Mission Design of a Two-Person Mars Flyby by 2018
International Student Design Competition
Team Mars18 - www.mars18.de
Margret Barkmeyer
Winfried Burger
Felix Düver
Eduardo Finkenwerder
Dan Fries*
Stefan Fuggmann
Sören Heizmann
Christina Herr
Mirjam Schmidt
Rolf Stierle
Lukas Teichmann
Tobias Torgau
Daniel Wischert
Ferdinand Leinbach
Victor Mosmann
Fabian Müller
Paul Nizenkov
Duncan Ohno
Adrian Pfeifle
Minas Salib
Marcel Scherrmann
Nils Hoffrogge
Heiko Joos
Peter Jüstel
Jochen Keppler
Ronja Keuper
Alexander Kunze
Jonas Lay
Hong Anh Le
University of Stuttgart
Table of Contents
Page
1 Introduction
1
2 Executive Summary
2
3 Mission Architecture
3.1 Trajectory . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.2 Launcher Selection & Manifest . . . . . . . . . . . . . . . . . . . . . . . . . .
3.3 Reentry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3
3
4
7
4 Human Factors
4.1 Astronaut Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
4.2 Crew Health . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
11
11
12
5 Spacecraft Design
5.1 Configuration & Structure . . .
5.2 Subsystems Design . . . . . . .
5.3 Scientific Payload . . . . . . . .
5.4 Systems Engineering & Budgets
.
.
.
.
15
15
17
35
36
6 Programmatic Issues
6.1 Cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.2 Roadmap & Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
6.3 Risk Management . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
38
38
41
43
7 Conclusion
46
Bibliography
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*Point of Contact: Dan Fries, [email protected]
1 Introduction
To increase the pace in manned Mars exploration, the Mars Society in collaboration with
Dennis Tito’s Inspiration Mars Initiative called for students around the world to develop a
complete mission concept for a manned Mars flyby in 2018. The ultimate goal is not only
to complete this very ambitious mission but to spark more interest in comparable missions
around the world and further a technological competition to put a human being on Mars on
a peaceful level for the greater benefit of mankind.
Previous efforts are limited to robotic missions and so far not even a sample return mission
has been accomplished. While robotic missions are certainly sufficient to gather simple,
pre-determined scientific data, they are not fit to actually expand humanity’s sphere of
influence. In-situ research and human settlements, however, enable access to resources and
hold future potential on a completely different scale. Future problems, like the overpopulation
of Earth and lack of essential resources like water, could be tackled right now. Even in
the present day, solar system exploration efforts would immediately result in a myriad of
scientific and technological developments, employment opportunities and eventually direct
financial gain. Furthermore, private space companies and NASA are already working towards
heavy-lift launch vehicles, but so far without a clear application. All of the factors mentioned
above contribute to the overall result that manned space exploration is not only desirable
but also achievable.
Mars18 is the student-led team at the University of Stuttgart, Germany, that has taken
on the challenge proposed by the Mars Society and Inspiration Mars. The team’s goals are
the meaningful contribution to the worldwide efforts in space exploration, education of highpotential students in a hands-on project, engagement of public interest in space exploration
through cooperation with local media and achieving a high ranking in an internationally
acclaimed competition. As a mission like this has never been attempted before, the design
presents a special challenge that also sparks creativity in every person involved. The team
consists of about 29 students, most of whom are aerospace engineers. But it is obvious that
such a project cannot be successful solely through aerospace technology. From the beginning,
the team attempted to achieve a multi-disciplinary composition of motivated participants.
Mars18 is comprised of students from medical sciences, social sciences, electrical engineering
and economics. Moreover, the team managed to gain professional support from the Institute
for Space Systems (University of Stuttgart), Astos Solutions, Constellation (distributed
platform for aerospace research) and Airbus Defense & Space.
The presented work attempts to show how a manned mission to Mars could be executed
realistically by 2018. In general, conservative assumptions were preferred over optimistic ones,
in both technological and cost issues. Key technologies that would further access to space in
general and for this specific mission were identified and a time schedule developed that would
allow for their implementation. Although a certain amount of technologies is employed that
have to be qualified, this is only done in absolutely necessary cases or because it presents a
considerable advantage. Human factors were evaluated and accounted for. Finally a complete
cost estimate was conducted.
1
2 Executive Summary
During the entire development, the Mars18 team followed four principles: simplicity, safety,
low cost and feasibility. To evaluate the amount of development still required an estimated
Technology Readiness Level (TRL) is used.
Through rigorous optimization and evaluation of available systems it is possible to lower
the total mass below 15 t and the low Earth orbit (LEO) mass amounts to ∼ 63 t. A concept
is devised that allows to launch the entire mass with only two starts of currently or soon to be
available carrier rockets. Trans-Mars injection (TMI) is accomplished via staged propulsion
of two modified Delta IV 4-m Second-Stages.
The system consists of modified versions of
the Enhanced Cygnus (referred to as Cygnus)
and the DragonRider (referred to as Dragon).
Both modules have already been tested in their
basic configuration and are currently under further development. An important mission like
this requires absolute priority among the deep
space communication systems on Earth (i.e. Deep
Space Network (DSN), ESTRACK). To reduce
mass several of the AOCS’ thrusters are resistojets, using waste products from the life support
system. Additionally, a fuel saving model predictive control (MPC) algorithm is employed for attitude control. To account for the dangers
of radiation exposure outside of the Earth’s magnetosphere, a protection scheme is devised
that works highly synergetic with the equipment of other subsystems and provides a storm
shelter for solar particle events (SPE). The life support system is designed from scratch and
follows a virtually close-loop approach. It also introduces two devices that have not been
used on previous missions but are able to reduce the required initial mass considerably and
increase the synergies with other systems. As the flight trajectory, a free-return option is
chosen that requires ∼ 4.8 km/s from a 350 km LEO. At return to Earth’s atmosphere, the
reentry capsule will have a relative velocity of ∼ 13.8 km/s. The resulting kinetic energy is
dissipated during two passes through the atmosphere before descending to the ground.
The total mission duration from TMI is ∼ 501 days. The presented concept to deal with
human factors handles physical as well as psychological issues in a very isolated and confined
space. The well-being is of great importance, since they should perform experiments and
document as many things as possible during their journey. Thus, the endeavor will result in
the maximum scientific benefit for future missions and possible spin-offs.
