Conference PaperPDF Available

Critical Analysis and Review of Current Mars Mission Scenarios for SpaceX Starship

September 2022

Conference: 73rd International Astronautical Congress
At: Paris, France

Authors:

Bjarne Westphal

Simtec Systems GmbH

Volker Maiwald

German Aerospace Center (DLR)

Human space exploration is currently aiming at the lunar environment in the frame of the ARTEMIS program, including the Lunar Orbital Platform Gateway and also lunar landings. In addition, there are mission plans by most prominently SpaceX for establishing a human exploration of Mars utilizing SpaceX Starship, a two-stage heavy launch vehicle and spacecraft for transfer to and landing on Mars. The currently discussed scenario includes landing on Mars, setting up of in-situ resource utilization (ISRU) for propellant generation and resupply of Starship for the return to Earth. Such a mission would not only be a huge step forward in human development, but would also require technological advances beyond what is currently possible. This paper analyses the currently available information about SpaceX mission plans for Mars based on Starship, extrapolates requirements, necessary technology developments and based on key figures evaluates the feasibility of these mission plans. Key figures are launch mass, payload mass and unloaded mass, technology readiness and costs. It is shown that two major parts of the mission scenarios, i.e. power supply and ISRU propellant production have low technology readiness, which is driving costs, mass and volume and timeframes expected to close technological gaps are not fitting SpaceX mission plans. System elements which require smaller technological advances, but are still critical include power supply for Starship during transfer and elevator technology to reach the ground after landing. Overall, the analysis shows that current plans are not feasible and therefore recommendations are made to achieve feasibility for Mars missions using Starship.

Stuffed Whipple Shield for Starship with two bumper shields (BS) and one back wall (BW), after [14]

…

Deployment mechanism of the MegaFlex solar array [45]

…

Process for propellant production on Mars, from harvested water and collected CO2, after [43]

…

10 kW (left) & 650 kW (right) FSP System [33]

…

MOROCAS concept with different modules [36]

…

Figures - uploaded by Volker Maiwald

Content may be subject to copyright.

Content uploaded by Volker Maiwald

Content may be subject to copyright.

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IAC-22-A5.2.3 Page 1 of 16

IAC-22-A5.2.3

Critical Analysis and Review of Current Mars Mission Scenarios for SpaceX Starship

Bjarne Westphala,*, Volker Maiwaldb

a Technical University of Braunschweig, Braunschweig, Germany, bjarne@westphalkiel.de

b German Aerospace Center (DLR), Institute of Space Systems, Department of System Analysis Space Segment,

Bremen, Germany, volker.maiwald@dlr.de

* Corresponding Author

Abstract

Human space exploration is currently aiming at the lunar environment in the frame of the ARTEMIS program,

including the Lunar Orbital Platform Gateway and also lunar landings. In addition, there are mission plans by most

prominently SpaceX for establishing a human exploration of Mars utilizing SpaceX Starship, a two-stage heavy launch

vehicle and spacecraft for transfer to and landing on Mars. The currently discussed scenario includes landing on Mars,

setting up of in-situ resource utilization (ISRU) for propellant generation and resupply of Starship for the return to

Earth. Such a mission would not only be a huge step forward in human development, but would also require

technological advances beyond what is currently possible. This paper analyses the currently available information

about SpaceX mission plans for Mars based on Starship, extrapolates requirements, necessary technology

developments and based on key figures evaluates the feasibility of these mission plans. Key figures are launch mass,

payload mass and unloaded mass, technology readiness and costs. It is shown that two major parts of the mission

scenarios, i.e. power supply and ISRU propellant production have low technology readiness, which is driving costs,

mass and volume and timeframes expected to close technological gaps are not fitting SpaceX mission plans. System

elements which require smaller technological advances, but are still critical include power supply for Starship during

transfer and elevator technology to reach the ground after landing. Overall, the analysis shows that current plans are

not feasible and therefore recommendations are made to achieve feasibility for Mars missions using Starship.

Keywords: Feasibility analysis, SpaceX Starship, Mars mission, human spaceflight

1. Introduction

SpaceX’s Starship will bring the first humans to Mars,

according to Elon Musk’s vision [1]. This achievement

would pale any human spaceflight mission that has

occurred in the past six decades and is one of several

plans for re-introducing the space environment beyond

Low Earth Orbit (LEO) into humanity’s theatre of

activity. This leap forward in humanity’s capabilities is

ambitious, considering the fact that NASA’s ARTEMIS

program is still in its infancy as are all other plans for

leaving LEO.

The mission to Mars is associated with several

challenges, which have to be addressed. They range from

human physiology and the health risks of spending

prolonged time in a low-gravity environment to funding

challenges and technical obstacles, e.g. a life-support

system which can operated reliably for several years

during a Mars mission – where replacement of parts,

beyond what you took along for the ride, is not possible.

Mission scenarios for SpaceX’s Starship involve

refuelling on Mars with fuel obtained, resp. produced on

the Martian surface [2]. This is one of the major

challenges as the fuel has to be produced reliably if not

for the crew to be stranded on Mars without a way home.

The technology has not only to be reliable, but this huge

infrastructure, larger than anything NASA brought to the

Moon during the course of the Apollo-program, has to be

transported to Mars.

These are just some challenges associated with

SpaceX’s Starship mission to Mars. This paper

investigates the feasibility of Starship to be applied for

the proposed Mars mission. For this purpose, first the

currently available information is compiled and then

weighed against the necessities of a Mars mission as

given by the current mission scenario [2]. Data is

extrapolated where necessary, based on existing

technology of other entities, to paint an as complete

picture as possible, involving in-situ-resource-utilization

(ISRU) technologies for generating fuel on the Martian

surface or nuclear reactors for power generation.

Subsequently, the feasibility is analysed and

discussed e.g. considering the technology readiness level

of the required technologies, available payload mass or

and thus propellant mass required.

2. Method

To review the feasibility of Starship’s Mars mission as

proposed by SpaceX, all relevant data for the spacecraft

and mission components were compiled. This data was

obtained from publications by SpaceX (e.g. [3] [4] [2])

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IAC-22-A5.2.3 Page 2 of 16

or about SpaceX (e.g. [5] [6] [7]) were the former were

not available. In case of contradicting information, the

most recent one was selected, to consider possible

updates on the design. Where no information was

available about Starship, data was extrapolated from

existing systems, e.g. based on ISS technology. The

system design also includes ISRU-technology.

Since the topic of Starship is still new, the search was

conducted purely via digital sources. Information also

comes, for example, from videos in which Elon Musk is

interviewed, in which he shows and explains the progress

of Starbase, as well as from presentations he has given.

In the search for further components and technologies for

the Starship and ISRU, NASA and other space

companies were frequently consulted for existing ones

and those currently in development.

As a backdrop for the analysis a mission scenario has

been formulated derived from information supplied by

SpaceX, to evaluate how the above design is fitting that

mission scenario. In addition, further requirements have

been set for the Starship and the ISRU components that

still have to be fulfilled. A possible launch window and

trajectory including flight duration and the required

speed difference were also selected (see Table 2).

With both in mind a system design has been set up as

a compilation of information given about SpaceX

Starship. The following steps were taken:

1) definition of the subsystems,

2) set-up of the respective system designs and

requirements,

3) estimation of mass and power budgets where

possible

With this compiled system design, the mission feasibility

concerning the given mission scenario has been analysed

and evaluated subsequently. For this feasibility analysis

the most relevant key figures have been identified, which

can be addressed with the available information. These

key figures have been:

a. Launch mass

b. Unloaded mass (mass w/o crew supporting

equipment)

c. Payload mass (mass w/ crew supporting

equipment)

d. Technology Readiness Level

e. Costs

Figure 1: Mission scenario as described by SpaceX. First (1) the crewed Starship is transferred into orbit, where it

separates into booster and upper stage (the actual Starship). The booster returns to Earth (3) and uncrewed transports

launch into Earth orbit (4). There, the transports refuel the crewed Starship (5), before landing back on Earth. Once

refuelled Starship makes the interplanetary transfer to Mars (6), where it conducts aerobreaking (7) and lands (8). ISRU

is used to refuel Starship (9). Afterwards Starship can launch from Mars (10), transfers back to Earth (11), where it

uses aerobreaking once more (12) to slow down and finally land (13). Source: [2], Mars and Earth images: NASA,

public domain, overall image: own

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IAC-22-A5.2.3 Page 3 of 16

3. Baseline Mission Scenario

Figure 1 shows the mission scenario compiled from

information published by SpaceX for planned Mars

missions using SpaceX Starship [2]. It relies on ISRU for

generating fuel on Mars, which is a major element to be

regarded for feasibility. Another major aspect is

aerobreaking capability.

