable

storage volume onboard the Dragon capsule has been factored in, a total habitable volume of 24.4

16

m3 is obtained. This volume is within the acceptable range of habitable volume in order for the

two-person crew to function efficiently, and the mass of this proposal is possible with the current

mission architecture and chosen launch vehicles.

VII. Re-entry Technologies The most critical part of the entire mission is the re-entry phase. If the crew fails to re-enter

safely, the mission will have been a catastrophic failure. As such, it is the primary driving force

for the design of the entire mission. A return capsule must be chosen that can withstand the re-

entry heating that will be experienced coming back into the Earth’s atmosphere. Based on

simulations, the re-entry vehicles would need to handle speeds of 14 km/s, potentially up to as

much as 14.2 km/s. If it cannot withstand these speeds, the capsule will have to be slowed to a

survivable velocity before re-entering the atmosphere. This will significantly increase the size of

the launch vehicles and in-space propulsion stages. Once the capsule and re-entry method are

determined, launch vehicle architecture can be chosen to support the rest of the mission.

No matter what capsule is chosen, there are other critical issues that must be addressed.

This mission will require re-entry velocities of up to 14.2 km/s, and to date, the fastest re-entry

that has ever been achieved was 12.8 km/s, achieved by the Stardust sample return capsule [19].

No re-entry system has been tested to velocities higher than that in practice, and facilities do not

currently exist that can simulate the re-entry velocities and temperatures required. In order to

properly test the capabilities of the heat shield to the required extremes, existing facilities will have

to upgraded or new facilities constructed, which will take time and will cost a significant sum of

money. There are however, currently plans to upgrade existing facilities; this will offset the cost

of development and construction of the new facility [20].

For the choice of re-entry vehicle, four different options will be considered and are

presented in the table below:

Table VII-1: Re-entry Capsule Selection Matrix

Vehicle Heat Shield Material Plausible Manned Launch Date

Orion MPCV AVCOAT 2021

Orion Pathfinder AVCOAT 2023

CST-100 Boeing Lightweight Ablator 2016

Dragon PICA-X 2015

The Orion MPCV is one of the more promising candidates for the re-entry capsule. It is

meant to be used for deep-space and interplanetary missions, it will be meant to handle extreme

re-entry velocities and long duration missions outside of Earth’s sphere of influence. Orion’s heat

shield will be made of a reformulation of AVCOAT, the material that was proven on the Apollo

missions to withstand a re-entry velocity of 11 km/s [21].

Even though this capsule will be able to withstand high re-entry speeds, it is a very massive

capsule. Orion is expected to have a mass of 9,820 kg fully loaded [22]. As a result, it will require

17

the use of larger launch vehicles and in space propulsion stages than other, lighter capsules. The

Orion capsule is designed for a crew of two to six people. An even more critical issue with utilizing

Orion is that at the current rate of development it will not be ready for a manned mission to deep

space in time for this flyby mission. The first unmanned flight of Orion is planned for September

2014 as a test of its basic subsystems and as a test to verify its ability to withstand re-entry. If this

first test mission goes well, the next unmanned mission is a Lunar flyby tentatively scheduled for

December 2017 to prove Orion’s ability to facilitate a manned mission to the Moon within a few

years later, potentially as early as 2021 [23]. This second unmanned mission would be launching

at essentially the same time as the Mars flyby opportunity. Using Orion in 2018 would mean using

a largely unproven capsule, and it will require significantly advancing NASA’s development

timetable. If Orion undergoes a failure in its first launch or some other major setback occurs, it

will delay the program significantly. Any delay would be disastrous to meeting the tight schedule

of this flyby mission.

It has been proposed that a lighter variant of Orion, called Orion Pathfinder, be developed

in order to address the mass concerns relating to Orion’s excessive size [24]. This proposal,

however, ignores the time requirements of this mission. The Orion Pathfinder capsule is a

completely unrefined design which has not been seriously developed yet and will almost definitely

not be available in time for an early 2018 launch.

Utilizing the standard Orion MPCV for the flyby is a dubious proposition at best that would

require accelerating Orion’s development. Proper development has not even been begun on the

Pathfinder variant, even though it would have striking similarities to the basic Orion system, years

of development and a significant budget will be necessary to bring Pathfinder to fruition. To

design, develop, build, and test Pathfinder sufficiently in less than 4 years to a level which could

justify a manned mission would require an extraordinary level of dedication to the project not seen

since the Apollo era, and it is a hopelessly unrealistic expectation, as even one minor setback will

likely render a 2018 timeline unreachable [20].

Another capsule that presents itself as a candidate for the crew ascent and re-entry vehicle

is the Boeing manufactured CST-100 capsule. The CST-100 capsule is primarily meant for

transporting astronauts between the ground and LEO. This capsule is only designed to reach 400

km orbital altitude, the maximum altitude of the ISS; thus rendering it useless for deep space travel.

It is less massive than the Orion MPCV, but is also a much simpler vehicle, keeping in line with

its basic mission requirements as a taxi vehicle. This simplicity includes a limited heat-shield that

can only withstand up to 1700 Co [25]. This weak heat shield coupled with its relatively low

maximum altitude make the CST-100 capsule a mostly obsolete choice as the re-entry vehicle for

this mission.

The final option for a re-entry capsule is the SpaceX Dragon capsule. The Dragon capsule

is designed to be used for commercial resupply missions as well as a deep space exploration

vehicle. One of the reasons Dragon can operate as a deep space vehicle is its advanced heat shield.

Dragon has a heat shield made of PICA-X, a variant of the PICA material which was used on the

NASA Stardust mission which achieved the highest Earth re-entry velocity of any vehicle to date

of 12.8 km/s. According to SpaceX, Dragon’s heat shield is much thicker than required for LEO

missions; thus enabling it to serve as a deep space vehicle [26].

Dragon indeed does have several other advantages over the other choices of re-entry

capsules. First off, Dragon is lighter than the other deep space vehicle under consideration, the

Orion MPCV. The manned version of Dragon is estimated to weigh 6,000 kg fully loaded,

contrasted with the 9,820 kg of the Orion MPCV. This number has been estimated from released

18

data about the capsule by SpaceX since they have not released the final number yet for the manned

version of Dragon. The unmanned version currently has a 4,200 kg dry mass, is a reasonable

estimate based on the proposed changes [5]. Dragon is significantly lighter than Orion making the

more economical choice for a two person fly-by mission; consequently, this makes Dragon a clear

choice over Orion. As stated above, Dragon is intended to serve as both a commercial resupply

vehicle and deep space transportation vehicle. This simple fact it designed to go outside of LEO

makes it a clear choice over Boeing’s CST-100 capsule.

Overall, the Dragon capsule best suits the needs of this mission. It is relatively lightweight

at 6,000 kg. Its heat shield, constructed of PICA-X, has been proven to withstand re-entry at 12.8

km/s which is the highest ever successful re-entry speed recorded [27]. Additionally, Dragon has

sufficient space for a two person re-entry mission since it is designed to hold up to seven people.

Additionally, Dragon’s operational life is rated to be up to two years which is less than the overall

time of this fly-by mission [26]. In conclusion, the SpaceX Dragon capsule is the best suited for

this mission.

Before the Dragon capsule can be used as the re-entry capsule, it is necessary to ensure that

capsule can survive the high speed re-entry at 14.2 km/s. There are three main ways that this can

be done. First off, the Dragon capsule as it is manufactured by SpaceX with no modifications can

be used as a viable option, because the company’s technical specifications which state that the

Dragon Capsule can withstand standard Mars return trip re-entry speeds [26]. Secondly, a simple

option to bolster the standard heat shield of Dragon would be to simply make it thicker to be more

resilient to account for the increased Mars return velocity. The final option would be to add a retro-

propulsive rocket motor to Dragon to slow the vehicle down to 12.8 km/s or less as it hits the

atmosphere.