The total cost of such a mission is estimated using current prices, heritage data and cost
models. Thus providing an amount of 4.3 B$ that should be around an upper limit for the
presented design. In a simple risk analysis not only the dangers stemming from technological
aspects but also programmatic issues are presented and how they might be mitigated.
2
3 Mission Architecture
The approach of choosing the most favorable trajectory is presented in this chapter. Furthermore, available launcher systems are compared in detail and finally reentry is discussed.
3.1 Trajectory
The objective of the trajectory optimization is to find a feasible flyby trajectory to Mars.
By solving the Lambert problem, options for such a trajectory are investigated. Therefore,
astrodynamics and interplanetary spaceflight have to be considered.
3.1.1 Tools & Boundary Conditions
POINT is a Lambert-solver by Astos Solutions. Ephemerides provided by Jet Propulsion
Laboratory (JPL) are used to solve the Lambert problem, providing an accurate initial
estimation. POINT determines and optimizes trajectories based on constraints and boundary
conditions. Possible optimization constraints are minimum flight time as well as low departure
and arrival C3 energies. In the following, different trajectories are compared according to
mission requirements. GMAT (General Mission Analysis Tool) is an open-source space
mission analysis tool provided by NASA. It enables the simulation of gravitational forces of
all celestial bodies in the solar system. GMAT is used for verification of the final trajectory.
The following boundary conditions constrain the trajectory design. Launchers starting
from Cape Canaveral will lift the spacecraft into a circular LEO. The assembly orbit is
at an altitude of 400 km and likely to decrease due to atmospheric drag. Therefore, the
interplanetary trajectory is set to start at a 350 km altitude with an inclination of 28.5° in
the equatorial plane. Additional mission requirements are a flyby altitude of 100 km at Mars
to prevent aerodynamic drag and to start the mission in 2018.
3.1.2 Trajectory Trade-Off
Multiple possible trajectories were found with POINT. The focus is on the ∆v at Earth and
Mars, total mission time and excess velocity v∞ for arrival at Earth. With fixed mission costs,
higher ∆v leads to a smaller throw mass. Shorter mission duration leads to higher v∞ for
return to Earth, which in turn significantly increases the loads and stresses on the thermal
protection system. Since all investigated trajectories result in v∞ higher than at every reentry
that has been done, trajectories with low v∞ are favored. As all mission concepts have a
duration of at least one year but only those shorter than 120 days have a significant impact
on the Environmental Control and Life Support System (ECLSS) design, it did not influence
the trade-off. Although, it can be concluded that a shorter mission duration decreases the
total system mass. Thus, a trade-off is necessary between mission duration, total mission
mass, the resulting ∆v as well as v∞ upon arrival at Earth. Three suitable trajectories found
with POINT are presented in Table 3.1.
The first trajectory with a start date in May 2018 and a duration of 435 days requires a
∆v of 1560 m/s at Mars and has a v∞ of 11 km/s on return to Earth. However, an additional
3
3 Mission Architecture
3.2 Launcher Selection & Manifest
Table 3.1 Considered Earth-Mars trajectories found with POINT
Start Date - Arrival Date ∆v Departure ∆v Flyby v∞ Arrival Duration
1
2
3
02/05/2018 - 07/12/2019
07/15/2018 - 07/14/2019
01/04/2018 - 05/19/2019
3586.2 m/s
9342.8 m/s
4825.5 m/s
1558.6 m/s 11 070.1 m/s
0
12 835.0 m/s
0
8786.3 m/s
435 d
361 d
501 d
propulsion system is required to provide a ∆v at Mars, which increases system risk and adds
a single point of failure to the mission. To conclude, the decrease in flight duration does not
justify the additional propulsion module.
The second trajectory offers a shorter
flight time, but the required ∆v is not feasible. The third trajectory is a good trade-off
Launch
Earth Orbit
between ∆v, duration and return velocity.
The detailed data for this free-return trajectory verified with GMAT can be found in
Table 3.2. A C3 of 38.7 km2 /s2 results in a
Sun
v∞ of 6220.7 m/s at TMI. For a circular departure orbit of 350 km altitude, with a velocity of 7702.0 m/s and a hyperbolic velocVenus Orbit
ity at the same altitude of 12 543.4 m/s, the
Reentry
required ∆v for TMI is 4825.5 m/s. This
Flyby
change in velocity has to be provided by
Mars Orbit
the propulsion system. No further impulFigure 3.1 Planned free-return trajectory
sive maneuver is required to return to Earth
safely after a flight duration of 501 days. A
schematic of the selected trajectory is presented in Figure 3.1.
Table 3.2 Free-return solution: right ascension (RLA) and declination (DLA) of the outgoing
hyperbolic asymptote relative to the Earth/MarsMJ2000Ec-frame of GMAT
Date
Departure Earth 04.01.2018 18:59:35.501
Mars flyby
20.08.2018 18:31:38.676
Arrival Earth
20.05.2019 16:28:42.508
v∞ [m/s]
6220.7
5375.0
8855.7
DLA
RLA
−89.3◦ 16.6◦
−1.53◦ −121.9◦
3.6◦
−72.9◦
vperi [m/s] C3 [km2 /s2 ]
12543.4
7258.1
14186.9
38.7
28.9
78.4
3.2 Launcher Selection & Manifest
Launchers are responsible for launching the Mars Transfer Vehicle (MTV) and propulsion
module (PM) into the defined parking orbit in LEO, where the TMI will be carried out.
3.2.1 Design Process & Requirements
In general, a launch concept with the least launches is preferable in order to keep down
mission and system complexity while increasing crew and mission safety accordingly. This
requirement, however, demands both a man-rated launch vehicle which is capable of launching
the MTV as well as a second launch vehicle capable of launching the PM.
4
3 Mission Architecture
3.2 Launcher Selection & Manifest
In order to find the ideal launch concept for the mission, carrier rockets from medium- to
heavy-lift of the last 60 years were systematically analyzed and the results can be found in the
appendix [36]. Launchers, which fulfill the criteria of high payload mass and high reliability,
were taken into further consideration. A smaller subset of launch vehicles will be man-rated
in 2018 and most of them are not able to launch the complete MTV by a single launch. The
collected information was used to develop different concepts of launcher combinations with
the respective payloads. The most promising concepts were further analyzed. Due to the fact
that there is no launch vehicle in 2018 which is man-rated, flight-proven and able to launch
the PM and the MTV by a single launch, a concept with two launchers is chosen.