According to current planning, the first two uncrewed

cargo Starships could use the next but one launch

window in 2024 to make their first flight to Mars [8]. In

the following launch window, two more cargo and two

crewed Starships (10 to 20 people onboard) are set to be

launched [9, p. 4]. For this work, it is assumed that the

two crewed Starships will each have a crew of ten on the

first mission and that a flight trajectory with a longer stay

of 368 days on Mars [10] will be chosen, as the propellant

for the return flight of both crewed Starships must be

produced during this time.

According to mission plans [2], the crewed Starships

will launch, (1) in Figure 1, from Boca Chica, Texas

and/or from launch site LC-39A at the Kennedy Space

Center (KSC) in Florida [11, p. 7 ff.]. The crewed

Starship will remain on orbit (2), while the booster will

return to Earth (3). Subsequently, transport variants of

Starship will launch into orbit (4), where they refuel the

crewed Starship and their booster will once more return

to Earth. Once refuelled the crewed Starship will transfer

to Mars (6), where it will use aerobreaking (7) to remove

excess energy and land on the Martian surface (8).

During the mission stay ISRU will be used to generate

fuel and refuel Starship (9) for the return trip. Once the

mission is over (and sufficient fuel has been generated)

Starship will leave Mars (10) and go on its return trip to

Earth (11). Again, using aerobreaking (12) Starship will

eventually land on Earth and end the mission (13). [2]

4. Compiled System Design

The assumptions and designs made in this Chapter refer

to the crewed version of Starship, because this is the

version with the highest launch mass on the return flight

from Mars. The cargo Starships would have a mass of

100 t less on a return flight (not on the first missions), as

their payload would remain on Mars.

SpaceX’s only design specifications for subsystems

are for the main engines and tanks, and the stainless-steel

structure. Therefore, the subsystem designs are mainly

based on own assumptions made in this work or on

existing systems such as the ISS or the Orion space

capsule.

4.1 Starship

Structures

The structures subsystem comprises all structural

elements, including protection against cosmic and solar

radiation. To protect the most important areas, such as the

crew’s sleeping compartments and the control centre,

these should be covered with polyethylene, as this is a

well protective material [12]. For additional protection,

water pipes, which are used to supply the crew and

transport waste water, are to be laid in the spacecraft in a

way that encloses as much habitable space as possible, as

water is a well protective substance as well [12]. Since

the mass of the radiation protection should be kept as low

as necessary, materials that have to be onboard anyway

should also be used as additional protection. These

include for example equipment and food. Thus, in the

event of a strong solar flare, a protective shelter could be

built in which, in addition to covered walls made of

polyethylene, the crew can surround themselves with

food and equipment containers and wait it out. This

increases the density of material around the crew,

resulting in better protection. This process is also being

pursued for the Orion capsule [13].

For micro-meteoroid protection, Starship, similar to

the Columbus module of the ISS, is to have a protective

layer reinforced with Kevlar and Nextel, a so-called

Stuffed Whipple Shield (SWS), which bursts incoming

objects with three layers of protective material and thus

prevents them from penetrating [14]. The three layers

should consist of two bumper shields (BS) and the back

wall (BW), as shown in Figure 2.

Furthermore, Starship must be designed and built in

such a way that its structure can carry the payload of up

to 100 t with empty tanks, because they will be almost

empty by the time it arrives on Mars. As with the current

prototypes, 3 mm thick 304L stainless steel is assumed to

be used for Starship’s outer skin [15].

Environment Control and Life Support System

For Starship, the ECLSS is should be modelled after that

of the ISS. For additional protection against strong solar

storms, special vests are to be available onboard Starship,

which should be worn when a solar flare occurs. One

such vest is the AstroRad vest, which will be tested on

Figure 2: Stuffed Whipple Shield for Starship with two

bumper shields (BS) and one back wall (BW), after [14]

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IAC-22-A5.2.3 Page 4 of 16

the Artemis missions [16]. Furthermore, the ECLSS is to

be expanded to include a radiation warning system that

will warn the crew when solar storms occur and they have

to seek shelter. The HERA (Hybrid Electronic Radiation

Assessor) radiation warning system, which is used

onboard the Orion capsule, is to be used for this purpose

[13].

Communications System

Communication onboard Starship should be possible

with a local network. For astronauts to communicate with

each other at any time, there should be a panel in every

room (including all living compartments) for audio

transmission only, with which every other room and thus

every other person can be contacted. In this way, other

people can be quickly informed of problems or warned in

case of emergencies. The transmission of video and data

is also to be made possible via mobile devices and a

wireless network.

Between Starship and the control centre on Earth,

communication during the flight should be carried out via

an optical communication system (laser communication

system). One such system, NASA’s Orion Artemis II

Optical Communications System (O2O), will be tested

on the Artemis II mission. It is supposed to deliver 10 to

100 times faster data transmission than conventional RF

(radio frequency) systems, so that even videos in UHD

can be received and sent on Mars. Furthermore, optical

communication systems are smaller, lighter, more

energy-efficient and more secure than RF systems. [17]

However, as a back-up, Starship should also have a

conventional RF antenna that can be used to

communicate via the Deep Space Network (DNS), as it

is the case with the Mars rovers and probes, for example.

Electrical Power System

The Electrical Power System (EPS) is responsible for the

generation, conversion, distribution and storage of

electrical power onboard Starship. Solar arrays, which

are to be stowed in the engine section during launch and

landing and to be deployed during the flight, allow

electrical power generation during the flight. Therefore,

they must not only be deployable but also retractable.

Similar to the Orion capsule, the solar arrays are

supposed to have a mechanism that allows them to

constantly align themselves with the sun so that they can

deliver full power.

Orion’s four 7 m long and 2 m wide solar arrays, each

consisting of three foldable panels, provide 11.2 kW of

power for a crew of four people [18]. Therefore,

Starship’s solar arrays should have about ten times the

power, 100 kW. In addition, the radiation intensity

decreases by about half during the flight to Mars. In order

for the solar arrays to deliver the required power near

Mars, they need to deliver at least twice as much power

near Earth, due to the reduced radiant energy on Mars of

590 W/m2 compared to 1,36 kW/m2 near Earth. [19, p.

19] With some margin for failing solar cells, for example,

an output of around 250 kW (the power of 100 kW with

the factor of two plus a margin of 50 kW) is required near

Earth. One solar panel that should be able to deliver this

amount of power is the MegaFlex from Northrop

Grumman (formerly ATK), which is foldable and unfolds

into a round panel by rotating 360°, as shown in Figure

3. The MegaFlex is a scalable system that is currently still

being tested, but its smaller version – the UltraFlex – is

already being used on, for example, the Cygnus

spacecraft and the InSight lander on Mars. So, the

technology is already proven and has a flight heritage. A

system consisting of two MegaFlex arrays, each with a

diameter of around 24 m, should be able to deliver this

power according to Orbital ATK. Together, the two

arrays have a mass about 2 t. [20]

Thermal Control System

The TCS should consist of two separate circuits – an

internal (ITCS) and external (ETCS) system, which is to

be based on that of ISS and Orion. Like them, the ITCS

should have water as a coolant, as this is not dangerous

to the crew in the event of a leak. Cold water is to be used

to cool systems. Water heated by waste heat from

systems is first used to heat certain areas before being

cooled by the ETCS through an Interface Heat Exchanger

(IFHX). The coolant of the external system absorbs the

heat of the internal system and releases it into space via

radiators. [21] HFE7200 is supposed to be used as the

coolant for the ETCS because it has a low freezing point

and low toxicity [22, p. 7]. For redundancy, there are to

be two internal and two external TCS.

In addition, there should be electric heaters and Multi-

Layer Insulation (MLI foil), which provides additional

low radiation shielding.

Extravehicular Activities

Airlocks are needed to carry out extravehicular activities

(EVAs). Similar to the concept of the HLS Starship,

Starship should have two airlocks [23]. This way, several

astronauts can go outside at the same time and

redundancy is ensured.

Furthermore, elevators are needed, as seen in the HLS

Starship concept, to bring the astronauts to the surface of

Mars. The elevators are to be suspended from two

extendable crane arms and have a platform with which

Figure 3: Deployment mechanism of the MegaFlex solar

array [45]

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IAC-22-A5.2.3 Page 5 of 16

the astronauts and the payload can be transported. In

addition to low-maintenance operation, the elevators

should also be able to function in strong winds so that

outside work does not have to be interrupted for weeks

due to dust storms.

Propulsion System

The propulsion system consists of the main engines, the

control thrusters (RCS thrusters), the main tanks for

liquid methane and liquid oxygen and the helium tanks

for pressurizing the main tanks. The system for orbital

refuelling of Starship is also included.