If it were determined that the Dragon heat shield cannot withstand the re-entry speeds of

up to 14.2 km/s, the re-entry velocity of Dragon will have to be lowered to a velocity that it will

survive re-entry. Dragon’s PICA-X heat shield material was based off of the NASA developed

PICA material which was used on the Stardust probe which was proven to a re-entry speed of 12.8

km/s [27]. This 12.8 km/s consequently represents the velocity threshold to be targeted for retro-

propulsion until a better number is identified via testing the heat shield material.

Additional mass margins for the mission will be put towards adding a thicker heat shield if

it is determined that the basic Dragon heat shield is not sufficient for Earth re-entry at the required

velocities; as well as there will be additional budgets set aside to pay for the development and

testing of said heat shield. However, there currently are no facilities that can test re-entry speeds

of 14.2 km/s. There are plans for test facility upgrades, but they will only be to sufficient to

simulate a heating level of 2000 W/cm2, which represents be a significant improvement to the

current 1400 W/cm2 [20]. In theory, this increase will be enough to test the improved heat shield.

Once this facility operational, testing would be paid for in order to determine how much thicker of

a heat shield was needed than the stock thickness.

The ideal scenario for adding a retro-propulsive rocket motor would be in the case where

only the crew and the most essential supplies remained in the capsule, and only a minimal amount

of RCS propellant remaining. If a STAR 63F solid rocket motor were attached to Dragon for the

purposes of doing a retro-propulsive maneuver to slow it for re-entry, based on our calculations,

an additional 1700 m/s of delta-V can be acquired assuming a nearly empty Dragon mass of 5000

kg [28]. This would add an additional 4600 kg of mass to the re-entry unit of the spacecraft, and

bring the spacecraft velocity down to 12.3-12.5 km/s depending on the exact re-entry velocity of

the craft, well under the 12.8 km/s threshold. Based on our calculations, this additional mass will

19

result in an increase in the mass of the in-space propulsion stage of upwards of 30%, representing

a corresponding increase in launch vehicle size. In order to accommodate the increase in payload

to orbit required, the mission will likely require a third launch, or require the utilization of SLS as

a heavy-lift launch vehicle. This would further increase the cost and complexity of the mission

significantly. The preferred option would be to increase the thickness of the Dragon heat shield so

that it will be able to withstand the re-entry heating that will be experienced at unmitigated

velocities, as it represents an option that is operationally much simpler and much cheaper than

implementing retro-propulsion. For this reason, retro-propulsion will be left as a last resort, in the

event of a failure to prove or develop a heat shield sufficient for the mission.

Overall, simply making the heat shield thicker on the Dragon Capsule outweighs the cost

of using retro-rockets because of the mass and economic savings. The driving force behind this

decision is that SpaceX’s technical specifications for the Dragon Capsule rate it to survive

traditional Mars return velocities. Since this mission will be returning with a non-standard Mars

return velocity, it will be necessary to increase the thickness of the heat shield which will be

determined through testing at the updated facility mentioned above. Based on rough calculations,

a 2-2.5 times the stock thickness of Dragon should be acceptable from this mission factoring in a

reasonable safety factor.

In the final days of the mission, as the re-entry capsule nears Earth, all non-essential

supplies and equipment will be moved from the capsule to the habitat as a lighter re-entry capsule

will be slowed down more effectively during re-entry. The habitat will then detach from the

capsule two hours before entry begins, and the capsule will orient itself for re-entry. Dragon will

then undergo re-entry through Earth’s atmosphere at 14 km/s to bring the capsule safely back to

Earth. The parachute landing systems of Dragon will be utilized as the primary landing system due

to the simplicity and reliability that they offer, rather than the propulsive landing system; the fuel

for the propulsive landing system will be used both for attitude control and for midflight course

corrections. Once the capsule has slowed sufficiently and reached 13.7 km altitude the drogue

parachutes will deploy, followed by the main parachutes at around 3 km altitude [29]. A water

landing is preferable to a land touchdown if orbital adjustments are able provide for it, but land

touchdown is acceptable if there is no viable alternative, as just making re-entry is of greater

concern than making a specific landing area for the purposes of this mission. Any attempt to target

a landing site will have to be done during the course corrections that will initially be quite sensitive

to the corrections, and will become increasingly difficult to change as the end of the mission

approaches and the re-entry corridor restricts possible trajectories.

VIII. Power The power storage and generation systems for the mission will be largely localized in the

Cygnus habitat’s service module. The mission will utilize solar power, along with batteries to

cover any possible power shortages due to shadowing or any other technical issues. Nuclear power

was not considered due to timeline constraints and a lack of in space development for these

technologies, and it was not considered necessary, as this mission is close enough to the Sun that

solar power is still cost and mass effective at the maximum mission aphelion distance.

The SpaceX Dragon capsule will utilize Lithium-polymer batteries for its various power

requirements during launch and flight. The solar arrays on the connected Cygnus habitat module

will provide power to the Dragon capsule whilst in orbit, as the Dragon trunk will not be used.

20

The default Cygnus habitat module will produce around 3.5 kW of solar power at 1 AU

distance from the sun, with the stock design on the Cygnus capsule [30]. Additional solar arrays

will be necessary, as the power requirements for a long-term manned mission are higher than those

required for low-Earth orbit loitering. In addition, Mars will be at a distance of 1.39 AU at mission

aphelion, very close to Mars perihelion, which means that the solar radiation levels there would be

52% of normal Earth levels. The solar panels will only be able to produce about half the amount

of power that is available in low-Earth orbit; less than at any other point in the mission. This means

the solar panels must be sized according to the amount of peak power that would be required at

the mission aphelion. Excess power will be used to charge on-board batteries, however, since

shadowing will be minimal, the rest of the excess power will have to be shunted at the arrays.

Ni/H2 batteries will be used as primary batteries in the habitat for times when shadowing is present

or if the power system cannot otherwise cope with power demand. 28 VDC and 120 VDC and

other electronics, similar to those used on the ISS, will be used for all systems of the IMMORTAL

mission [31].

NASA predicts a power requirement of 18kW for a 500-day habitat configuration based

off of ISS modules, which requires approximately 36 m2 of solar arrays at 60 W/kg [32]. Based on

the habitat’s much smaller volume, 10 kW will be sufficient to power all spacecraft components

in this mission, which requires approximately 20 m2 of solar arrays at 60 W/kg. However, since

available sunlight at a Martian distance from the sun is about half, a total solar array area of 40 m2

will be needed.

The solar arrays will be fitted with Multi-junction Inverted Metamorphic (IMM) Solar

Cells [32]. The technology readiness level of such solar arrays is 8. Lighter ATK Ultraflex solar

arrays could be used to provide additional mass savings if they prove to be ready by launch date,

but at the date of this writing, they are at a technology readiness level of 6 [33]. No additional

funding will be provided within the scope of the IMMORTAL architecture for these panels, but

should they be developed independently they could be substituted. The entire power system with

standard solar panels will have an approximate mass of 445 kg, including wiring, solar arrays, and

batteries.

IX. Communications By examining the current technologies in development and available, the final technology

choices have been made regarding the communications equipment.