With only two launches, the risk for the crew is minimized in addition to ensuring low cost
and high mission safety. One lighter, man-rated and highly reliable launch vehicle to carry
the crew is used and another to carry the PM. This concept is preferred due to its reduction
of complexity for rendezvous maneuvers, lower boil-off of propellant and the design lifetime of
the electrical power system (EPS) and attitude and orbit control system (AOCS) of the PM.
3.2.2 Final Launch Concept
The PM is launched with a slightly modified Falcon
Table 3.3 Payload distribution
Heavy on December 21, 2017 from Launch Complex 39
Launcher
Payload Mass
at Kennedy Space Center (KSC). The crewed MTV is
launched on the December 24, 2017 by an Atlas V 441 Falcon Heavy PM
48 027 kg
from the planned launch complex for manned Atlas V
Atlas V 441
MTV
15 000 kg
missions also from KSC. The payload distribution is detailed in Table 3.3. All launches from KSC during the past five years in December and
January were analyzed and it can be concluded that the weather conditions are not a critical
factor for the launch window, whereas technical malfunctions can indeed be a critical variable.
Therefore, three launch possibilities for each vehicle with an interval of one day for the Falcon
Heavy and a three day interval for the Atlas V 441 are scheduled. The begin of each launch is
timed to achieve the right orbit parameters for parking orbit. After the PM has reached the
orbit, it transforms into the injection configuration and acquires a gravity-gradient stabilized
position to minimize the required fuel for the AOCS. The change of the configuration is
further explained in Section 5.2.5. Once the MTV reaches LEO, the crewed Dragon capsule
detaches from the Dragon trunk, flips over and docks to the docking port of the Dragon
trunk. After orbital alignment is arranged by the MTV within 7 h, it docks to the PM. After
assembly completion, there are at least two days to test all systems and to prepare the MTV
for TMI. Figure 3.2 illustrates the launches and on-orbit assembly.
On January 4, 2018, the first stage of the PM is ignited to achieve a high elliptical orbit,
which provides the correct Keplerian elements of the Mars transfer orbit for ignition of the
second stage. As burnout of the first stage is complete, the MTV and second stage separate
and perform trajectory corrections at the apogee. At perigee, the second stage is ignited and
the final TMI is performed. Both ignitions produce a maximum acceleration of 0.5 g, due
to the relatively low engine thrust of 110 kN. This acceleration is easily tolerable and was
already proven and tested during the Gemini 10 mission.
In order to validate the required modifications of the launch vehicles, slight changes in
already scheduled missions are planned. Firstly, negotiations have to be conducted with the
5
3 Mission Architecture
3.2 Launcher Selection & Manifest
Figure 3.2 Batchart of launch and assembly
U.S. Department of Defense or Intelsat to perform their planned Falcon Heavy launches with
the modified Asymmetric Payload Fairing (APLF) in 2015 and 2017, respectively. Therefore,
it is intended to pay 75% of the launch costs to the respective customer, which is also
considered in the overall mission costs. Secondly, a test rendezvous is planned between the
Cygnus module and the Dragon at COTS-6 from SpaceX and COTS-7 from Orbital Science.
The Cygnus module stays docked at the ISS until the manned Dragon also docks with the
ISS. After undocking of Cygnus and Dragon, they will perform a rendezvous maneuver before
their reentry to validate the modified docking port.
Moreover, the proposed launch concept takes full advantage of the Commercial Crew
Development (CCDev) program, which is initiated by the U.S. government and promotes the
upcoming private space industry in the U.S. For the improbable case that it is not possible for
SpaceX to modify the Falcon Heavy with the APLF by 2018, an alternative launch concept is
provided. This backup intends to replace the Falcon Heavy with an Atlas HLV, which carries
a Delta IV 5-m Second-Stage, and a Delta IV Heavy carrying a Delta IV 4-m Second-Stage.
These stages are modified to be able to perform a docking maneuver. Further information
about this alternative launch concept can be found in the appendix [36].
3.2.3 First Launch (Falcon Heavy)
To execute the mission with as few launches as possible, usage
Table 3.4 Heavy-lift
of heavy-lift launchers is necessary. An initial evaluation was
launcher trade-off
performed during the design process, after which several launch
Payload
vehicles were selected for further consideration as shown in Table Launcher
to LEO
3.4. The most promising launch vehicle is the Space Launch
61–81 t
System (SLS). However, as the first launch is scheduled as late SLS
Falcon
Heavy
53 t
as 2017, this mission would be its maiden flight. Therefore, and
Angara A5
24.5 t
because of the high cost, it is not considered for the launch concept.
Delta IV Heavy
23 t
This is why SpaceXs Falcon Heavy is utilized. It will have its Ariane 5
21.5 t
first launch in 2014 and, hence, has a sufficient trial period. To Proton
20 t
launch the PM with the Falcon Heavy, it is necessary to enlarge
6
3 Mission Architecture
3.3 Reentry
the payload fairing so that it offers enough space for the PM. An aerodynamically optimized
payload fairing is proposed, which keeps the aerodynamic loads equal and increases the
weight only by 30% compared to the default fairing [44]. The used APLF is 18.5 m tall and
9.5 m wide and is able to host the PM in its launch configuration. In addition, the fairing is
equipped with ventilation in order to release the escaping hydrogen of the cryogenic tanks
of the PM. Furthermore, the upper stage of the Falcon Heavy is equipped with additional
thrusters and more fuel for its AOCS to be capable of controlling the changed center of
gravity of the payload. To account for these modifications, such as the increased structure
mass and the parking orbit above the 200 km reference orbit given by the launch provider, a
10% margin on the overall payload capacity is included in the calculation.