Since one engine has a mass about 2 t including the

mounting structure [5], this results in about 12 t for the

six main engines – three SL Raptor and three RVac engines.

Since Space Shuttle had 44 RCS thrusters [24], 50

RCS thrusters are assumed for the larger Starship. As a

rough estimate for the mass of a thruster, the 220 N RCS

thruster of the Orion capsule is used, which has a mass of

approximately 2 kg [25]. This results in a mass of

approximately 100 kg for Starship’s RCS thrusters.

For the mass estimation of the main tanks, those of

the Super Heavy booster are used. These currently have

a mass of approx. 80 t [5] for a propellant mass of 3,600 t

[5], but are still somewhat too heavy, which is why 70 t

are assumed. Based on Starship and a propellant mass of

1,500 t [26], the mass of the main tanks is about 29 t. The

helium tanks are assumed to have a mass of about 5 t.

Table 1 lists the assumptions made from the previous

subsystem chapters. For the crew, the assumption is made

that the ten people have a mass of around 80 kg each and

are allowed to carry 40 kg of luggage, considering that it

is a long mission duration. This results in 1.2 t for the

crew and its luggage. Since no information is available

on the masses of the remaining smaller subsystems

(EVAs, robotics, heat shield etc.), it is assumed that these

add up to 7 t. The total mass without propellant and

payload is assumed to be 200 t. Masses of subsystems

that are necessary for the crew or required for them to a

greater extent are not counted as unloaded mass, but

instead as payload mass. It is estimated in this work that

half of the masses of the EPS, TCS and others margin are

also needed for unmanned Starships such as the cargo

version, as these require less power due to the lack of life

support systems and thus also smaller solar panels and

batteries as well as no such extensive TCS and, for

example, no equipment for EVAs. Therefore, only the

halves of the masses of the EPS, TCS and others margin

as well as the masses of the structure, the meteoroid

protection and the propulsion system are counted as

unloaded mass. The unloaded mass of Starship is

therefore around 125 t according to Table 1. Since

Starship is supposed to be able to carry a payload of 100 t,

a total dry mass of 225 t is assumed in the following.

Therefore, in addition to the 200 t, an additional 25 t of

payload, such as the rovers, can be carried.

In addition to the total mass, the

to be applied is

required to calculate the propellant mass required for the

flight to Mars and back, as this is also important for the

propellant mass that needs to be produced on Mars. The

for the selected trajectory from NASA’s Trajectory

Browser refers to a launch from an Earth orbit at an

altitude of 200 km, which can be assumed for Starship,

because it still has to be refuelled. The assumed

for

arrival at and departure from Mars refers to the escape

velocity in a 200 km orbit around Mars. According to

NASA’s programme, landing on Earth will presumably

take place with parachutes, because no

is specified for

this. Table 2 shows the

’s with these assumptions in

the second column. For Starship’s mission, however, the

values for arrival at and departure from Mars must be

adjusted, because Starship will not enter and leave a Mars

orbit, but will land and launch directly on the surface of

Mars. For the supersonic retropropulsive landing burn at

an altitude of 2.5 km [27], a

of 1 km/s should be

assumed, as well as for the landing on Earth.

Now the

for the launch from the Martian surface

and the acceleration to the escape velocity in a 200 km

Martian orbit is to be calculated, to which the value

Table 2: Total !" for the planned mission

!" (km/s)

NASA [10]

!" (km/s)

modified

Start Earth

3.63

Arrival Mars

0.623

Total !" outbound flight

4.253

4.63

Start Mars

0.887

6.267

Arrival Earth

Total !" return flight

0.887

7.267

Table 1: Total dry mass budget of a crewed Starship

Mass (t)

Total mass (t)

Structures

Radiation shielding

101.3

Meteoride shielding

21.3

Structure

ECLSS

Radiation vests

0.27

20.27

Margin

EPS

Solar arrays

Cables

Batteries

Components

TCS

MLI

2.21

10.21

Margin

Propulsion

Main engines

51.1

RCS thrusters

0.1

Main tanks

Helium tanks

Pipes etc.

Margins

Crew + Luggage

1.2

8.2

Other

(EVAs, Robotics,

Heat shield etc.)

Total

200

Payload

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IAC-22-A5.2.3 Page 6 of 16

already calculated by NASA for the return flight is then

added. Including assumed velocity losses of 0.5 km/s at

launch and the velocity change for injection to Earth of

0.887 km/s, the change of velocity at launch is as follows

Δ𝑣!"#$" =𝑣%&'()** +Δ𝑣+,&& +Δ𝑣-./

(

4.88+0.5+0.887

)

-km

=6.267-km

(1)

With the simplified assumption that the

to be applied

is burnt in one piece, the following can be calculated

using the rearranged rocket equation:

𝑚*=𝑚0∙𝑒12

(2)

the mass

before and the mass

after burnout and

thus the required propellant mass

𝑚3=𝑚*−𝑚0

(3)

can be calculated. The exhaust velocity can be calculated

by the mean of the specific impulses of both engines [4] into

𝑣%(! =𝐼&4 ∙𝑔*=(3∙355 +3∙ 380)-s

6∙9.81-m

=3,605-m

(4)

Since the atmosphere of Mars is so thin and the pressure

so low, the specific impulses under vacuum conditions

are assumed for the launch on Mars. For the outbound

flight (EM = Earth-Mars), it is assumed that

corresponds to the total mass of Starship plus 10 t of

propellant residuals (which is a bit more than the

assumed 0,5 % by Elon Musk for the booster [5]), thus

235 t, resulting in the following values (included

propellant residuals):

𝑚*(56 =848.9-t

(5)

𝑚3(56 =623.9-t

(6)

For the return flight (indexed ME), the assumption is

made that with consumed food, no payload and 10 t of

propellant residuals,

is 200 t. This results in the

following values (included propellant residuals):

𝑚*(65 =1,501.4-t

(7)

𝑚3(65 =1,311.4-t

(8)

For the outbound flight, 623.9 t of propellant are needed

with a landing mass of 235 t, and for the return flight with

a landing mass of 200 t, 1,311.4 t of propellant are

needed.

4.2 In-Situ-Resource-Utilization

Propellant Production System

In order to produce the two propellants liquid methane

(LCH4) and liquid oxygen (LOX) on Mars, a propellant

production plant is needed.

To produce methane and oxygen on Mars, different

processes have to be used. Water and carbon dioxide are

needed to produce methane and oxygen. The water is to

be extracted from ice deposits located near the landing

site just below the Martian surface or from those found

on the surface. A suitable landing site with such deposits

must be found beforehand, as this is essential for

propellant production and thus also for the return flight

to Earth. Electrolysis is then used to separate the water

into its two components: hydrogen and oxygen [28]:

2 H2O ® 2 H2 + O2

(9)

The oxygen obtained is now in gaseous form and must be

liquefied. The hydrogen is further used in another

process, the Sabatier process [28]:

CO2 + 4 H2 ® CH4 + 2 H2O

(10)

In this, CO2 extracted from the Martian atmosphere, of

which the Martian atmosphere consists of 96 % [29], is

converted into methane and water together with the

hydrogen obtained from electrolysis. The methane, again

gaseous, must also be liquefied and can be stored in tanks

afterwards. The resulting water can either be fed to the

water electrolysis system or provided to the ECLSS to

supply the astronauts. Figure 4 schematically shows a

possible process for propellant production on Mars.

To produce and pump the water, a drilling and

pumping device is needed to melt the ice and bring water

to the surface. In order to extract CO2 from the

atmosphere, a filter system is needed first to remove

Martian dust from the air that is sucked in. Then the CO2

must be separated from the other gases in the air in a

separator. In addition, pumps, condensers, compressors,

coolers and transport pipes are needed. [28] For the first

missions, such a system could be transported completely

onboard one Starship as a complete system for propellant

production and would remain onboard. Thus, initially

only pipes would have to be laid and the drilling,

pumping and melting equipment for the ice and water

would have to be set up.

For the estimation of the propellant production

system (PPS), the 1,311.4 t of propellant calculated in

equation (8) should be assumed, which will be needed for

the return flight. However, since two crewed Starships

are to fly back, 2,622.8 t are required. With a duration of

368 days (Chapter 3) on Mars and a 30-day safety buffer

that the propellant should already be completely

produced before the return flight, 338 days are available

for production, resulting in a production rate of

7,760 kg/day.

A completely integrated propellant production

system that has already been tested on Earth under Mars-

like conditions is the Integrated Mars In-Situ Propellant

Production System (IMISPPS) from Pioneer

Astronautics. It has a single reactor that produces both

propellants. The system has a production rate of

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IAC-22-A5.2.3 Page 7 of 16

1 kg/day, a mass of 50 kg and requires 700 W of power.