Table IX-1: Communication Trade Study [34]

Infrastructure

Requirements Distance Data Rate

Top Technology

Choice TRL Final Technology Used TRL

Mars Low Orbit to

Mars Orbiters ~400 km 10 Mbps Optical 4 UHF 9

Mars Orbiters to

Earth-Sun, L3, L4

Relay Link ~2.5 AU 100 Mbps Optical 4 UHF 9

Sun, L3, L4 to

Earth Orbit/Ground 1 AU 100 Mbps Optical 4 X-band 9

Emergency Coms.

High Grain 2.5 AU 10 Mbps X-band 9 X-band 9

Emergency Coms.

Low Grain 2.5 AU 1 Mbps Ka-band 9 Ka-band 9

21

Data will be transmitted over three different antennas: A high-gain antenna to handle the

high rate communication directly to Earth or via DSN. It has to be able to “aim” itself to within 2-

3 degrees of its target to uplink and downlink data. A UHF software designed radio will be used

to send large quantities of data to the DSN for the fastest communication with Earth. The UHF

will be the primary form of communication for the spacecraft. A low gain antenna will

continuously communicate with other spacecraft and satellites as part of the DSN. The low gain

antenna can only handle low data rates so it will be primarily used to receive data.

The data rates needed for communication between a spacecraft carrying humans and Earth

range between 1 Mbps and 100 Mbps, and sometimes even greater and need to be transmitted with

Ultra-High Frequency (UHF) relay and X band technology. The low gain antenna will have a low

data rate, around 1 Mbps, and the UHF will have speeds of around 150 Mbps. Data will primarily

be transmitted through the Deep Space Network (DSN) along the bidirectional backbone data relay

and secondarily through direct communication with antennas on Earth.

The spacecraft will also be equipped with a transponder. The transponder will have a

transmitter to generate the tone and radio frequency called a carrier wave to be amplified by the

antennas mention above. The receiver in the transponder will take the incoming radio signals or

uplinks and convert them into a perceptible form.

When the controlling body wants to communicate with the spacecraft directly or indirectly

through the DSN it will use the low gain antenna to determine the exact location and orientation

of the spacecraft and will then use that information to position the high gain antenna to

communicate with the satellite via the DSN [35].

X. Attitude Control and Navigation Throughout the duration of the mission, it is critical that the module is able to maintain

heading and bearing in the vast emptiness of space. Due to the nature of the mission, being a single

burn, free-return trajectory, makes it imperative that there be no flaws in the implementation of the

Reaction Control System (RCS) or navigational aids.

The RCS package which designed for this mission incorporates data collected from a pair

of Inertial Measurement Units (IMU), in order to calculate vehicle position and orientation.

However, due to the nature of IMUs, which rely on a combination of gyroscopes and

accelerometers to estimate position through successive calculation based on elapsed time and

spacecraft velocity and heading, the mission will also employ dual Solid State Star Trackers (SS),

which utilize star maps to detect the attitude of the craft, the data from which is then used to correct

for compounded IMU error [36]. These sensors should be mounted so as to allow for the largest

field of view possible, which will help to further reduce uncertainty in calculation. The last

navigational tools that the IMMORTAL mission will include are a Sun Sensor, which simply

locates the position of the Sun relative to the vehicle, for added accuracy in attitude estimation and

therefore increased precision in position calculation by the onboard flight computer, as well as a

Horizon Sensor, which will be used to further minimize error upon final approach to Earth for re-

entry, perhaps the most critical part of the mission.

In addition to these navigational aids, the module will employ two pairs of 3 Aerojet MR-

107V hydrazine thrusters, configured to provide omnidirectional attitude control without

imparting a translation to the vehicle. These 220N thrusters will provide the necessary impulse for

effective attitude control and trajectory manipulation throughout the mission, to ensure that the

spacecraft does not deviate from the flight plan [37]. For further precision, two sets of momentum

22

disks will be used, also arranged to provide 3-axis control, as these disks can create slight attitude

changes due to their small mass compared to that of the vehicle itself.

The Attitude Control & Navigation system includes several redundancies to protect against

the loss of vehicle maneuverability and the degradation of accuracy in positioning. Such situations

would prove catastrophic for mission success, since without fuel for a return burn, crew recovery

will be impossible unless accurate headings are maintained throughout the duration of the mission.

The seamless integration of these sensors and actuators will allow for slight changes in

attitude to be imparted to the vehicle during the 7 predefined TCM windows, which are outlined

in Table XI-1, shown below:

Table X-1: Trajectory Correction Maneuvers

Maneuver TCM Window Purpose

TCM-1 Days 15-20 Place vehicle on Mars fly-by trajectory

TCM-2 Days 110-115 Correct for TCM-1 execution errors

TCM-3 Days 190-195 Further corrections, if necessary

TCM-4 Day 225 Align vehicle for optimal free-return trajectory

TCM-5 Days 300-305 Return trajectory adjustment

TCM-6 Days 400-405 Correct for TCM-5 execution errors

TCM-7 Day 499 Final vehicle alignment for orbital insertion before re-entry

The initial launches of the Habitat and Propulsive Stage provides a unique challenge insofar

as the two modules must dock before continuing towards Mars. In order to complete this maneuver,

the propulsive stage will provide positional data from its onboard computers to the Cygnus module

in real-time, to facilitate the docking procedure between the two.

As the craft completes the initial burn before entering the cruise phase of the mission, it

must maintain a heading which will ensure that the propulsive stage does not strike Mars, as both

the stage and habitat will continue in the same direction after separation. Since the propulsive stage

will not have been decontaminated, it must be guaranteed that foreign microbes are not introduced

into the Mars atmosphere according to planetary protection protocols TCM-1 will then be used to

alter the course of the module to the planned fly-by route [38].

Through the implementation of the above systems, the IMMORTAL mission will find no

issue in executing its fly-by maneuver past Mars on day 227, and then continuing on to complete

its 501-day traverse of local space following re-entry upon arrival back at Earth.

XI. Environmental Control and Life Support System The purpose of the ECLS systems is to regenerate air, water, and food in a manner that

minimizes overall logistical burdens, and minimize demands on space habitat resources while

promoting self-sufficiency and ensuring habitability. It will designed to ensure maximum

redundancy as well as ease of access for repair in the case of any failure. The ECLS is responsible

23

for the effective environmental control and monitoring as well as waste management and water

recovery. It does all this while still providing for in situ maintenance. The ECLS will use closed

loop subsystems to maximize efficiency. As the IMMORTAL architecture does not allow for any

extra vehicular activity (EVA), all critical sections of the ECLS system must be serviceable from

within the crewed space. Most of these systems will be located inside the service module, but

access to them should be incorporated into the design of the modified Cygnus.

The Water Recovery and Management system is responsible for the physicochemical

systems used to increase efficiency and decrease cost of water recovery. Its purpose is to reduce

the Equivalent System Mass of water recovery subsystems and integrated systems while providing

for and supporting long duration integrated life support systems. The WRM is divided into two

separate collection and distribution loops. One is used to recover condensate to drinkable standards

for crew consumption, and the other is used to recover waste hygiene water back to hygiene

standards for crew bathing and equipment use. Crew urine is collected and processed separately

through distillation and then added to the waste collection side of the Hygiene Loop. The resulting

water of the WRS meets the highest standards for potable use while also playing a crucial role in

life support systems mainly feeding the IMMORTAL habitat’s oxygen generator which works off

of electrolysis and the Sabatier method. The Water Recovery and Management system for the

IMMORTAL habitat will have a total system mass of 675 kg, including spares for components

that are likely to fail. This number is in addition to the 1225 kg of water that will be kept on board

for fueling the oxygen generators, the WRS, and the ATCS water coolant loops.