3.2.4 Second Launch (Atlas V 441)
As shown in Table 3.5, only a few launchers are man-rated and
Table 3.5 Man-rated
capable of launching enough payload to LEO. Soyuz-FG is fully
launcher trade-off
developed and flight-proven in multiple missions but can only carry
Payload
light payloads. According to SpaceX, Falcon 9 will be man-rated by
Launcher
to LEO
2016. Additionally, Falcon 9 is cheaper, but Atlas V is able to carry
SLS
61–81 t
heavier payloads. The Atlas V family is flight-proven. It is one of
Atlas
V
8.9–17.7
t
the most reliable launch vehicles in the world and every U.S. mission
Falcon 9
11 t
to Mars in the last decade was launched by an Atlas V. In addition,
Soyuz-FG
7.1 t
Boeing plans to launch their CST-100 capsule in 2016 with an Atlas
V and also the Sierra Nevada Corporation plans to launch their DreamChaser in November
2016 with an Atlas V. Therefore, Atlas V is man-rated and extensively tested with multiple
crewed capsules until 2018. The Atlas V family can be combined in different ways and the
chosen launcher will be an Atlas V 441. It uses four solid rocket boosters and a modified four
meter payload fairing. In addition, the upper stage of the Atlas V 500 series is utilized to
account for the heavier payload and its shifted center of gravity. To host the MTV, the fairing
is modified to cover the Cygnus only. The upper part of fairing is removed and replaced by
the Dragon capsule, so it carries the weight of the Dragon. With this modification, Cygnus is
not carrying the weight of the Dragon at launch and SpaceXs launch abort system, which is
integrated in the Dragon, can be used. In order to take account of these modifications, a
margin of 8.5% on the payload capacity of Atlas V 441 is included in the calculation.
3.3 Reentry
Two main factors impose restrictions on the reentry: Loads must be kept below 8 g and the
duration must not exceed 14 h. The time constraint is due to limited battery capacity and
heat sink, which is used to dissipate waste heat after detachment from the trunk and the
Cygnus module.
Considered Reentry Scenarios
The reentry maneuver is challenging and, if not altered, will be the fastest manned reentry.
Several efforts and ideas to reduce the speed were considered:
7
3 Mission Architecture
3.3 Reentry
ˆ Slowing down the spacecraft with an additional propulsion system
ˆ Capturing the spacecraft and slowing it down with a pre-deployed capture vehicle
ˆ Using the Moon for a gravity assist maneuver to slow down the spacecraft
ˆ Aerobraking with multiple passes before reentry
ˆ Using a spring/tether mechanism to achieve a slowdown of the vehicle (quickly discarded
due to high mass/∆v ratios)
The first and second option exploit the same idea: employ propulsion systems to slow
down the spacecraft. For the first option, only solid rocket boosters are feasible due to the
boil-off of cryogenic propellant and a relatively low Isp of storable propellant. This results in
a drastically increased mass of the propulsion module and, thus, in high launch cost. While
the second option could utilize a pre-deployed propulsion module with electric propulsion, it
increases mission complexity and requires an additional launch as well as the development
of an additional module. Also, both options suffer from a single point of failure issue: An
engine failure and/or a missed capture/rendezvous would lead to a loss of crew (LOC) as
well as loss of mission (LOM).
The gravity assist option at the Moon is not possible since its position is on the opposite
side of Earth at the time of arrival for the selected trajectory. Therefore, only aerobraking is
further investigated. To approximate aerobraking reentry scenarios, the tool ASTOS from
Astos Solutions was used. ASTOS is a software package to simulate and optimize launch and
reentry trajectories.
Assumptions of the Reentry Trajectory
Reentry loads depend on a range of factors, including spacecraft geometry and trajectory. Although the latter is known, geometry data is unavailable due to proprietary issues. Important
geometric characteristics of a reentry probe are: cD /cL values (and the respective ballistic
coefficient) and nose radius. The Apollo capsule [41] serves as a reference for cD /cL values
since it is similar in size, weight and heat shield diameter. The cD /cL values are considered as
a function of altitude at constant velocity, since for constant altitudes and varying velocities
aerodynamic values do not change significantly. To account for deviations from these values,
a sensitivity study was conducted. The study demonstrated that variations of cD /cL within a
reasonable range only have a minor impact on the reentry trajectory. For every combination
of tested cD /cL values within the range of the Apollo values, a viable reentry trajectory could
be found. Graphical approximations from drawings in the official DragonLab fact sheet led
to estimations of the nose radius. From these, a calotte was calculated with a nose radius of
r = 4.7 m. The respective data sheet and sensitivity study for the aerodynamic coefficients
can be found in the appendix [36].
The heat flux consists of the convective heat flux and radiative heat flux. For all trajectory
calculations the 1976 US Standard Atmospheric Model is used. Convective heat flux is
covered by the Chapman heat flux model, which is given by
q̇conv = C ρn V m ,
8
(3.1)
3 Mission Architecture
3.3 Reentry
where C = 1.705 × 10−4 , n = 0.5 and m = 3 are model parameters. ρ and V denote the air
density and velocity, respectively. This model yields good estimates for convective heat flux
and is within a margin of error verified for a wide range of existing missions.
Estimating the radiative heat flux is more challenging. There are two widely known
radiative heat flux models, the Tauber-Sutton model [53] and the Detra-Hidalgo model [14].
The former is known to be fairly accurate, but is only applicable for nose radii between
0.3 m < r < 3 m. The determined nose radius is therefore not within the applicable range. A
study and verification of both models was performed for the Stardust Return Capsule. The
nose radius was varied over a wide range to show steadiness and applicability. The result
suggests using the Tauber-Sutton model for the approximation. The heat flux calculated by
Tauber-Sutton for the investigated nose radii is larger by one order of magnitude than the
results of Detra-Hidalgo. Furthermore, the heat flux during reentry of Stardust computed
with the model of Tauber-Sutton compares well with literature values [34]. Finally, it is
demonstrated that by using Tauber-Sutton out of range of its applicability, ”the resulting
error is still generally within the range of uncertainty found in computing the radiative
heating with a more computationally intense method” [13].
Results of the Reentry Simulation
10
7500
Altitude
Velocity
5000
5
2500
0
0
1
2
3
4
5
0
Time [h]
Load factor [g]
Altitude [km]
10000
Velocity [km/s]
Based on the final reentry mass of 3900 kg and self-imposed constraints (load factors less
than 8 g and total reentry time less than 14 h), a path analysis was performed. The altitude
and velocity as well as the load factor over time are shown in Figure 3.3.