[30] Assuming technological progress over the next few

years, the production rate of the system is estimated at

2 kg/day with a mass of 75 kg and a power of 1 kW.

Based on the required production rate and therefore

multiplied by 3,880, this results in a mass of 291 t and a

power of 3.88 MW.

However, additional power and additional mass will

be added for the water extraction, because in the

IMISPPS the hydrogen for the Sabatier process was

supplied from tanks and not extracted in advance [30].

Since no exact data is available for such a system, it is

estimated that the mass and power of the water extraction

system is one fifth of the propellant system, so that an

additional 59 t and 776 kW are added, giving a total mass

of 350 t and a power of 4.656 MW for the PPS.

In addition, LOX and LCH4 tanks must be built for

storage, whereby the tanks of the landed Starships are to

be used for this during the first missions. Based on

SpaceX’s mission plans, there should already be four

unmanned Starships on Mars ready for LCH4 and LOX

storage when the first two crewed Starships arrive in

2026. Propellants can be loaded and unloaded via the

ports that allow for orbital refuelling. This would allow

space for other payloads on the first missions. For the

transfer of propellants from the Starships converted into

storage facilities to the crewed Starships that are to return

to Earth, flexible transport pipes must be laid or

refuellable rovers used. To prevent the pipes from

becoming too long and to keep the distances as short as

possible, the Starships must all land close to each other,

which is possible thanks to the precise control system.

The risk of damage from kicked-up dust and stones

should be investigated beforehand.

Power Supply System

A power supply system (PSS) is needed for propellant

production, Starships, rovers, future habitats and all other

activities on Mars. Nuclear reactors are to be used as the

primary power source, because the use of a solar system

as the main power source comes with some

disadvantages. For one thing, the received energy output

on Mars is only half as much (590 W/m2) as on Earth

(1,361 W/m2), due to the further distance from the sun.

In addition, solar panels have an efficiency of only about

30 %, which further reduces the usable energy. [19, p. 19]

The panels can also only provide power during the day

and would be very limited during months of dust storms.

Dust also accumulates on the panels over time, which

also reduces power. Nuclear reactors operate

independently of ambient conditions, provide power

even at night and do not consume as much space in terms

of area as comparably powerful solar systems.

For the first mission, in addition to two crewed

Starships with ten people each, the propellant production

plant and rovers must also be supplied by the power

supply system. Assuming two Starships, each requiring

100 kW of power (Chapter 4.1), plus 4.656 MW

(previous Chapter) for the propellant production plant

and additional power for the rovers, it is assumed that the

PSS must provide 5 MW of power.

The Kilopower system launched and demonstrated by

NASA, whose follow-up project Fission Surface Power

(FSP) is now being continued together with the Space

Nuclear Power Corporation (SpaceNukes), is a scalable

system in which a reactor core made of enriched uranium

heats sodium in heat pipes, which in turn lead to Stirling

engines that then convert the heat into electrical power

[31] [32]. It is a small and light system for its power, very

safe and four such reactors shown on the left in Figure 5

should be able to supply a base with four people [33]. The

10 kW system is expected to have a mass of about

1,500 kg and be able to produce electricity for up to ten

years [34] [35].

In Figure 5, the reactor block (black) and the reactor

casing (silver), in which the heat pipes are located, can be

seen at the bottom with the Stirling engines above them.

The shield above the FSP system, which is folded up

Figure 4: Process for propellant production on Mars, from harvested water and collected CO2, after [43]

Published by the IAF, with permission and released to the IAF to publish in all forms.

IAC-22-A5.2.3 Page 8 of 16

during transport and only unfolded on Mars, is the

radiator.

Two 2 MW systems, which are required for this

power demand, will have a mass of approximately 32 t

each. [34]. Such a system will look similar to the 650 kW

system shown on the right in Figure 5, only larger. It can

be seen that there is a fan at the top above the Stirling

engines, which is intended to provide better cooling due

to the increased heat generated. It should be ensured that

this can dissipate sufficient heat given the low

atmospheric density on Mars. For transport, the reactor

block below the fan can be completely stowed away in

the upper large casing. In addition to these two systems,

a 1 MW system is needed to deliver the required 5 MW of

power. Its mass is estimated at around 20 t.

To protect astronauts and surroundings from

radioactive radiation, either the reactor must be

surrounded by a protective shield or the reactor must be

embedded in the Martian soil, whereby the second option

saves additional mass and should therefore be preferred.

The stated masses of the systems already include the

protective shield. A 4.8 m deep hole is needed to bury the

reactor block in the Martian soil [34]. With an estimated

cylinder diameter of 2 m, the volume to be excavated is

15 m3 and with a density of the Martian soil of

1,680 kg/m3 [36], this corresponds to a mass of around

25 t. Furthermore, it must be ensured that in the event of

a launch failure, the reactor remains switched off and will

not be activated.

Solar panels could still serve as a back-up and

additional power source. Solar panels such as the

MegaFlex could be used again for this purpose. To ensure

that the power does not decrease over time due to settled

dust, a mechanism for removing Martian dust should be

developed for solar panels on Mars. Lithium-ion batteries

should be used for temporary storage and power supply

at night, which can also provide a short-term back-up and

ensure the power supply in the event of a power failure.

Transportation System

Unlike the robotics subsystem, which is responsible for

transporting payloads onboard Starship, the

transportation system is responsible for transport on the

Martian surface. Different rovers are needed to transport

astronauts and objects and to build infrastructure. To

facilitate the construction of infrastructure, rovers are

needed that can move the heavy and bulky payloads from

the Starships on the Martian soil. Rovers with shovels

and drills, like NASA’s RASSOR rover, will help create

flat surfaces for future habitats and the propellant

production plant. Future habitat modules could also be

covered with Martian regolith or built into rocks for extra

radiation shielding. If transporting the two propellants

from the production plant to the tanks or from the tanks

to the Starships to be refuelled by means of transport

pipes is too costly, for example because the distances are

too long, there should be rovers with tanks that take over

this task and constantly shuttle back and forth. For the

construction of paved roads and landing zones for

Starships, rovers with a printing head should be used.

These melt the loose regolith and solidify it, creating a

solid flat surface. With paved roads, rovers can travel

faster and easier between different locations. The

advantage of paved landing zones is that there is less dust

when landing and it can be done on a flat surface. For the

astronauts, there should be both rovers that, like the lunar

rover of the Apollo missions, can be used with a spacesuit

for short distances and ones in which the astronauts can

travel longer distances without a spacesuit.

In order to avoid having to develop and build a

completely new rover for each application, a simplified

rover system such as the concept of the Modular Robotic

Construction Autonomous System (MOROCAS), which

is illustrated in Figure 6, is a suitable solution [37, p. 4].

As the name suggests, this concept is a modular

autonomous rover system.

The basis is a chassis that can accommodate various

modules and tools in a standardised dock and thus

perform many different tasks. For example, it can hold a

module with a tank, a shovel, a drill and one for crewed

transportation. [37, p. 4] The shovel module must have

Figure 6: MOROCAS concept with different modules

[36]

Figure 5: 10 kW (left) & 650 kW (right) FSP System

[33]

Published by the IAF, with permission and released to the IAF to publish in all forms.

IAC-22-A5.2.3 Page 9 of 16

an excavator arm about 6 m long and be strong enough to

dig the 4.8 m deep holes and 25 t of Martian soil for each

of the reactors. In addition, there could also be modules

with a printing head, a gripper arm and a cargo bed. Such

a modular system saves costs and additional mass. In the

report of the concept, it is estimated that a rover needs

10 kW/day of power [38, p. 18]. Unlike the concept, the

rovers should be powered by batteries instead of with

Multi-Mission Radioisotope Thermoelectric Generators

(MMRTGs). Since the two Mars rovers Curiosity and

Perseverance have a mass of around 1,000 kg [39], it is

assumed for the MOROCAS concept that the base has a

mass of 800 kg without scientific instruments, but with

batteries. Five such rovers are to be transported.

Including additional modules, a total mass of this

subsystem of 10 t is assumed.

5. Feasibility

The feasibility analysis of the planned baseline scenario

with regard to Starship and ISRU systems and the

associated requirements is conducted on the basis of the

following key figures: Launch mass, Unloaded mass,

Payload mass, Technology Readiness and Costs.

5.1 Launch Mass

The payload to be transported and the

to be applied

are decisive for the evaluation of the launch mass. The

required for a launch from the surface of Mars and the

flight to Earth was already calculated in Chapter 4.1 at

7.267 km/s, as well as the resulting propellant mass of

1,311.4 t and launch mass of 1,501.4 t. So, there is still a

buffer of 188.6 t until the maximum permitted launch

mass of 1,690 t is reached. With these, if used purely for

the propellant, for a landing mass of 200 t (see Chapter

4.1) an additional

$%!"#$"%&%#'' & %(%! '

(

+,-./#0

1//#0

3)*

(

+,4/+56#0

1//#0

& /561-#78

(11)

could be generated.