The main purpose of the Air Revitalization System is to control the relative humidity of

the cabin and keep it between 30% and 75%, provide cooling to the IMMORTAL habitat’s cabin

compartment while monitoring temperature and ventilation, and maintaining carbon dioxide and

carbon monoxide at nontoxic levels. The Air Revitalization Systems incorporate water coolant

loops, cabin loops, and pressure controls to have the habitat’s two person crew atmospheric habitat

safe and ventilated. The habitat will utilize current technologies already aboard the ISS Destiny

module including several removal techniques for separating CO2 such as permeable membranes,

liquid amine, adsorbents, and absorbents. These processes efficiently contain and entrap carbon

dioxide and transport it around the module. The technologies going to be used on the IMMORTAL

include the Carbon Dioxide Removal Assembly (CDRA) [39]. The CDRA on the IMMORTAL

will utilize four beds each with a CO2 sorbent bed. The IMMORTAL will utilize a closed loop

CDRA system which could selectively remove carbon dioxide from the cabin air supply and

reroute it to a carbon dioxide reduction system. There the oxygen will be recovered reducing the

byproducts. The total weight of the CDRA and the Air Revitalization System will be reduced due

to the two person cabin.

The Waste Hygiene Compartment (WHC) is responsible for the collection of waste, its

processing and recovery of water, minerals and oxygen from organic trash and biological waste of

the crew. It also is responsible for odor control and hygiene maintenance in the spacecraft.

Trash must be collected, compacted, stabilized and stored such that it is harmless and at the same

time performs certain useful functions. An example is the case of the waste management aboard

the ISS. The waste collected is ‘wet’ and high in water content and acts as a radiation shield in

addition to acting as a source of water. Most existing waste management technologies use physical, chemical or a combination of both

processes for the treatment of waste and extraction of resources out of it. Some of these include ‘Super

Critical Water Oxidation Method’, pyrolysis and electrochemical incineration among others [40].

Some other technologies that can be considered are using biological agents for the treatment of waste

and extraction of useful products out of them.

24

The purpose of the Active Thermal Control systems is to provide constant heat rejection to

the spacecraft at all points throughout the mission and maintain components at acceptable

temperatures. The ATCS comes into play when the modules’ generated heat exceeds the system

capabilities of the Passive Thermal Control System. To combat heat gain, the ATCS utilizes three

functions: heat collection, transportation, and rejection. The ATCS consist of two Freon - 21

coolant loops, avionics units, liquid-liquid heat exchangers, four heat sink systems for rejecting

excess heat outside the spacecraft, and radiator panels [41]. The coolant loops and radiator panels

will be sized so that they can reject up to 20 kW of heat from the spacecraft, enough for the worst

case thermal loads, when all of the power from the spacecraft is dissipated as heat, in addition to

solar radiation of the craft being at a maximum near perihelion. The Freon 21 coolant loops

transport excess heat from the different modules and deliver them to heat sinks. The waste heat is

removed through cold plates and heat exchangers which are cooled by circulating ammonia. The

system will be based heavily off of ISS heritage which has been proven over the past 15 years.

Like on the ISS, the ATCS consists of a multitude of parts that fall into two categories: The Internal

Active Thermal Control System (IATCS) and the External Active Thermal Control System

(EATCS) [41]. The EATCS is the heat rejection system that transports heat to the outside of the

spacecraft and rejects it to the space environment. The IATCS transports heat around the spacecraft

and transfers it to the EATCS for dissipation into space.

Currently oxygen generators on board the ISS discard the hydrogen gained from hydrolysis

as well as the carbon dioxide produced when the oxygen is consumed. The amount of water

necessary for such procedure is acceptable on the ISS but would not hold for future long duration

space missions such as the IMMORTAL. Electrolysis in conjunction with the Sabatier reaction

can be used to recover water from exhaled carbon dioxide and the hydrogen discarded from

electrolysis. The released hydrogen would then be recycled back into the Sabatier reactor leaving

a deposit of pyrolytic graphite which is easily disposable. The Sabatier reaction is a key step in

reducing the total cost of the IMMORTAL project. By utilizing this in situ resource, weight can

be saved. Using this method, the amount of atmospheric oxygen & nitrogen contained on the

IMMORTAL habitat would total up to about 485 kg.

XII. Human Factors Human factors is a crucial portion of this project as the well-being and safety of the

astronauts are vital for the success of the mission. Since a mission of such length has never been

attempted before, many of the conclusions have been drawn from testing on the ISS as well as on

Earth and many assumptions have been made.

Firstly, the construction and layout of the crew quarters is very important as the astronauts

will spend a large portion of their time in the living quarters relaxing and resting. The crew quarters

will be designed to limit noise that is produced from the rest of the spacecraft as well as the have

the ability to shut out external light to allow the astronauts to have adequate rest at any time of the

day. To reduce external noise for the crew, the spacecraft will be designed such that there is

minimal activity in the vicinity of the crew quarters [42]. Since the duration of the mission is longer

than usual, it is also very important to construct this area in as comfortable and homely manner as

possible. Astronauts are known to be susceptible to psychological effects such as depression due

to the time spent apart from their family. To ensure the astronauts have sufficient protection from

radiation while they are resting, the living quarters will be placed as far to the center of the module

as possible and lined with ultra-high molecular weight polyethene (UMHWPE). On a whole, the

25

crew will have living quarters that are more spacious, equipped with a computer work station,

lighting, storage space with electric bungees and Velcro patches. To ensure sufficient airflow, a

ventilation system consisting of two fans will draw air into the living quarters above the crew and

draw air from the cabin below the crew. The crew quarters should also be well lit and include

power sockets to allow the astronauts use their computer during their time off [43]. The crew will

also bring media entertainment and e-books that they will be able to use whenever they are not

working. This will keep the crew members busy and reduce the opportunities for them to feel

homesick.

One of the factors that has to be considered while planning the mission would be the

amount of provisions that the crew requires throughout the duration of the trip. The recommended

minimum caloric intake is 3000 calories and the meals that are given to the astronauts must be

planned according to such dietary recommendations to ensure the astronauts remain in the best

condition possible. During the duration of the trip, it is calculated that 3,006 meals will be

consumed, assuming that the crew consume 3 meals a day. However, it is encouraged for more

meals to be carried on board. Current meals in space are thermostabilized, rehydratable and have

a shelf life of around 18 months, hence they will be suitable for use during this mission. In addition,

assuming that the crew consumes about 5 drinks per day, about 5,010 dehydrated drink power bags

will be brought along [42]. It is also recommended that a number of supplements, medication and

a set of medical tools be brought on board due to the long duration of the trip, in case of medical

emergencies, and a lack of resources. Common medicine such as that for cold, cough, and muscle

discomfort as well as supplements and vitamins will be included, some of which will enable the

crew to counter the effects of microgravity which will be discussed later.

The long duration of the mission and stresses present within it introduce a psychological

factor that must be taken into account. Despite common beliefs, research has shown that prolonged

isolation does not play a part in the psychological health of the crew but external stresses and

internal stressors do [44]. As the duration of the mission increases, chances of psychological issues

arising would increase. This is important to note as it is crucial during the selection process that

the mission directors not only select crew members who are passionate and well prepared mentally,

but also ensure that they are able to cope and are well prepared in the event that their partner does

develop a psychological issue during the course of the mission. The selected crew should also be

prepared during pre-flight training for unexpected situations such as when communications

between Earth and the spacecraft is lost or experiences a considerable delay. He or she must then

have the presence of mind to analyze and rectify such situations should they occur [45].