8
15
12500
6
4
2
0
0
1
2
3
4
5
Time [h]
(a) Velocity and altitude
(b) Load factor
Figure 3.3 Reentry at perigee altitude and bank angle of: 63.5 km and 98.5◦ (first pass), 65 km
and 65◦ (second pass)
Figure 3.4 presents the total heat load integrated over time and the heat flux. A worst-case
scenario at the lower reentry window at 60 km is assumed here, which differs from the nominal
trajectory with three atmospheric passes during aerobraking.
Moreover, it was found that a precise control of the bank angle µ is crucial. Hence, keeping
the bank angle within a certain margin during reentry is an important requirement for the
AOCS. Generally, it can be said that higher perigee altitudes require higher bank angles
9
3 Mission Architecture
3.3 Reentry
1000
3
800
101
Heat Flux
Heat Load
600
10−1
400
200
10−3
0
1
2
3
4
5
6
7
8
9
10
11
Heat Load [MJ/m2 ]
Heat Flux [W/cm2 ]
10
0
12
Time [h]
Figure 3.4 Thermal stresses for perigee altitude of 60 km and bank angle of 79.5◦ (worst-case)
(thus shifting the lift vector downwards) and result in lower load factors and heat flux. At
the upper and lower ends of the reentry window, the flight path is more sensitive to the
bank angle. At a perigee altitude of 63.5 km, the bank angle leading to a reentry within
the constraints is in the range of ∆µ = 8.5◦ . This is the favored trade-off since it is a local
maximum of the bank angle range at these perigee altitudes. At the upper limit of 71.45 km
with a bank angle of 180◦ , it is possible to shift to left and right and thus double the range
of the bank angle. However, the gradient of the bank angle range is very steep. At 71 km,
the range drops to ∆µ = 2◦ . Therefore, it is not recommended to enter at this altitude.
Fluctuations of the atmosphere or inaccuracies of the predicted trajectory would lead to a
reentry outside of the constraints. Finally, with bank angle adjustments during the second
pass, a water landing can be guaranteed. Table 3.6 shows at which perigee altitude the bank
angle has the largest margin. Results of the analysis can be found in the appendix [36].
Table 3.6 Reentry window (bank angle for first pass, 0◦ for following passes)
Perigee Altitude
Upper window
Selected trajectory
Lower window
Bank Angle
Load Factors
71.45 km 174.5◦ – 185.5◦ 4.3 – 4.3 g
63.5 km
94◦ – 102.5◦ 6.9 – 7.9 g
60 km 79.5◦ – 80◦
8.0 g
10
Duration
9.7 – 14 h
1.3 – 14 h
12.1 – 13 h
4 Human Factors
The primary focus of this section is on the challenges the human health system will face
in a microgravity environment. Due to effects of microgravity on the physiological system,
solutions are needed to ensure physical health during the whole flight. This includes solutions
for the cardiovascular system and bone loss problem in space. To ensure medical care for
all potential situations, suitable preventative measures, medication, a systematic training,
nutrition protocol and electronic monitoring have been developed. Besides physical health,
the focus is also on the mental fitness of the astronauts. Mental health is as necessary and
essential for a successful mission as medical care. The concept is outlined in Figure 4.1.
2
Ensure physical health
The crew is prepared for all medical
risks and is provided with medical
treatment possibilities.
3 E-Health
24/7 monitoring and
documentation of medical
parameters through a
health vest and Microflow.
1 Preselecting & Preparation
Criteria for the preselection (age,
experience, profession,..) are set up.
Moreover, the astronauts have to be
prepared mentally and physically.
4
Training & Food
To prevent muscle degradation due
to microgravity, training equipment
and a suitable nutritional protocol is
provided.
5 Ensure mental health
Mental health is ensured by using
audio-visual stimulation, a motivation
and entertainment kit as well as a daily
schedule.
Figure 4.1 Health care concept
4.1 Astronaut Selection
The crew consists of a male and a female astronaut, which have both successfully completed
their astronaut training. In January 2016, the crew selection process will start to pick one
primary and one backup crew for the mission. Preparations and training procedures, which
are especially developed for this mission, should start 18 months before the launch. At this
time, the crew selection process will be finished and the astronauts will start to specialize
their training routine.
Mental disorders such as anxiety, post-traumatic stress, insomnia or depression can develop
unexpectedly. A study has indicated that the average age of onset depression for healthy
people is 41 years [55], so the age of selected astronauts will range from 26 to 37 years. Some
crucial factors for astronaut selection are listed below.
ˆ Age ≤ 37 years
ˆ University degree (or equivalent)
ˆ Free from any disease, any dependency
on drugs, alcohol or tobacco
ˆ No intensified radiation exposure until
launch
11
4 Human Factors
4.2 Crew Health
ˆ Normal range of motion and functionality in all joints
ˆ Visual acuity in both eyes
ˆ Curiosity
ˆ Ability to trust
ˆ Creativity / resourcefulness
ˆ Free from any psychiatric disorders
ˆ Resiliency
ˆ Blood pressure below 140/90
ˆ Adaptability
ˆ Standing height 1.57 m ≤ h ≤ 1.91 m
4.2 Crew Health
4.2.1 Physiological Challenges of Microgravity
Human physiology adapts to microgravity, which sets off changes to the human body. When
considering physiological risks during the mission, some of the major problems occur because
of altered blood composition. These pathological changes are well-known and treatable in
their appearance on Earth [58]. According to data from previous missions (e.g. Gemini or
Skylab), the erythrocyte and blood plasma ratio decreases daily by up to 1% and blood
pressure drops by 10 mmHg after three to six months in microgravity. These effects are due
to decreasing vascular resistance. On the other hand, the autonomic nervous system reacts
to microgravity conditions and performs counter-regulation, which also leads to increased
cardiac output [58].
Consequences of microgravity can lead to various diseases. Due to a loss of blood plasma,
there are comparatively too many thrombocytes in the blood, which can lead to pulmonary
embolisms. With decrease of hematocrit, total blood volume drops and can thereby carry less
oxygen. On account of these effects, blood viscosity can fluctuate in both directions. With
too few erythrocytes, which carry oxygen through the body, an under-supply of the brain
occurs, which inflicts irreversible damage after two minutes, fainting and eventual death. In
addition to diseases due to the cardiovascular system, there are numerous others. Some of
them are temporary, like the space adaptation syndrome, which are not further discussed.