Likewise, the launch from a 200 km Earth orbit has

already been calculated, where only 623.9 t of propellant

are required for a

of 4.63 km/s, since most of the

Earth’s gravity has already been overcome here. The

launch mass from orbit after refuelling is 848.9 t. If the

tanks were fully filled with the additional possible 876.1 t

of propellant, with a landing mass of 235 t (see Chapter

4.1) an additional

$%!"#$"%)*%#'' & %(%! ' ()*(+,:14#0

1;4#0 23 )*(<6<5.#0

1;4#0 22

& 1544-#78

$%&'(

would be achievable. Thus, either a shorter travel time of

only 192 instead of 304 days [40] or the transport of more

payload could be realised. A shorter travel time in turn

increases the length of stay on Mars, which in turn

provides more time for propellant production, allowing a

reduction of the PPS and PSS.

The only launch mass that remains to be verified is

that of the entire Starship system during a launch from

Earth into a 200 km Earth orbit. In the following, it will

therefore be investigated whether the launch is possible.

For this purpose, the launch mass and the required

are

calculated. The launch mass results from a fully fuelled

and fully loaded Starship (1,725 t) and a Super Heavy

booster (3,760 t [5]), which is also fully fuelled, at

5,485 t. The required

is then calculated. Now the

to be expended must be calculated. With a special

gravitational constant of the Earth of

𝜇5=398,599-km7

s)(

$%)'(

the orbital velocity in a 200 km Earth orbit with a radius

*#,%!!

of 6,578 km can be calculated to be

𝑣'()** =

$#,%&& =7.78-9:

;

$%,'(

Furthermore, the Earth’s rotational velocity at launch

latitude, in this example from Boca Chica in Texas (

- .

&/0

), calculates to

Δ𝑣$,"(<= =)>$

)?@A =)>B"CD; E

FG?**@; =0.42-9:

;+(

$%1'(

When launching from Earth, losses for the steering angle

(0.1 km/s), air drag (0.2 km/s) and gravity (1.5 km/s)

must be overcome and added to the circular orbital

velocity and the Earth’s rotational velocity at launch

latitude must be subtracted [41, p. 69]. In addition, a

margin of roughly 5 % (

!"&. 2+,&(3456

) is to be

added according to ESA [42, p. 7]:

$%!"#$"%) & %+%,-- =$%!"(($./0 =$%1$#0 = $%2$#3."4

3$%$5"%67 = $%&& .54<#78

$%/'(

According to equation (7) converted to

, the booster

with the launch mass

, the mass

after burnout of

1,925 t (1,725 t Starship + 160 t empty mass booster +

30 t propellant for booster landing + 10 t propellant

residuals booster) and an exhaust velocity

"',(

3,237.3 m/s according to equation (4) (330 s

7)*

)

produces a velocity change of

Δ𝑣<,,&"%$ =𝑣%(< ∙lnBH&

H'C= 3.390-9:

(17)

After stage separation, Starship has to generate the

remaining 6.19 km/s. If 1,725 t are used as the mass

and the previously assumed exhaust velocity are inserted

into equation (2) converted to

and into equation (3),

this results with added propellant residuals of 10 t already

assumed in Chapter 4.1 in a propellant mass of

𝑚3(!"#$&IJ4(!"#$" =1,425.2-t.

(18)

Published by the IAF, with permission and released to the IAF to publish in all forms.

IAC-22-A5.2.3 Page 10 of 16

This is less than the maximum propellant mass of 1,500 t,

i.e. the launch is possible. Afterwards, 74.8 t of useful

propellant remain onboard Starship, which requires

correspondingly less mass to be refuelled. The limiting

flight phase is therefore the launch from Earth, since

95 % of the maximum propellant mass is used for a

payload of 100 t, whereas this is only 41.6 % for a launch

from Earth orbit.

5.2 Unloaded Mass

For better comparability with other Starship variants and

with the given data, as well as to determine the maximum

possible additional payload, the masses of the subsystems

of a crewed Starship, which are necessary or required to

a greater extent for the crew, were not included in the

unloaded mass but in the payload mass in Chapter 4.1.

The unloaded mass of Starship was therefore about 125 t

according to Table 1. With a stated unloaded mass of

100 t, this means a significant deviation of 25 %. Since

current prototypes are presumably neither equipped with

a TCS and EPS nor with meteoroid protection and have

a mass of around 100 t, the estimated mass could

nevertheless be realistic. Conversely, this means that

either 25 t less payload can be transported in order not to

exceed the 200 t (100 t unloaded mass + 100 t payload),

or, as assumed in this work, a regular 100 t payload is

added, which then increases the originally planned

launch mass.

For a cargo Starship with an unloaded mass of 125 t

and a payload of 100 t, this also results in a total dry mass

of 225 t. What distinguishes a cargo Starship from the

crewed version is that it will initially be fully loaded only

on the outbound flight to Mars, and on the return flight it

will be about 100 t lighter and therefore either the

required

or the propellant mass would be lower. Since

for both versions the propellant masses required for

launch from the surface of Earth and Mars and from a

200 km Earth orbit are below the maximum possible

1,500 t, it follows that a 25 % increase in unloaded mass

with an additional 100 t payload is feasible for both a

cargo and crewed Starship and does not represent a

critical condition.

5.3 Payload Mass

The payloads to be transported for the first Starships are

the propellant production system, the power supply

system and the rovers. A system mass of 350 t was

assumed for the PPS. The total mass of the PSS,

consisting of two 2 MW systems of 32 t each and one

1 MW system of 20 t, is 84 t. Five rovers of 800 kg each

plus additional modules were estimated at 10 t.

According to the baseline scenario, a total of four cargo

Starships with a payload capacity of 100 t each and two

crewed Starships with an available capacity of 25 t each

for additional payload are available until 2026 to bring

the required systems to Mars. Considering the mass

alone, 444 t would have to be distributed among six

Starships with a payload capacity of 450 t. This would be

feasible, but it presupposes that the volumes of the

payloads can also be distributed appropriately among the

Starships as the payload volumes of the Starships are

limited. Thus, the PPS, a 2 MW reactor and the 1 MW

reactor could be distributed among the four cargo

Starships. The rovers could then be transported onboard

one of the two crewed Starships and the second 2 MW

reactor onboard the other. Since the reactor would exceed

the payload capacity of 25 t, the following examines how

much additional payload mass could be transported if the

maximum propellant mass of 876.1 t is utilised. Since the

PPS should ideally remain onboard the Starships after

landing in order to reduce the logistical effort, but it must

be divided among four Starships, the Starships must be

very close to each other so that pipes for connecting the

individual units are as short as possible.

Both during the launch from Earth and from a 200 km

Earth orbit, the maximum propellant mass was not yet

fully utilised, as was shown in Chapter 5.1. Instead of

using this additional possible propellant for a shorter

travel duration, the maximum possible payload with an

unchanged travel duration will be calculated in the

following. Calculating the maximum possible payload

mass now is an optimisation task, because with it the

launch masses of the booster also change and thus also

the velocity changes achieved. This is calculated for

launch from Earth, as this is the limiting flight phase.

Equation (19) is a combination of equations (2), (3) and

$%/'.

>8%9#:%!"#$;<.=%!"#$" &?+,:14=@A#0

3?+,:14= @A0

B>3!"#$",&?3',(@ABCD%EFDG:

H%I,DG:JKL

3',!

& +,6./#0

(19)

Mass

includes not only the additional possible payload

mass but also structural mass needed to support the

additional payload mass. It is included in the total mass

of Starship, the booster and the mass after the booster has

burned out and was steadily increased in small steps until,

with

𝑥 = 81.7(

$&2'(

the maximum useful propellant mass of 1,490 t (the

maximum of 1,500 t minus the 10 t residuals) was almost

reached at 1,489.96 t. SpaceX assumes that for every

tonne of payload, another tonne of structure and

propellant will be added [6]. The payload and structural

mass of 81.7 t is therefore added to the propellant mass

of 74.8 t from Chapter 5.1 and then divided, resulting in

the maximum possible additional payload of

𝑚K(H#L =MF0NOPO?NFQ@R

)=78.25≈78-t.

(21)

Published by the IAF, with permission and released to the IAF to publish in all forms.

IAC-22-A5.2.3 Page 11 of 16

Assuming 10 t of propellant residuals after landing, the

onward flight to Mars from a 200 km Earth orbit

subsequently requires

>8%)&%M9#: &;+-5:#0'B>3&)

3',! 3;+-5:#0 & <1:5;/#0

(22)

of propellant according to equations (2) and (3).