A 520-day simulation involving an international six person team of volunteers living inside

a 550 cubic-meter spacecraft-like facility in Russia revealed that long term space travel would

result in alterations to life-sustaining sleep patterns and neurobehavioral consequences. The results

also revealed that the volunteers became more sedentary further into the experiment and majority

experienced some form of disturbance to sleep quality, alertness deficits or disrupted sleep cycles

and times. Hence it is recommended that the living conditions and sleep cycles during the mission

replicate those that are on Earth [46]. As stated above in the section focusing on crew quarters, it

would be beneficial if the crew quarters have systems to completely shut out light to allow the

crew to get uninterrupted and sufficient rest. Currently, up to 50% of the astronauts utilize sleeping

pills and other medication to assist with sleeping. Even then, they still sleep about 2 hours less a

night compared to the 8.5 hours allotted to them, resulting in a sleep deficiency and affecting their

abilities to perform their tasks in space. NASA has even deliberately altered the sleep cycles of the

astronauts prior to launches to ensure that they are at their best condition during the launch [47].

26

To combat sleep deprivation, NASA intends to test out a solid-state lighting module

(SSLM) containing LEDS that produce blue, white and red light made by Boeing with a $11.2m

budget in 2016. The aim of installing the SSLM is to simulate night-day cycle to minimize sleep

disruption. In order to do so, the SSLM will emit blue light to induce production of melanopsin

and suppress melatonin which makes a person feel alert. On the other hand, red light would have

the opposite effect and encourage the feeling of sleepiness. If the tests are successful, such

technology will not only be used for space exploration, but also have the potential to beneficial for

use on people who suffer from sleep loss or insomnia by being able to alteration sleep patterns

[48]. It is also vital that frequent contact with family members are scheduled during the course of

the mission. The contact with their family members would undoubtedly provide the crew members

with a psychological boost during the course of the missions and even more so when they are

facing difficult challenges. Having such human contact and interaction will not only be helpful

during the course of the mission but also help the crew readjust to society after returning to Earth.

Although a space flight of such duration has not been attempted before, applying the above

methods to reduce the psychological effects of the mission will definitely help in reducing mission

risk.

The most important factor is the well-being and health of the crew. This is brought about

by the long duration spent in micro-gravity which is expected to affect the crew’s health negatively

[49]. One of the most common health issues suffered by crew members spending long durations

in space would be the loss of bone mass. This occurs when cells called osteoclasts remove existing

bone tissue when daily mechanical loading falls below normal levels. However, that is not the sole

contributing factor that results in loss of bone mass. Factors such as reduced fluid pressures in legs,

altered nutritional intake and metabolic processing would also result in the loss of bone mass.

Measurements taken from crew members who spent long durations at the MIR space station and

ISS revealed that there was a 1% loss of bone density in the spine, 0.4%-2.7% at the hip and 2.6%

loss of bone fracture strength at the hip for every month spent in space. Similarly, measurements

taken from the same group of crew have also revealed that there was a 2.2% loss of muscle volume,

5.3% loss of peak calf muscle power and 0.7%-4.0% loss of muscle strength per month. By losing

such a significant amount of muscle mass and strength each month, there may be a possibility that

the crew members will no longer be capable of safely carrying out activities that are required to

maintain and repair the vehicle. By the end of the journey, it is estimated that the crew will lose

up to 33% of their fracture strength at the hip bones, 48% of their muscle strength at the knee and

32% of muscle strength at the ankle. Long term durations spent in micro-gravity would also impair

the healing of fractures of the crew. In-flight and ground studies with the use of animals have

demonstrated that microgravity not only caused a delay in the recovery process in abdominal

incision wounds in rats, but also produced an inflammatory response and scars. Since mechanical

loading is known to be vital to healing in bones, it is suggested that the reduced gravity has played

a role in the prolonged time taken for wounds to heal. As such, this factor must be taken into

account when considering the quantity of medical supplies that should be brought on board the

spacecraft.

Dietary measures and pharmaceuticals are viable and cost effective options to combat the

loss of bone and muscle mass in space. For example, a crew member who consumed vitamin K

supplements halfway through a 6 month Mir mission produced an increase in osteocalcin and

alkaline phosphatase, which are indicators of new bone formation, during the time in which she

consumed the supplements. Current pharmaceuticals that are used to increase bone formation such

as Anabolics may also be suitable for trial runs at the ISS in the near future these. The use of

27

Angiotensin-converting enzymes can be further researched into in the near future as a 3 year trial

on elderly subjects have shown to be able to slow the decline in muscle strength. Another possible

method of reducing bone and muscle loss would be through inhibiting the production of myostatin

in skeletal muscle cells. A study done on mice on a 13-day shuttle mission showed that the

inhibition of myostatin prevented muscle and bone loss as well as enhanced fracture healing [50].

However, in all these pharmaceuticals, there has been limited or no testing in the context of

spaceflight and thus would still need to go through rigorous testing and trials to find out before

they are suitable for use during spaceflight.

A countermeasure currently being used is the Advanced Resistive Exercise Device

(ARED). This is a device that stimulates free-weight exercises in normal gravity and used to

maintain muscle strength and mass during the time they spend in space [51]. However a similar

idea being mooted is the combination of using whole-body vibration with resistive exercise to

enhance the effectiveness in combating muscle loss when using the ARED. Studies have shown

that a vibrating foot plate did prevent or reduce bone loss in postmenopausal women and disabled

children and reduced invertebral expansion in a 90-day bed study [50]However, similar to the use

of pharmaceuticals, the use of body vibration in microgravity is still unknown as of yet and further

study and research, possibly in the ISS, will be the only way to understand.

XIII. Radiation Protection Trapped electrons, and protons (<10 MeV), protons and light charged particles (>10 MeV),

Galactic Cosmic Radiation (GCR) and secondary photons, secondary charged particles, and

neutrons are the radiation types that need to be considered for this mission, as specified by the

International Commission for Radiation Protection (ICRP) [52]. The shielding material making up

the exterior of the spacecraft and EVA suit will be able to stop trapped protons (<10 MeV). Other

types of radiation may penetrate the EVA suit and spacecraft, significantly increasing risk to both

the health of the astronauts and the mission success. To meet this risk advanced shielding material

concepts, rigorous radiation and dose monitoring and dose optimization measures will be

integrated into the IMMORTAL mission.

GCR is modulated by the solar cycle. This occurs due to some of the GCR being deflected

by the solar wind. GCR presents a greater challenge versus solar particle emissions for this mission

due to the 2018 launch date coinciding with a solar minimum. Thus, this requires shielding

primarily against the high energy particles of GCR. GCR has a greater potential to produce harmful

secondary radiation when interacting with the spacecraft shielding compared with a Solar Particle

Event (SPE). This is supported by the results from high charge and energy transport software

(HZETRN) simulations done by Boeing Research [53]. These simulations suggest that

implementing a mission during a solar maximum will mitigate overall dose compared with a solar

minimum. However, due to launch window constraints the IMMORTAL mission must launch

during a solar minimum and will instead optimize against harmful GCR as best as possible.