For the problems mentioned above, several meaTable 4.1 Human factors summary
sures are taken to control the cardiovascular system
Mass Volume
and replace missing blood components. Measures
consist of a combination of diet, physical exercise
Astronauts
189 kg
0.2 m3
and medication. A specific diet plan with dietary Launch entry suits
40 kg
0.5 m3
0.3 m3
supplements controls the iron and vitamin deficit. Personal items & drugs 12 kg
1.1 m3
A positive effect on blood pressure and blood circu- Medical & biofeedback 24 kg
Training devices
158 kg
1.2 m3
lation can be achieved through a low-sodium diet.
Total
423 kg
3.3 m3
Only in the case of acute changes or pathological
concerns should the latter dietary supplements be
taken. This is detailed in Section 5.2.4. With a purpose-built training schedule and modified
exercises, astronauts counteract the inevitable reduction of bone and muscle mass. Positive
side effects of these workouts are maintenance of the cardiovascular system and physical
fitness of the crew. Furthermore, it supports the muscle-vein-pump, which helps to ensure
oxygen supply. To get used to these exercises, the crew should start purpose-built training
at least six months prior to launch. If abnormal changes or problems occur during one of
12
4 Human Factors
4.2 Crew Health
the monitoring sessions, drug treatment to ensure the health of the astronauts should be a
last resort. In this case, in addition to dietary supplements, the crew has access to drugs
for hypertension, osteoarthritis, arteriosclerosis or anticoagulants to prevent thrombosis (e.g.
iron, calcium, magnesium, etc.). Drugs are administered orally in order to avoid wounds.
Through this combination, the crew is prepared for all foreseeable eventualities.
4.2.2 Physical Training
Bone Mass Density[%]
The musculoskeletal system is mostly unencum100
bered, but it disintegrates and loses tone and mass
during a space mission. Because of loss of minerals,
90
bones become weaker and trunk and leg strength
decreases by 10% to 20% [31]. Figure 4.2 shows an
80
age-related bone mineral density (BMD) loss rate
combining men and women. It is based on a simulation, which is strongly correlated to the combined
70
osteoporotic bone loss in BMD.
Therefore, a specially-designed workout concept
60
10
20
30
40
is used to keep bone and muscle loss within a reasonTime in Space [m]
able limit. After completing cardiovascular training
with an ergometer, the crew has to practice with a
Figure 4.2 Predicted bone loss [48]
combination of barbells and bungee cords. By using
different resistance, intensity of the workouts can be varied.
4.2.3 Monitoring and E-Health System
To monitor crew health and physical condition, important vital parameters are checked
several times a day. Because a of lack of erythrocytes and other aforementioned problems,
viral infections can be reactivated more easily than on Earth. Without any medical staff on
board, the crew has to be able to make medical decisions independently and provide mission
control with health data. Therefore, two devices for data collection are planned.
One is the LTMS-3 vest by CSEM in coorporation with ESA. The vest takes pulse, measures
electrocardiography, blood pressure, respiration rate and body core temperature without
affecting the astronauts in a negative way. The crew wears their vests at least three times a
day for one hour to guarantee sufficient monitoring. These vests are currently used in clinical
tests at Concordia Base in Antarctica and will have at least TRL 8 by 2018 [16].
Additionally, Microflow [11] is employed. This technology analyzes blood in real time
and detects infections. It is also able of rapid viral or bacterial identification. Furthermore,
Microflow is able to analyze radiation exposure and stress level of the crew by evaluating
blood consistency. These measurements are taken once a week. Monitoring data from the
LTMS-3 vest and Microflow is sent to Earth frequently. Hence, the medical team is able to
notice conspicuous variations sufficiently early.
13
4 Human Factors
4.2 Crew Health
4.2.4 Mental Health - A Risk Reducing Strategy
The success of human space flight depends on astroTable 4.2 Daily schedule
nauts remaining mentally healthy in order to survive
Time
social isolation and extreme physical environments. Activity
It is crucial for the crew to remain alert and vigi- Getting up/hygiene
8:00 am
lant while operating complex equipment. Therefore, 1. Meal (0.5 h)
8:30 am
9:00 am
getting enough sleep is an important factor. Mi- Mental activity (2h)
Medical
measurement
(0.5
h)
11:00
am
crogravity, noise, vibrations and loss of the natural
Sporting activity (1.5h)
11:30 am
day-night cycle make sleeping difficult in space [35].
1:00 pm
Furthermore, studies have shown that irregular work 2. Meal (0.5h)
Medical measurement (0.5 h)
1:30 pm
and day schedules, high workloads and varying envi- Brainwave entertainment (1h) 2:00 pm
ronmental factors have negative effects on sleep and Entertainment (2h)
3:00 pm
crew performance [35]. Moreover, carbon dioxide and Sporting activity (1.5h)
5:00 pm
6:30 pm
radiation are other factors which negatively impact Medical measurement (0.5 h)
neurobehavior and performance. To avoid these neg- 3. Meal (0.5h)
7:00 pm
Privacy
time
(0.5h)
7:30
pm
ative effects as much as possible, a well-structured
Video
call,
scientific
work
(2h)
8:00
pm
daily schedule is established and depicted in Table 4.2.
Hygiene/sleep
10:00 pm
This schedule provides the necessary variety of sleep,
work, physical training and privacy. It will ensure a specific rhythm during this long-term
mission to prevent shifting of the sleep-wake cycle. A daily routine also prevents muscle
atrophy and bone loss. However, after some weeks in space, when everything has settled
into a routine, issues of boredom and monotony may appear. To prevent such situations, the
astronauts must be given challenging, meaningful tasks [25]. Meaningful work is consistently
correlated with psychological benefits, as it is associated with overall well-being [20].
Another issue is offering as much privacy as possible within limited space. The astronauts
are allowed to take < 5 kg personal items (e.g. e-book reader, symbolic items) with them.
However, privacy has to be provided on a physical as well as on a mental level [7]. To create an
environment where they can retreat and be on their own, the astronauts use an audio-visual
stimulation kit. Audio-visual stimulation is a kind of brain entertainment, which adapts
the brain waves with external impulses. Due to the fact that frequencies of brain waves
correlate with specific mind states, a harmonization of brain waves can induce the related
mind state. By using audio-visual stimulation, brain wave frequencies from the highest state
of consciousness state (beta and gamma waves) up to the state of non-rapid eye movement
sleep (delta waves) can be evoked. Therefore, this kit provides a supportive method for
inducing sleep, privacy or a higher mental active state [1, 30].