With an additional payload capacity of 78 t per

Starship, the total payload capacity of the four cargo and

two crewed Starships would increase to 918 t. The

payload volume remains of course unchanged, but if the

volume is suitable, the PPS could be distributed over only

two cargo Starships and the PSS into one Starship. This

would create space for scientific instruments, more

rovers, back-up solar arrays, a habitat module or similar.

Since the crewed Starships still have a large portion

of their payload onboard when they are launched on Mars

due to the systems needed for the crew, it must be ensured

that a certain total dry mass is not exceeded. If the tanks

of the crewed Starship on Mars were to be completely

filled for a return flight with 1,500 t of propellant (1,490 t

for the flight + 10 t of propellant residuals after landing),

a maximum landing mass of 229 t including 10 t of

propellant residuals and thus a total dry mass of 219 t

would be possible at a required

of 7.267 km/s. With

the assumption already made in Chapter 4.1 that 10 t of

food were consumed at this point, 29 t of additional

payload could be transported back to Earth. If the payload

capacity of the crewed version were to be increased, the

additional structural mass required for this would have to

be subtracted from these 29 t.

According to Table 1, the mass of the subsystems

required for the crew of a crewed Starship is around 75 t.

The largest portion is accounted for by radiation

shielding at 40 t and the ECLSS at just under 20 t. This

should not be a problem for the first missions, but if up

to 100 people are to be transported in later missions, the

subsystems required for this would also have to be larger.

A larger volume, which has to be surrounded by the

heavy radiation shield, and ten times as many people,

who have to be provided with recycled water and

atmosphere, increase the respective subsystem masses by

at least double, it is estimated. Even with 78 t of expanded

payload capacity, but of which only 29 t of additional

mass may be onboard Starship on the return flight, this

scenario will be difficult to achieve.

Another problem besides the sheer mass is the power

supply for 100 people. The power of 100 kW already

required for Starship with a crew of ten, or 250 kW near

Earth, would have to be between 2-2.5 MW for such a

large crew. Solar panels that could deliver such power

would probably have to be 60-80 m in diameter if a pair

of two 40 m panels is to produce 700 kW and with a

slightly exponential power-to-size ratio [20]. Such large

panels not only entail the difficulty that they have to be

retractable, but also that they are relatively long,

estimated at 20-30 m when folded, and have to be stowed

on or in Starship. No solar arrays can be seen in current

renderings of Starship, only in the very first design of the

ITS (Interplanetary Transportation System). This

possibly indicates that SpaceX itself is moving away

from solar arrays and wants to rely on nuclear reactors

such as the FSP system. One of the advantages of these

is that the system does not have to be designed to be twice

as powerful near Earth in order to deliver the required

power near Mars. But then there would be the problem of

mass, which is estimated at around 20 t for a 1 MW

system, and the question of the compatibility of a nuclear

power supply and people onboard a spacecraft.

5.4 Technology Readiness

The radiation and meteoroid protection as well as the

components of the ECLSS and TCS are technologies that

have already been used on previous spacecraft or the ISS

and therefore have a high degree of technology readiness.

The first technology currently being tested on many

different missions is the optical communication system.

In 2026, Deep Space Operational Services with multiple

terminals forming a network are expected to commence

and be available to missions from that date onwards [43].

Based on this described extensive testing over the next

few years, it can be assumed that the technology

readiness of an optical communication system will be

high by 2026.

The MegaFlex solar arrays of the EPS are based on

the UltraFlex, which has a flight heritage and, with

TRL 9, the highest possible technology readiness level

(TRL). A MegaFlex array with a diameter of 9.6 m has

already been tested in the course of a TRL 5

demonstration, but it has not yet been tested in space or

in the size required for Starship. As the technology is

available but still needs to be scaled to the required size

and a mechanism for retracting the solar arrays needs to

be developed, it is quite possible that this could be

operational by the planned launch date of 2026.

If NASA’s xEMU are used as spacesuits for EVAs,

then their technology readiness should have reached a

TRL of 9, as they will already have been used on the first

Artemis lunar missions beforehand and will therefore

have flight heritage. It is considered feasible that the

modifications that need to be made for a mission in the

Martian atmosphere can be implemented by the time of

the launch in 2026. In contrast, the required elevators

must first be designed and tested.

The next new technology to be considered are

Starship’s main engines, the Raptor engines. At the time

of writing, they have not yet been used in space, but will

be from 2022 onwards during the orbital test flights.

However, they have already been used in numerous test

flights on Earth, so their level of technology readiness is

high. For the RCS thrusters, for which no more precise

specifications are yet available other than that they will

probably be cold gas thrusters, it can be assumed that

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IAC-22-A5.2.3 Page 12 of 16

existing thrusters or similar to these will be used here, so

that the technology readiness will also be given for these.

The system for orbital refuelling, however, has not yet

been developed. The feasibility of such a system must be

demonstrated by the launch of the first two cargo

Starships in 2024 and successfully tested during several

tests, which is considered feasible due to the need to carry

out all Mars missions on the scale SpaceX is planning.

The heat shield technology also still needs to be

extensively tested during re-entries and possibly adapted,

depending on what the tests reveal. However, it is

believed that the heat shield tiles will meet their

protection and reusability requirements by the first

launch in 2024, especially since SpaceX already has

experience designing a heat shield for the Dragon space

capsule.

That the technology to produce liquid methane and

oxygen would work on Mars has already been

demonstrated with the IMISPPS presented in Chapter

4.2. However, this system has only a fraction of the

propellant production rate needed to fuel two Starships.

In addition, such a system without a water extraction

plant requires with 3.88 MW a lot of power even with the

assumptions of technological progress that have been

made. Besides the mass and required power of such a

system, the volume could also become a problem. It must

be possible to distribute the system over several

Starships. If it has to be distributed on more than four

Starships because of the volume, the crewed mission is

not feasible in 2026, as only four cargo Starships will be

available by then. Therefore, it is seen as critical that such

a system is operational and flight-ready by 2026. The

current technology readiness and feasibility are therefore

low.

The technology readiness of the Fission Surface

Power System is also not yet very advanced. Technology

demonstration has already taken place successfully on

Earth with a smaller 1 kW reactor. The problem with the

FSP, however, is the estimated time to operational

capability. Systems with 10-20 kW are expected to be

flight-ready in 3-5 years, the required 2 MW system only

in probably ten years [34]. It can therefore be assumed

that this technology will not be available for the planned

2026 mission.

The last new technology to be discussed that was

proposed in the system designs is the MOROCAS rover

concept. This is only a concept so far, but individual

modules such as shovels and drills have been

demonstrated on other rovers such as NASA’s RASSOR

or VIPER and will be used on future missions. The

development of a common platform that can

accommodate the various modules and the different

modules is considered feasible, as technology from

previous rovers can be used.

5.5 Costs

In order to be able to estimate the rough costs of Starship

and booster Super Heavy, the costs that SpaceX forecast

for the ITS concept in 2016 should be used as a reference.

However, these values have to be scaled down because

the ITS had a larger fuselage diameter of 12 m (instead

of 9 m), thus larger tanks and consequently a higher

payload capacity of 450 t. Furthermore, the primary

structure was to be made of carbon fibre instead of

stainless steel. [44] Table 4 shows the estimated costs of

the ITS and next to it the costs related to the current

Starship.

For the manufacturing and maintenance costs of the

current Starship, those of the ITS concept are multiplied

by the factor 0.5625, the ratio of the different fuselage

cross-sectional areas, which takes into account the

smaller volume of the current Starship. In addition, the

different material must be taken into account in the

manufacturing costs. For this purpose, the product of the

structural mass of 40 t of Starship and 55 t of Super

Heavy (analogous to Starship in Chapter 4.1) multiplied

by the cost of carbon fibres, 130,000 US$/t, is subtracted

from the already adjusted manufacturing costs and

replaced by the product of the structural mass multiplied

by the cost of stainless steel, 2,500 US$/t. [4] It is to be

assumed that a Starship flying to Mars can complete ten

flights and the booster and tanker 100, as these two only

launch into near Earth orbit. The number of launches

refers to a single crewed Starship from the baseline

scenario, which has to be refuelled with 623.9 t in Earth

orbit. However, there is still 84.8 t of propellant onboard

(74.8 t + 10 t of residuals) if Starship is launched from

Earth with full tanks as in Chapter 5.1 and assuming the

100 t of payload described in the baseline scenario. The

tanker Starships therefore only have to provide 529.1 t of

propellant, since 10 t of residuals are still onboard. It is

assumed that the tanker Starships, like the booster, need

30 t of propellant for landing and 10 t of propellant

residuals, so that according to equation $&2' an

8 . ,9

and thus a maximum extended payload capacity of 47 t

results. The tanker Starships can therefore carry 147 t of

propellant for refuelling, which minimises the number of

refuelling flights, but does not quite match the 165 t [45,

Table 3: Cost comparison of different rockets

Atlas V [46]

Delta IV

SLS Block 1

Falcon 9 [47]

Falcon Heavy

[47]

Starship

Costs (US$)

163 M

350 M [48]

876 M [49]

62 M

90 M

39 M

Payload to Mars (t)

8 [50]

20 [51]

4.02

16.8

100

Costs/tonne (US$/t)

32.6 M

43.8 M

15.4 M

5.4 M

0.39 M

Published by the IAF, with permission and released to the IAF to publish in all forms.