A “defense in depth” approach will be applied to the passive shielding in the exterior and

interior of the spacecraft to mitigate the risks of harmful secondary radiation. Hydrogen rich

plastics (Polyethylene) will be used in multiple layers for the exterior of the spacecraft. Ideally,

thickness would approach 1.55 cm to optimize radiation protection. However, due to weight

constraints these thickness will be 1.05 cm per layer meeting NASA radiation protection

guidelines. A water wall comprised of multiple polyethylene tanks large enough to accommodate

an areal density of 20 g/cm2 of water will surround the sleeping quarters. This chamber will be an

28

effective shield in the event of a severe SPE according to the NASA Spaceflight Radiation Health

program [54]. Additionally, both polyethylene and water protect against GCR’s and other high-

energy particles by means of nuclear fragmentation. Incoming extra-solar particles traveling at

near-relativistic speeds impact the nuclei of each respective material and break apart into

secondary radiation in the form of much lower-energy particles. The total combined energy of the

resultant series of particles will always result in a lower dose being absorbed by the astronaut.

Polyethylene and water are both especially effective in this reaction due to each having a high

fraction of hydrogen by weight, approximately 14.4% and 11.2% respectively. Hydrogen has been

shown to cause the greatest fragmentation effect per unit mass over any other element, making

polyethylene and water highly useful in maximizing shielding while still maintaining a low mass.

Additionally, the low cost and availability of these materials makes them both a practical and cost-

effective choice. Concentration of these protective layers specifically around crew

quarters/sleeping area further reduces the cost, as total protection in all areas of the spacecraft is

both unnecessary and impractical due to structural constraints. Ideally the crew will spend almost

all of their non-operational time within this zone to minimize absorbed dosage.

Active and passive measures will be implemented to optimize radiation protection in this

mission. EVA time will be limited as much as possible throughout the mission. Also, when sensors

on the spacecraft detect spikes in radiation from either GCR or SPEs the crew will stop non-

essential operations and retreat to the water wall chamber until radiation levels return to baseline.

Consumables will also be used to counteract oxidative effects of radiation-induced free radical

species from prolonged exposure. Astronauts will consume compounds containing “α-tocopherol

(vitamin E), a known in-vivo antioxidant which protects cells and tissues from radical effects; 2)

ascorbic acid (vitamin C), a terminal biological antioxidant compound; and 3) β-carotene (vitamin

A precursor), a free radical quenching molecule [55].”

Due to the variety of radiation types the astronauts will experience during this mission,

multiple techniques will be employed for dose and radiation detection. These include personal

dosimeters that will be read-out on a daily basis, Radiation Area Monitors (RAM), Tissue

Equivalent Proportional Counters (TEPC), and Charged Particle Directional Spectrometers

(CPDS). There is operational experience with these techniques from their usage on the ISS. Due

to limitations in neutron detection, these devices may provide outputs not representative of the

dose received by astronauts. Furthermore, personal dosimeters only provide absorbed skin doses.

To mitigate these shortcomings biological dosimetry will be necessary in order to accurately

measure the dose as well as chromosomal damage occurring in cells. The Fluorescence in situ

hybridization (FISH) chromosome painting technique will be used to track the number of

aberrations taking place within blood lymphocytes on a monthly basis. These aberrations are

related to the damage taking places in cells [56].

Radiation exposure will meet standards adopted by NASA as a result of recommendation

by the National Council on Radiation Protection and Measurements (NCRP). These standards

serve as a basis for the implementation of several requirements that must necessarily be met to

ensure that the health and safety of those involved is not compromised. These requirements

include, but are not limited to: 1) use of techniques applies to a limited population, primarily small

astronaut crews, 2) compilation and maintenance of detailed flight crew exposure records, 3) pre-

flight hazard assessment, 4) planned exposures be kept as low as reasonably achievable, 5)

maintenance of NASA operational procedures and flight rules to minimize the chance of excessive

exposure. In addition to these requirements, specific limits on radiation exposure per organ have

29

been established in order to prevent potential illnesses resulting from said exposure, they are as

listed below.

Table XIII-1 Organ Specific Exposure Limits [57]

Exposure Interval Depth (5cm) Eye (0.3cm) Skin(0.01cm)

30 Days 25 REM 100 REM 150 REM

Annual 50 REM 200 REM 300 REM

Career 100-400 REM 400 REM 600 REM

Table XIII-2: Career Exposure Limits by Age and Gender [56]

Age

Gender 25 35 45 55

Male 150 REM 250 REM 325 REM 400 REM

Female 100 REM 175 REM 250 REM 300 REM

Monthly and annual limits given in the above tables exist in order to prevent the short-term

physiological side effects from moderate levels of exposure. Career limits are primarily in place

to contain the excess risk of future cancer mortality due to radiation exposure to around 3%.

Exposure limits vary by age and gender as shown by the above tables. Approximate dosages are

given in units of REM, which correspond to the SI biological equivalent dose unit of 1 Sievert =

100 REM.

XIV. Scientific Return Due to mass and size limitations, IMMORTAL will not involve a “science payload” in the

traditional sense. Rather, the scientific contributions of the mission will revolve around data

collection applicable to future, longer duration missions. The experimental design will be based

upon a philosophy of maximizing relevant data collected with minimal equipment. To the highest

practical degree, the information collected will be transmitted back to a scientific team on Earth

and the crew will have only a preliminary part in interpreting the data.

In order to maximize the benefit of the flight from Earth to Mars, radiation levels both

inside and outside the spacecraft will be recorded and transmitted back to Earth. Little to no

analysis of this data will be done in flight, in order to avoid unnecessary computing power and

time. While similar data has been collected in the past, the radiation dangers beyond Earth orbit

are still one of the major obstacles to long duration space travel. Further analysis of the radiation

encountered as well as the effectiveness of the chosen radiation mitigation techniques could prove

extremely helpful for the planning of future deep-space long-duration missions, all for the low

mass and power requirements of a few radiation sensors. Additionally, the growth of a variety of

small plants will be analyzed on the approach and return flights, hopefully yielding further

information on the feasibility of growing at least a portion of the required food on a longer mission.

Upon arrival, some imaging will be done of both Mars and its moons during the flyby, but these

30

images are not expected to reveal any new, scientifically relevant data, given the extensive imaging

already taken by the Mars Reconnaissance Orbiter. Barring any truly unexpected phenomenon,

planetary and space science payoff is anticipated to be minimal, but as that is expected given the

mission architecture and purpose. The primary scientific payoff of this mission will involve the previously unstudied effects of

long duration space travel on human physiology. The astronauts will wear clothing with biometric

sensors embedded at all times [58]. Their heart rate, temperature, blood pressure, and oxygen saturation

level will transmitted to a medical team on Earth in as close to real time as possible. This system of

constant monitoring will not only provide an early warning for any possible health issues, but will also

provide a way to track any slowly developing or previously unnoticed effects of long duration space

travel. Regular stethoscopes, sphygmomanometers, etc. will also be included with the medical

supplies as back-ups. In addition to the basic monitoring, a handheld ultrasound device, such as

those currently commercially available, will also be used for semi-regular exams, again with the

dual purpose of ensuring crew health and ascertaining possible risks involved with extended time

in space [59]. The data from the scans will be included with the other data transmissions back to

Earth. Blood samples will be taken and frozen every month for analysis upon return. The

diagnostic equipment required to analyze the blood samples on board would be too heavy to take,

so even though the samples might not be as helpful after such a long storage period, any form of

degradation of the samples should be minimal, and they will hopefully rule out any major

problems. Exhaustive pre and post-launch physicals will provide points of comparison for all the

information collected.