Despite all stressors, the astronauts have to cope with during the mission, given solutions
should minimize risk for the occurrence of mental disorders. Nevertheless, human factors,
such as physical closeness, communication and family support, also have to be taken into
consideration. As it is not possible to establish a live connection due to signal delay, astronauts
will be given the possibility to keep a blog and record or receive videos from Earth. The goal
is to maintain motivation and significance of the mission [7].
14
5 Spacecraft Design
Obtaining detailed information about the general properties of a spacecraft and its subsystems
from manufacturers proved to be a challenge. Manufacturers’ information was largely used
for the preliminary decision process and for the estimation of available volume. Finally, all
subsystems were designed from scratch for this demanding mission. Although this seems like
a large amount of modifications, special care was taken to keep cost, complexity and need for
extensive research and development at a minimum.
5.1 Configuration & Structure
The design of a spacecraft used for a mission largely depends on the specific mission requirements. In any case, all subsystems have to be integrated into the structure in a way,
that allows it to be launched by available rockets and assembled easily in orbit, if required.
For a manned mission, the spacecraft also serves as a habitat and has to fulfill stringent
requirements. Furthermore, as many subsystems as possible have to be accessible for repairs
without extra-vehicular-activities (EVA). Finally, the overall imperative of a safe, simple and
cheap solution imposes further constraints on the available choices of crew habitats.
During the decision process, key aspects for finding a suitable solution were cost, availability,
pressurized and unpressurized volume and mass. After evaluating multiple concepts, more
advanced concepts like inflatable habitats with artificial gravity systems (i.e. NASA NautilusX) were discarded. Possible advantages of large habitable volume and superior shielding do
not outweigh present problems of development time and unproven reliability when compared
to the traditional aluminum can design. Remaining spacecraft are compared in Table 5.1 to
find a combination of systems that best suit the needs of this mission. Due to availability
and promising development, the Dragon by SpaceX is chosen as the command module
and combined with different modules. To augment the available volume for the different
subsystems and as habitable space, a second element is needed. Options for that include the
European Automated Transfer Vehicle (ATV), Orbital Sciences’ Cygnus and an advanced
inflatable concept by Bigelow Aerospace. The TRL has been estimated according to the
presented mission requirements. The comparison is also based on constraints put upon the
structure by other subsystems and mission aspects.
Table 5.1 Structural trade-off
Dragon
+ ATV
3
+ Cygnus Enh.
3
+ Dragon
3
+ BA330
Habitable volume
Mass (no PL)
Total cost
TRL
58 m
33.8 t
440 M$
6+9
37 m
19.3 t
378 M$
6+7
20 m
26 t
280 M$
6+6
340 m3
34 t
140 M$ + n.a.
6+4
Results
Oversized,
expensive
Fulfills requirements best
Very small
Large volume, little information, lowest TRL
A trade-off shows that the Cygnus represents the best compromise between extra volume,
launch size, mass and cost. Relevant data sheets can be found in the appendix [36].
15
5 Spacecraft Design
5.1 Configuration & Structure
5.1.1 Final Configuration
Both Dragon and Cygnus have to be modified to accommodate the systems required for a
long-term, manned mission. To make optimal use of the trunk normally attached to the
Dragon capsule, it was decided to connect trunk and Cygnus structure via a bulkhead and to
perform an Apollo-style docking maneuver in orbit. A launch and deployed configuration is
shown in Figure 5.1. Furthermore, the trunk should be at least partially pressurized to allow
the crew to move freely between Dragon and Cygnus. It will also support radiators, two of
the solar panels and one of the high gain antennas. This requires extensive modification of
the trunk structure but benefits in volume and functionality are well worth it and allow for
further use of the trunk structure in future deep space missions.
(a) Configuration of spacecraft during launch
(b) Deployed spacecraft after TMI
Figure 5.1 Structural assembly of Dragon capsule, trunk and Cygnus
5.1.2 System-Layout & Structure
The structural mass of the spacecraft parts was
determined through simplified calculations assuming several worst-case load scenarios (reentry, launch). Thus, a minimum wall thickness
and mass for several frequently used spacecraft
materials was determined and the lightest one
chosen (aluminum (7075-T73), aluminum honeycomb). Moreover, the structure subsystem
includes International Standard Payload Racks
(ISPL) for storage of consumables. A 3D-Printer
is utilized for the production of spare parts. Up
16
Table 5.2 Structure system summary
Mass Volume Power
Dragon capsule
Dragon trunk
Cygnus enh.
Docking adapters
3D-Printer
ISPL-Racks
Tools & accessories
Total
968 kg
414 kg
776 kg
494 kg
165 kg
509 kg
276 kg
0.4 m3
0.3 m3
0.6 m3
1.3 m3
144 W
-
3500 kg
2.6 m3
144 W
5 Spacecraft Design
5.2 Subsystems Design
to 0.25 m3 of parts can be produced with 127 kg of raw material with 35% fill. Moreover,
tools and accessories such as test equipment, fixtures and restraints [31] are included in the
budget in Table 5.2.
The distribution of the subsystems across the spacecraft accounts for several critical
aspects: storage of H2 , dinitrogen tetroxide (NTO) and monomethylhydrazine (MMH) in
non-pressurized compartments, spatial distribution of critical systems, fairing restrictions,
low mass of reentry vehicle, little necessity for outside wiring or piping and current lack of a
docking system with an umbilical connection. Furthermore, it acts as a proof of concept that
enough living space remains for the astronauts.
Radiators
CS
T PS T
Rad.
Human
Factors
AOCS
ECLSS
COMM
Radiators
ISPL
TCS EPS
Docking Adapter
CO
MM
TCS
Science
Human
Factors
ECLSS
Rad.
Scienc
e
Docking Adapter
Heat Shield
CS
AO
Solar Arrays
EPS
Solar Arrays
Figure 5.2 Schematic system layout
5.2 Subsystems Design
This section details the design of the subsystems and the assumptions, trade-offs and design
choices are presented. Each section is concluded with an overview table of the relevant
parameters of the elements. Finally, mass, volume and power budgets as well as the systems
engineering approach are given in Section 5.4.
