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

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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.
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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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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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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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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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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
3.63
Arrival Mars
0.623
1
Total !" outbound flight
4.253
4.63
Start Mars
0.887
6.267
Arrival Earth
1
Total !" return flight
#
0.887
7.267
Table 1: Total dry mass budget of a crewed Starship
Mass (t)
Total mass (t)
Structures
40
101.3
21.3
40
ECLSS
0.27
20.27
20
EPS
2
9
1
4
2
TCS
2.21
10.21
8
Propulsion
12
51.1
0.1
29
5
5
Margins
1.2
8.2
7
Total
200
Payload
25
25
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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
s
=6.267-km
s.
(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*=𝑚0𝑒12
2!
(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*=(3355 +3 380)-s
69.81-m
s)
=3,605-m
s
(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]
73rd International Astronautical Congress, Paris, France. Copyright ©2022 by Mr. Bjarne Westphal and Dr. Volker Maiwald.
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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]
73rd International Astronautical Congress, Paris, France. Copyright ©2022 by Mr. Bjarne Westphal and Dr. Volker Maiwald.
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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
!"
of
$%!"#$"%&%#'' & %(%! '
(
)*
(
+,-./#0
1//#0
2
3)*
(
+,4/+56#0
1//#0
22
& /561-#78
9
(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
!"
of
$%!"#$"%)*%#'' & %(%! ' ()*(+,:14#0
1;4#0 23 )*(<6<5.#0
1;4#0 22
& 1544-#78
9#
$%&'(
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'()** =
?
8"
$#,%&& =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
9#
$%/'(
According to equation (7) converted to
!"
, the booster
with the launch mass
#!
, the mass
#"
after burnout of
1,925 t (1,725 t Starship + 160 t empty mass booster +
30 t propellant for booster landing + 10 t propellant
residuals booster) and an exhaust velocity
"',(
of
3,237.3 m/s according to equation (4) (330 s
7)*
)
produces a velocity change of
Δ𝑣<,,&"%$ =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)
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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
8
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
)=78.2578-t.
(21)
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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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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)
5
8 [50]
20 [51]
4.02
16.8
100
Costs/tonne (US$/t)
32.6 M
43.8 M
43.8 M
15.4 M
5.4 M
0.39 M
73rd International Astronautical Congress, Paris, France. Copyright ©2022 by Mr. Bjarne Westphal and Dr. Volker Maiwald.
Published by the IAF, with permission and released to the IAF to publish in all forms.
IAC-22-A5.2.3 Page 13 of 16
p. 4]. Four tanker flights are thus needed to refuel the
crewed Starship. The propellant costs are calculated from
the price of the propellant of 168 US$/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$)
230 M
130 M
200 M
122.36 M
68 M
107.4 M
Lifetime launches
1,000
100
12
100
100
10
Starts/Mars trip
6
5
1
5
4
1
Average maintenance costs/flight)
0.2 M
0.5 M
10 M
0.11 M
0.28 M
5.63 M
Propellant costs/launch (US$)
1.13 M
0.42 M
0.33 M
0.6 M
0.28 M
0.25 M
Launch site costs/launch (US$)
0.2 M
0.2 M
Total costs/Mars trip (US$)
11 M
8 M
43 M
13 M
6 M
20 M
73rd International Astronautical Congress, Paris, France. Copyright ©2022 by Mr. Bjarne Westphal and Dr. Volker Maiwald.
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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.
73rd International Astronautical Congress, Paris, France. Copyright ©2022 by Mr. Bjarne Westphal and Dr. Volker Maiwald.
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IAC-22-A5.2.3 Page 15 of 16
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... Nevertheless, on-orbit refueling will increase significantly the overall mission cost. As each launch will cost 200,000$ for the launch site (Musk, 2017), the 5 tanker launches required for each mission (Westphal & Maiwald, 2022) will have an additional 1,000,000$ of cost. These extra tankers would also increase the total cost of building extra launching systems, although Super Heavy and Tanker Starship have a life expectancy of 100 flights due to short distance travel (Westphal & Maiwald, 2022). ...
... As each launch will cost 200,000$ for the launch site (Musk, 2017), the 5 tanker launches required for each mission (Westphal & Maiwald, 2022) will have an additional 1,000,000$ of cost. These extra tankers would also increase the total cost of building extra launching systems, although Super Heavy and Tanker Starship have a life expectancy of 100 flights due to short distance travel (Westphal & Maiwald, 2022). Thus, the cost per mission of CLR will ultimately surge. ...
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Elon Musk Has Finally Predicted The Year Humans Will Land On Mars
  • 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
  • Spacex
SpaceX, "Making Life Multiplanetary," 27. September 2016. [Online]. Available: https://www.spacex.com/ media/making_life_multiplanetary_2016.pdf. [Accessed 28. September 2021].