The psychological effects of long duration space travel will also be studied as well. While

not the most sophisticated diagnostic tool, basic personality tests will be administered pre and post

launch, as well as every month or so throughout the mission. Consistency in these tests across the

mission would indicate that there is no significant psychological danger to space travel, however

a shift could indicate an unanticipated risk factor that would require further study before a longer

duration mission is attempted. Also, basic IQ, visual, and motor skills test would performed at the

same intervals. The effects of long duration space travel and confinement on basic skills and

processing ability has not been studied in such extreme circumstances and any degradation would

pose a very obvious danger for long duration missions [60]. Every two-three weeks, the crew

would be required to have a private video chat (or as close to a video chat as possible with the time

delay) with a mission psychologist back on Earth. While these sessions will be recorded, they

would be kept private from the other crew member to allow free communication [60]. The

psychologist would be focused on their overall mental health and attempting to pick up on any

subtler patterns that might indicate a psychological danger in space travel. In a very direct sense

the primary science payload of IMMORTAL is the crew itself. The mechanics and technology

required for space travel can be developed and tested sufficiently on Earth; however, true human

reactions and responses can only be studied by actually sending a crew into space, and thus the

science payoff of this mission is much more important than would originally be thought.

XV. Cost The IMMORTAL mission architecture focuses on using existing technologies to allow this

relatively near term mission to succeed without excessive rushed development costs. However,

some technologies still do need proving such as the modifications to the Cygnus spacecraft, the

31

testing of Dragon’s heat shield at high re-entry velocities and the development of a reliable long

duration ECLSS system.

Costs for this mission were calculated using published costs in the case of existing

hardware, or were estimated using NASA’s Project Cost Estimating Capability (PCEC)

framework. This framework is an excel based architecture that uses cost estimating relationships

between past missions to create reasonable A healthy margin of 20% was added on top of all

estimates in order to provide for cost overruns and changes in mission design throughout the life

cycle.

The final total costs for the mission come to $242M for development and, $172M for

systems test hardware, $736M for the manufacturing of the actual flight units, and $342M for

supporting ground operations during the mission. A more detailed breakdown of the costing can

be found in table XX. The total programmatic cost of the IMMORTAL mission is approximately

$1,493M. While this cost is more than 50% greater than the architecture proposed by the

Inspiration Mars Foundation, it is a more realistic estimate based on technologies that should

actually be available within the desired timeframe.

Table XV-1: Cost Analysis (All values in $MM USD)

Item DD STH FU Total:

Falcon Heavy (1 Launch) 0 0 135 135

Delta IV Heavy (1 Launch) 0 0 300 300

Dragon 0 0 76.83 76.83

Cygnus Modified Craft 80 115 160 355

4 Meter DCSS 20 35 50 105

Heat Shield Verification 7.80 5.39 4.15 17.35

Thermal 7.69 6 0.46 14.15

CCDH 0.18 0.52 0.40 1.114

Batteries 0.11 1.43 1.102 2.65

Solar Array 0.36 0.74 0.57 1.671

Communications 29.09 3.786 2.912 35.79

ECLSS 80.29 0.205 0.15 80.66

Attitude Control 2.49 1.788 1.37 5.657

Reaction Control 0.056 0.531 0.408 0.99

Crew Accommodations 2.46 0.038 0.029 2.53

Integration of Dragon with Cygnus 12 1.6 2.5 16.1

Table XV-2: Cost Summary

Total Design and Development 242.56

Total Standard Test Hardware 172.04

Total Flight Unit 735.90

Ground Operations 342.22

Overall Mission Cost 1492.72

32

XVI. Risk Due to a truncated timeframe before launch in addition to the elevated implications of

human factors involved in the mission, risk assessment and mitigation was given high

consideration in order to ensure mission success, and more importantly, the safe return of the two-

person crew on-board the Cygnus-Dragon vehicle.

The associated risks for each subsystem were identified and assessed according to the

NASA probability-impact strategy as described below in Table X-1 [1]. The probability of

occurrence and impact of each risk is assessed with a numerical value from 1-5. Using the ratings

for probability and impact, risks were categorized in a risk matrix (Table X-2) to determine the

priority of mitigation associated with each potential risk. Risk items in red should be prioritized in

terms of risk management and mitigation while those in green pose little to no threat in the overall

success of the mission. Yellow items have a moderate priority and for each risk, a mitigation

strategy is proposed for implementation.

The risk assessment revealed two items of high priority: the trans-Mars injection burn and

the heat shield for Earth atmospheric reentry. The TMI burn is an integral part of the Earth to Mars

orbital maneuver. During the procedure, there is a risk of complications in the maneuver that could

potentially utilize more fuel than intended. As it has never been done with humans before, the

design and development portion of the mission will seek to mitigate risks with attention to a strong

design for TMI. Additionally, the transfer vehicle will be loaded with extra fuel to offset any

potential complications.

Another item of concern is the Dragon heat shield for Earth atmospheric reentry. To

mitigate any risks of failure, there will be stringent prelaunch testing and verification. As the shield

is thicker than other transfer vehicles, there is a lower chance of failure that will occur.

Table XVI-1: Probability-Impact Scale

Number 1 2 3 4 5

Probability

Improbable Low Occasional Likely Definite

Rare, very low

chance of

occurring

Less than likely to

not occur

May or may not occur,

medium chance of

occurring

Occurrence

assumed to be

high, more than

likely to occur

Assumed to occur

in every execution

Impact

Insignificant Minor Moderate Major Catastrophic

Little to no

impact

Possible

functionality loss

but minor, could be

isolated or restored

Medium impact,

system still functional

as a whole with loss of

some functionality

Portions of systems

rendered non-

functional, mission

objective lost

Total system and

mission failure,

can lead to

fatality

33

Table XVI-2: Risk Matrix

Pro

ba

bil

ity

5 Very High

4 High

3 Moderate

2 Low

1 Very Low

Very Low Low Moderate High Very High

1 2 3 4 5

Impact

Table XVI-3: Risk Analysis for Launch/Deployment Systems

Component Risk Consequence Mitigation Probability Impact

Level

Rocket Launch Launch window

missed

Non optimal orbital trajectory with

added mission cost, potential

jeopardization

Stringent launch schedule and preparation with built-in buffer time

2 4

Engine Ignition failure Launch vehicle does not lift-off, launch scrubbed

Testing and regular checks prior to launch

1 1

Engine/Fuel Unstable engine burn,

explosion Potential fatality and loss of mission critical components

Use of reliable launch

systems/vehicles and regular check-ups and testing

2 5

First Stage Stage separation

failure Complications with upper stage ignition

Redundant pyrotechnic separation

charges, proper system wiring and

ground testing 1 2

First Stage Premature separation Unstable flight, potential inability to

reach orbit

Sections inspected for secure connection, wiring and electrical

harnessing inspected and tested 1 2

Avionics Electrical/Avionics

failure

Complications in flight guidance and

flight path

Extensive ground testing and system

redundancies 1 3

Dragon/Cygnus Docking

Maneuver fails Crew unable to reach habitat, fatal

Thorough testing of release

hardware, extensive preparatory

simulations 1 5

DCSS Docking Maneuver fails Crew unable to perform burn, mission failure, non-fatal