IAC-22-A5.2.3 Page 13 of 16

p. 4]. Four tanker flights are thus needed to refuel the

crewed Starship. The propellant costs are calculated from

the price of the propellant of 168 US$/t and the propellant

mass [44]. Maintaining Starship flying to Mars is more

expensive, because after a long journey to Mars, all

systems must be examined particularly thoroughly before

it sets off on a new Mars journey. The reason why the

tanker Starship is more expensive to maintain than the

booster is that the heat shield, among other things, has to

be inspected. The costs of 200,000 US$ per launch for

the launch site are taken over by the ITS concept.

The total costs per Mars trip of the current Starship

result from the manufacturing costs distributed over the

number of expected launches during a lifetime, the

maintenance, propellant and launch site costs multiplied

by the number of required launches and a margin of

20 %. A trip to Mars thus costs 39 M US$, which

corresponds to 390,000 US$/t for a 100 t payload. Table

3 compares the costs per tonne of payload of different

rockets.

As can be seen in the comparison, the cost of Starship

is only a fraction of the other rocket systems compared.

It should be noted that the prices and payload capacities

of the Falcon 9 and Falcon Heavy refer to a non-reusable

version with the maximum possible payload. Even if

their prices were to drop by half for a reusable rocket, the

cost/tonne ratio would also drop by half but would still

be significantly more than for Starship. Assuming that

Starship’s cost calculation is based on a single launch,

rather than distributed over the number of expected

launches over a lifetime, this would result in a cost of

12 M US$/t, which would still be at least a third of the

other rocket systems, with the exception of SpaceX’s

rockets. In this case, the Falcon Heavy would even be

55 % cheaper.

The five boosters and four tanker Starships must

therefore be reusable at least three times in order to

undercut the Falcon Heavy’s cost of around 5 M US$/t.

The Starship flying to Mars does not even have to be

reusable. Due to the low cost of Starship, even with a

lower number of possible launches within the lifetime,

the costs are seen as feasible.

For the technologies required, those with a low level

of technology readiness generally require higher costs

than those whose development is already more advanced.

For example, the costs of the radiation vest, the optical

communication system, the solar arrays, the space suits,

the thrusters, the heat shield and the rovers, which unlike

previous rovers do not have expensive scientific

instruments, are likely to be within a feasible range, since

they have either already been tested or will be tested in

the near future, or can be based on similar existing

technologies.

This is not the case for the nuclear reactors, the

propellant production system and the elevators, which are

very expensive to scale up or have to be designed in the

first place. For example, the development and

construction costs of the 2 MW reactor are estimated at

one billion US dollars [34]. Similar costs could also apply

to the PPS. Due to these high costs, it is not considered

feasible to implement in the available time window.

6. Discussion

The biggest problems that have arisen in the analysis in

this work are caused by the PSS, PPS and EPS and

concern the mass, their required power and their

produced power, respectively, and the technology

readiness. With the currently available technology for

propellant production, this system requires too much

power for the size it needs for production for two

Starships and is also too heavy. Technological progress

was already assumed in the calculation of the PPS.

Should this not occur, the required power and mass

would increase by 40 % and 30 % respectively compared

to the assumed values.

Distributing the PPS, PSS and rovers among the four

cargo and two crewed Starships with a standard payload

capacity of 100 t is feasible in terms of mass, but the

volume of the individual components and the available

volume of the Starships will probably be a problem, so

that the 100 t might not be fully utilised. In the case that

the payload volume is the problem, the possible extended

payload capacity of 78 t is of no use, because the volume

of the payload area of Starship does not change.

However, if it is not the volume but the maximum

payload capacity that is a problem, the possibility of the

extended payload capacity could be used, which would

increase the feasibility, because more space would be

available and the additional payloads could be

transported more easily.

As the main problem however, the high power

requirement of the PPS of 4.656 MW is seen, which leads

to heavy nuclear reactors with high power. In addition to

the mass, these have the problem that their technology

Table 4: Costs ITS [43] and current Starship

ITS

Current

Booster

Tanker

Starship

Booster

Tanker

Starship

Manufacturing costs (US$)

230 M

130 M

200 M

122.36 M

68 M

107.4 M

Lifetime launches

1,000

100

Starts/Mars trip

Average maintenance costs/flight)

0.2 M

0.5 M

10 M

0.11 M

0.28 M

5.63 M

Propellant costs/launch (US$)

1.13 M

0.42 M

0.33 M

0.6 M

0.28 M

0.25 M

Launch site costs/launch (US$)

0.2 M

Total costs/Mars trip (US$)

11 M

8 M

43 M

13 M

6 M

20 M

Published by the IAF, with permission and released to the IAF to publish in all forms.

IAC-22-A5.2.3 Page 14 of 16

readiness is not yet very high, which in turn leads to high

development and construction costs as well as a long time

span until flight readiness of about ten years.

The fact that an elevator does not yet exist that has to

bring astronauts and payloads to the surface of Mars even

during dust storms is also problematic, as this is

something that has not existed before and the

requirements for this system are very high. This is

because the elevators must be able to operate even during

dust storms. The moving components, which are then

particularly exposed to sand, must therefore be designed

in such a way that sand cannot penetrate the system

anywhere and lead to malfunctions.

For future flights to Mars with the planned 100 crew

members, there is another problem – the power supply

for Starship. Either huge solar panels would have to be

used, a big leap in solar panel efficiency would have to

happen by then, or reactors would have to be relied on,

but this will again raise political questions of

compatibility.

Starting the propellant production already two years

earlier would drastically reduce the required power of the

system. At 1,068 days, the production rate of the PPS

would have to be only 2,456 kg/day, with a mass of 92.1 t

and 1.23 MW of power, including the water extraction

system, that is about 115 t and 1.54 MW. However, the

feasibility of this idea is difficult because the system,

which is distributed over several Starships, would have

to be connected by robots to form one system. In

addition, the Starships would have to land practically

directly on an ice deposit so that it could be used directly

for production. In addition, there is the connection of the

PSS with the PPS and the transport of the produced

propellant into the propellant tanks of the second cargo

Starship for storage. All processes would therefore have

to be executed automatically by robots, whose control

cannot take place in real time either. If anything should

go wrong and the system cannot produce any propellant,

this can only be fixed when the crew lands two years later

and then the propellant production system is designed too

small to produce the required amount of propellant for

the return flight in the remaining time span. Of course,

such a problem can also occur during a crewed mission,

but a human being is better able to solve an initially

unknown problem. Another possibility would of course

be to wait until production has started and is functioning

and only then, when this has been achieved, to launch the

crewed Starships. But here the difficulty remains that the

system is distributed over several Starships and must be

automatically assembled and made functional by robots.

Starting propellant production only with the arrival of the

crewed Starships may therefore seem risky at first and

also definitely represent a risk factor, but in the end it is

probably the safer way. Moreover, even with the

extended production period, a 2 MW reactor is still

needed, and its availability of ten years remains

unchanged.

Due to the lack of alternatives for the problematic

systems described above, these hurdles cannot be

avoided more easily with other technologies. The use of

solar panels instead of nuclear reactors represents too

great a risk in dust storms, and there is no way around a

propellant production system, since transporting 2,622.8

tonnes of propellant to Mars is also not practical and

therefore not feasible. The problem with the elevator also

cannot be solved in any other way due to the design and

landing manoeuvre of Starship. For these reasons, it is

concluded that SpaceX’s expanded mission plans in the

baseline scenario are not achievable and feasible at this

scale and timeframe by 2024/2026.

If the time until the nuclear reactors of the PSS are

actually ready for deployment is ten years, this would be

deployable in 2032. This would also allow time for the

development and scaling of the PPS, which in the best

case can be made smaller, lighter and more power-

efficient by then. If these hurdles can be overcome by

then, the first cargo Starships could be launched in 2030

and the first crewed Starships in the following launch

window in 2032. For these launch windows, a feasibility

analysis must then be carried out again based on the

required velocity changes, the duration of stay on Mars

and thus the demands on the PPS. The issue of planetary

protection should also be considered in detail in order to

keep human contamination of Mars for scientific

experiments as low as possible. However, this cannot be

completely avoided when astronauts set foot on the

surface.