Thorough testing and simulation training to ensure success

2 5

Propellant Boil-off during

assembly

If crew loiters too long, insufficient

propellant to perform trans-Mars

injection due to propellant boil-off

Effective and efficient procedure to

ensure quick execution of TMI

maneuver 2 4

TMI Burn Inaccurate trajectory Unable to have enough fuel to make

return home

Extra fuel and strong pre-maneuver

plan 3 5

34

Table XVI-4: Risk Analysis for Power/Thermal Systems

Component Risk Consequence Mitigation Probability Impact

Level

Solar Arrays Deployment malfunction

Unable to power spacecraft for duration of mission

Rigorous testing before launch, robust design concepts

2 5

Solar Arrays Power collection

failure

Unable to power spacecraft for duration of mission or reduced power

collection

Secondary battery system to store

backup energy, spacecraft ability to

defer to emergency power saving mode

2 4

Solar Arrays Damage from high-

velocity debris

Unable to power spacecraft for

duration of mission or reduced power

collection

Secondary battery system to store

backup energy, spacecraft ability to defer to emergency power saving

mode

1 4

Batteries Battery failure Unable to store excess power

collected from solar array

Usage of reliable Li-Po batteries as

well as primary solar array system 1 2

Batteries Battery

overheating/explosion

Loss of batteries and potential

damage to housing compartments

Usage of reliable Li-Po batteries and

preflight testing procedures 1 1

Electrical Wiring Disconnection,

shorting, damaging

Loss of various on-board systems,

potentially fatal

Circuit redundancies and ground

testing 2 4

Fluid Loops Leakage Loss of fluid pressure and potential damage to sensitive spacecraft

components

Preflight testing and robust design, tools available on-board for en route

repairs 4 2

Fluid Pump Malfunction

Unable to run ATCS, must be

repaired quickly to ensure proper heat rejection

Redundant pump systems and tools

available on-board for en route repairs

3 3

Table XVI-5: Risk Analysis for ECLSS/Human Factors Systems

Component Risk Consequence Mitigation Probability Impact

Level

Water Recover

Loop Leakage

Loss of fluid pressure and complications in water recycling and

recirculation

Preflight testing and robust design, tools available on-board for en route

repairs

2 3

Air Revitalization

Element Failure

Inability to regulate CO, CO2, and

relative humidity levels and provide clean oxygen to crew

Multiple design redundancies and

training to fix any issues on-board 2 5

Waste Processing

System Failure Inability to manage solid waste

Multiple design redundancies and

training to fix any issues on-board 2 1

Airlock System Failure/Leakage Potentially deadly if unable to ensure airtight seal for crew

Stringent testing and design before flying mission

1 5

Radiation

Exposure

Excessive radiation

dosage Assorted adverse health effects

Well-designed radiation shielding in

crew habitat, with water,

Polyethylene, and metals, and dosimeter detectors

1 3

Physical Health Illness

Impaired functioning of human crew,

potential spread and contamination throughout capsule

Vaccinations, pre-mission health

exams, included medical equipment

and relevant training, and robust air & water filtration to eliminate

pathogens

2 1

Mental/Emotional

Health

Depression/Morale

Issues Performance of crew hampered

Rigorous crew selection criteria, routine fitness and health regimens,

human interaction 3 1

35

Table XVI-6: Risk Analysis for Avionics, Controls, and Navigation Systems

Component Risk Consequence Mitigation Probability Impact

Level

Startrackers/Sun

Sensor Malfunction

Unable to determine spacecraft

attitude

Redundancies of devices on spacecraft (IMU and dual SS

Trackers)

2 4

On-board

Computer Malfunction

Unable to compute navigation

commands for spacecraft guidance Redundancies in computer systems 1 5

Antennae Failure Unable to carry communications

between spacecraft & Earth

Redundancies through multiple

antennae on spacecraft 1 3

Aerojet Thrusters Failure Unable to fire for course and attitude

corrections

Rigorous on ground testing and

usage of space proven thrusters in different areas on vehicle

2 4

Transponder Malfunction Cannot transform radio frequencies

to send out

Redundancy transponder with

prelaunch testing 1 2

Hydrazine

Propellant Leakage/Boil Off

Loss of total impulse able to be

imparted on vehicle

Preflight checks for securing the

propellant 2 3

Table XVI-7: Risk Analysis for Spacecraft Structure

Component Risk Consequence Mitigation Probability Impact

Level

Cygnus Habitat Damage from high-

velocity debris Hull damage, potential breach Installation of MMODs 2 3

MLI Damage from high-

velocity debris

Thermal leaking around damaged

area Installation of MMODs 2 2

Heat Shield Performance failure Unable to re-enter Earth atmosphere,

fatal

Prelaunch inspection, prior success with Dragon capsule with thicker

heat shield

3 5

Propellant Tanks Leakage/Failure Loss of valuable propellant

Prelaunch inspection, ensuring seals

and structural integrity of

component

3 4

Momentum Disks Mechanical failure Unable to accurately actuate

spacecraft attitude

Redundant sets of 3-axis momentum

disks 1 1

36

XVII. References

[1] J. Chilton, "A New Framework for Human Space Exploration," October 18, 2012. [Online].

Available: https://info.aiaa.org/Regions/SE/HSV_AIAA/Downloadable%20Items/AIAA-

Chilton_18Oct2012_Final2.pdf.

[2] S. J. H. a. D. J. Creech, "SLS Dual Use Upper Stage (DUUS) Opportunities," 2013. [Online].

Available: http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20130013953.pdf.

[3] "Falcon Heavy," Space Exploration Technologies, 2014. [Online]. Available:

http://www.spacex.com/falcon-heavy.

[4] United Launch Alliance, "Delta IV Launch Services User’s Guide," June 2013. [Online].

[5] D. A. e. a. Tito, "Feasibility Analysis for a Manned Mars Free-Return Mission in 2018," IEEE,

2013.

[6] Jet Propulsion Laboratory, "Mars Reconnaissance Orbiter: Trajectory Correction Maneuvers,"

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XVIII. Appendix A: Mass Budget Subsystem Mass

Command and Data Handling 45

Attitude Determination and Control Hardware 120

RCS Propellant 210

Communication 105

Food 1250

Water 1225

Water Recycling + Spares 675

Air 485

Air Recycling + Spares 850

Thermal + Spares 1220

Crew Accommodations/Interior Structures 265

Crew Personal Effects 30

Dragon Capsule 6000

Cygnus Super Enhanced PCM and Service Module Bus 2300

Solar Arrays 325

Batteries 120

Integration of Dragon with Cygnus 150

Payload Attachment Fitting 350

Crew Person Mass 150

Total 15875

ID Task Name Duration Start Finish

1 Program Definition 32 days 5/15/2014 6/14/2014

2 Core Mission Design 120 days 6/26/2014 10/19/2014

3 Development 240 days 9/14/2014 5/3/2015

4 Testing 450 days 2/18/2015 4/27/2016

5 Construction/Manufacturing 650 days 4/13/2015 12/30/2016

6 Integration/Assembly 440 days 4/11/2016 6/9/2017

7 Launch Preparation & Check 150 days 8/2/2017 12/24/2017

8 Launch Window 10 days 12/24/2017 1/2/2018

9 In-Orbit Vehicle Assembly (Predicted) 0 days 1/1/2018 1/1/2018

10 Trans-Mars Injection Burn (Predicted) 0 days 1/1/2018 1/1/2018

11 Earth To Mars Transit 227 days 1/1/2018 8/8/2018

12 Mars Flyby 1 day 11/13/2018 11/13/2018

13 Mars To Earth Transit 274 days 11/13/2018 8/4/2019

14 Dragon Earth Reentry 1 day 11/29/2019 11/29/2019

15 Splashdown 0 days 11/29/2019 11/29/2019

16 Capsule & Crew Recovery 0 days 11/29/2019 11/29/2019

1/1

1/1

11/29

11/29

Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4

2013 2014 2015 2016 2017 2018 2019 2020

Task

Split

Milestone

Summary

Project Summary

Inactive Task

Inactive Milestone

Inactive Summary

Manual Task

Duration-only

Manual Summary Rollup

Manual Summary

Start-only

Finish-only

External Tasks

External Milestone

Deadline

Progress

Manual Progress

Page 1

Project: IMGantChart.mpp

Date: 3/15/2014