7. Conclusion

This work analysed the feasibility of current plans for

SpaceX’s Starship. The analysis was based on the current

system, the assumptions made and selected technologies,

to give an overview of the subsystems and to identify

existing problems, can be said to have been achieved.

The final conclusion of the analysis, that the current

mission plans are not feasible at this scale and timeframe,

was made mainly due to the masses (350 & 84 t),

volumes, required and produced power (5 MW) as well

as the technology readiness of the propellant production

system and power supply system. In addition, there is a

lack of feasible alternatives. Even though the elevator

and the larger electrical power system required for future

missions with 100 crew members are major hurdles to

overcome with Starship, the feasibility of the missions is

significantly influenced by the required ISRU

components. If these hurdles can be overcome, a Mars

mission with Starship could be possible at a later time.

Published by the IAF, with permission and released to the IAF to publish in all forms.

IAC-22-A5.2.3 Page 15 of 16

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Cryogenic Liquid Rocket and Nuclear Thermal Rocket: A Survey of the Current Technology and a Comparison for Future Mars Missions

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Feb 2024

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Full-text available

Oct 2020

Searching for an alternative planet to be inhabited by humans is currently quite a popular topic, especially considering the current state of technology development and future opportunities. Mars is the closest planet in the Solar System, located in the ecosphere, which despite the difficult conditions on the surface, is planned to be colonized. However, before this happens, the Red Planet must be better known. The first challenges that await the future colony are to choose the right location and develop technologies to facilitate the water acquisition (e.g., from ice caps on the poles or hidden under the surface of regolith). In order to enable human survival in the initial colonies, it is necessary to create a habitat to enable the everyday functioning of the first settlers and to protect them from low temperatures, a rare atmosphere whose composition is mainly carbon dioxide and where the amount of oxygen is below 1%. Lack of Mars magnetic field makes it impossible to protect against cosmic and solar radiation. Another problem is the occurrence of sandstorms, which can last for months, limiting access to natural light. A natural consequence is to design infrastructure adapted to current conditions and chosen location. Habits such as a cave system or the location of the base in a canyon, whose walls would additionally protect against difficult conditions, are considered. The aim of present paper is evaluation of Mars colonization possibility depending on conditions on this planet and also the location of the colony in accordance with the landscape. The paper presents a proposal to minimize the impact of negative factors on the lives of future Mars residents, related to, inter alia, the architecture and location of the base, food production, energy generation and transport of colonists, based on the Ideacity project proposed by the InnSpace team.

A Feasibility Analysis of a Variable Mass Multiple Asteroid Retrieval Mission

Conference Paper

Full-text available

Mar 2020

This paper aims to analyze the feasibility of a space-mining operation by estimating hardware acquisition and refueling operations expenses for a 30-year Multiple Asteroid Retrieval Mission (MARM). The study builds on previous work on possible Near-Earth Asteroids (NEA) capable of being captured using upcoming rocketry for space-based mining. It combines the re-usability of rocket architecture with refueling capabilities to reduce costs. Furthermore, it expands the mission requirements to include a variable retrieval mass, whereas previous work only considered missions that allowed for the entirety of an asteroid to be retrieved. This paper developed a high-level point of view of the mining industry, which proposes to constrain the operation under an estimated amount of investment and logistical cost. Asteroid Retrieval Missions (ARMs) depart from L2 Halo orbit. The time window for the space mining operation is proposed to be three decades, starting in 2030. The logistics of fuel from the Earth's surface to L2 Halo is presented to refueled each ARM. Previously developed data mining online tool that searches for Near-Earth Asteroids (NEA) close approach database from JPL and the small body database from NASA was used as a dataset for possible candidates. The final solution was the retrieval of 60,430 metric tons with the cost per kg of USD 333.44. It reflects an 87.62% reduction in cost per kg compared to previous work.

RESOURCE UTILIZATION ON MARS

Conference Paper

Full-text available

Jul 2020

Cole C. Pazar

Resource utilization of the atmosphere, surface, and upper subsurface of Mars is of utmost importance for scientific exploration, permanent colonization, and the development of sustainable city infrastructure. This presentation demonstrates the current knowledge, systems, technologies, and processes for resource utilization on Mars. The stages of space resource utilization on Mars are: prospecting, acquisition, processing, production, and manufacturing. The six primary resource requirements to sustain human life on Mars are: water, air, metals, power, fuel, and food. For the acquisition of water, there is solid water ice readily available in the subsurface of the Martian Polar Regions, and quite abundant in specific ice-filled craters. Oxygen can be initially acquired by separation of carbon dioxide from processing the Martian atmosphere and surface ices, and can later be partially acquired through biological photosynthesis. The leftover nitrogen can be concentrated and pressurized for the inert component of breathable air. Metals (iron, aluminum, silver, gold, REE's, etc.) for technology manufacturing and construction can be acquired from surface and subsurface deposits, and asteroid mining, utilizing the byproducts of mining for concrete and agriculture. Electrical power generation will be fulfilled by a combination of solar, wind, nuclear, and combustion. Fuel for plastics (syngas) and propellant (methane) can be created via the Sabatier process, and potentially extracted from subsurface reservoirs where they exist. Food production can be enabled through resource processing of Martian soil and acquiring essential minerals for biomass growth. I address the current geologic knowledge of available Martian resources, modern advancements in space systems and robotics, state-of-the-art ISRU technologies, and the construction processes of developing this infrastructure. The costs of this infrastructure are addressed at various scales, and will be funded through asteroid mining and resource supply chains going through the future Mars–Moon–Earth transportation network. Conference dates: October 15-18th 2020 – abstract accepted, presentation time TBD

NASA's Space Launch System: Launch Capability for Lunar Exploration and Transformative Science

Conference Paper

Mar 2020

Stephen D Creech

A surface gravity traverse on Mars indicates low bedrock density at Gale crater

Article

Feb 2019

Teaching Curiosity to do gravimetry Gravimetry—the measurement of tiny changes in gravitational fields—can be used to weigh mountains. Large-scale gravimetric mapping can be done from orbit, but examining small details requires a vehicle on the ground. The Curiosity rover on Mars carries several accelerometers used for routine navigation. Lewis et al. recalibrated these accelerometers to allow them to be used for gravimetry. They measured how the local gravitational field changed as the rover moved through Gale crater and began to climb Aeolis Mons (Mount Sharp). The resulting density of material under Gale crater shows that it is relativity porous, disproving a theory that the crater floor was once buried under several kilometers of rock. Science , this issue p. 535

MegaFlex - The Scaling Potential of UltraFlex Technology

Conference Paper

Apr 2012

David Murphy

Integrated Mars In Situ Propellant Production System

Article

Jan 2013

Pioneer Astronautics has developed a system that harvests CO2 from a simulated Martian atmosphere and reacts it with H2 into rocket propellant at a rate of 1 kg/day. A prototype system operated autonomously for five consecutive days during which it maintained a 100% conversion rate (to detectable limits) of CO2 and H2 into products, maintaining a constant O2:CH4 ratio in the product stream with only minor adjustments. The Integrated Mars In-Situ Propellant Production System (IMISPPS) uses a mixed catalyst bed to conduct the Sabatier methanation reaction and the reverse water gas shift (RWGS) in a single reactor. A recycle loop makes performance resistant to operational changes, with product ratios stable regardless of reactor temperature changes of as much as 100oC. An optimized IMISPPS system is projected to produce 1 kg/day of O2:CH4 propellant and have a mass of 50 kg with a methane purity of 98+% while consuming 700 W of electrical power. With hydrogen from the Earth, the system demonstrates a mass leverage of 18:1 in useful rocket propellant produced on Mars compared to feedstock imported.

Elon Musk Has Finally Predicted The Year Humans Will Land On Mars

May 2022

C Nwali

C. Nwali, "Elon Musk Has Finally Predicted The Year Humans Will Land On Mars," May 2022. [Online]. Available: https://www.futurespaceworld.com/elon-muskhas-finally-predicted-the-year-humans-will-land-on-mars/. [Accessed 30. July 2022].

Making Life Multiplanetary

Sep 2016
27

Spacex

SpaceX, "Making Life Multiplanetary," 27. September 2016. [Online]. Available: https://www.spacex.com/ media/making_life_multiplanetary_2016.pdf. [Accessed 28. September 2021].

Critical Analysis and Review of Current Mars Mission Scenarios for SpaceX Starship

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