5.2.1 Attitude and Orbit Control System (AOCS)
The attitude and orbit control system controls orientation and orbit of the spacecraft by
disturbance rejection. This is done cooperatively by the sensor and actuator suite as well as
the utilized control algorithms. Identified control modes, disturbance environments and the
relevant actuator suite are presented in Table 5.3.
Sensor and actuator suites are selected and distributed such that all autonomous spacecraft
meet the requirements for their particular control modes. An overview is given in Table 5.4.
While attitude determination is conducted solely with on-board sensors, orbit determination
depends on support from Mission Control Center. Doppler orbitography and radiopositioning
17
5 Spacecraft Design
5.2 Subsystems Design
Table 5.3 Control modes and disturbance environments
Control Mode
Selected Type of AOCS
Orbit insertion
Stabilizing the PM
Transposition & docking
Orbit adjustment MTV
Docking PM & MTV
Stabilizing in Earth orbit
Launch vehicle controls
Gravity-gradient stabilization, thrusters
Thrusters Dragon
Thrusters Cygnus
Momentum wheels, thrusters
Gravity-gradient stabilization, momentum
wheels, resistojets
Momentum wheels, resistojets
Momentum wheels, resistojets
Thrusters Dragon
Deep space flight
Mars orbit control
Orbit & reentry control
Major
Disturbance
Gravity-gradient
Gravity-gradient
Gravity-gradient
Gravity-gradient
Gravity-gradient
Solar, internal
Gravity-gradient
Aerodynamic drag
with DSN and Sun sensors is conducted as well as GPS positioning in Earth orbit. The
attitude of the spacecraft is obtained by a combination of inertial measurement units aligned
by star trackers and GPS.
For assembly in LEO, thrusters and resistojets are utilized, while after the rendezvous
maneuver momentum wheels are added to the actuators. Resistojets operate with backup gas
for low thrust operations. However, after TMI, attitude is solely controlled by momentum
wheels desaturated with waste gas from the ECLSS (0.86 kg of CO2 and 0.43 kg of CH4 ) by
using resistojets. From 4450 N m s up to 11 120 N m s of angular momentum, depending on the
axis, can be desaturated daily. To save electrical energy, resistojets can operate as cold gas
thrusters as well. By using resistojets in combination with momentum wheels, small thrust
impulses are feasible while the dimensions of the momentum wheels enable the spacecraft to
rotate fast enough in case of a SPE. For synchronization in Earth orbit, rendezvous maneuver
and orbit control, 400 N m bipropellant thrusters are provided using MMH and NTO.
On the basis of estimated ∆v needed
Table 5.4 Attitude & orbit control system summary
for synchronization, rendezvous, orPower
bit control and reentry the propellant
Mass Volume
Average
Peak
type is chosen regarding its required
mass. Bipropellant MMH/NTO is
Actuator suite
325 kg 0.27 m3
420 W 2295 W
Sensor suite
43 kg 0.08 m3
76 W
76 W
used due to mass savings compared
3
Docking
system
127
kg
0.28
m
84
W
126
W
with a monopropellant system. For orPropellant tanks 184 kg 1.93 m3
bit synchronization, a trade-off is done
Utilities
25 kg 0.96 m3
54 W
90 W
between synchronization time and rePropellant
1166 kg
0.0 m3
quired fuel mass due to propellant boil3
Total
1871 kg 3.53 m
634 W 2551 W
off of the PM. A sufficient amount of
propellant is provided to complete docking in seven hours. For assembling the spacecraft in
Earth orbit, additional systems and sensors are necessary. GPS provides the information
needed for a rough approach of the MTV towards the PM. Docking maneuvers are conducted
with a combination of a laser-reflector system (RVS-3000) with video support. Redundancy
is ensured by a radio telemetry system (KURS). For the rendezvous maneuver of the MTV
and PM, both units are equipped with the those systems for redundancy reasons.
18
5 Spacecraft Design
5.2 Subsystems Design
Algorithm
αMPC
βMPC
γMPC
αPID
βPID
γPID
0.3
0.2
Control effort [N m]
Angle [rad]
MPC gained popularity as a method for feedback control due to its ability to optimize
closed-loop performance of plants while constraints on inputs, internal states, rates of change
and outputs are taken into account. The availability of optimization solvers, i.e. Quadratic
Programming solvers [29], and sufficient computational performance, dictated by time step
intervals, enables the application of MPC in these fields.
MPC explicitly takes future time steps and the impact of the input on the system into
account and, therefore, it is superior to the backup PID controllers. As a result of the
optimality of the input signal for future time steps, propellant mass and electrical energy can
be saved. This can be achieved with the same sensor and actuator suite; only the computer
systems and algorithms have to be changed. In fact, improved capability of the control
algorithm enables the use of an actuator suite with lower nominal performance, though
without losing performance of the overall control system. The input cost saving for a generic
example is shown in Figure 5.3. The MPC controller needs only 44.8% of control effort
compared with the PID controller without a resulting permanent control deviation.
0.4
0.2
0.1
0
0
5
10
15
20
25
30
Time [s]
MPC
PID
0.15
0.1
0.05
0
0
5
10
15
20
25
30
Time [s]
(a) Euler angles step response
(b) Accumulated control effort
Figure 5.3 Comparison of MPC and PID controller with step response of α = 0.1 rad, β = 0.2 rad,
γ = 0.3 rad and similar reaction time
As for the Curiosity Rover, reliability of control software and reduction of fatal errors
can be achieved by using a development technique presented in [21] which is based on three
principles: implementation of risk-based coding rules, using tool-based code review and a
logic model-checking tool to formally verify mission-critical code segments. By making the
control code accessible from Earth, it can be debugged during the flight which results in
higher reliability. A more detailed approach is described in the appendix [36].
5.2.2 Communication System
For long duration missions, communication is vital for the psychological well-being of the
astronauts. This leads to challenging requirements concerning coverage, reliability and
availability of the communication system. The onboard system consists of two redundant
high-gain parabolic reflectors that can be aimed individually and two low-gain backup
antennas. They are fed by four redundant transponder systems, with 70 W output power
19
5 Spacecraft Design
5.2 Subsystems Design
each. If required, power can be saved by deactivating channels separately. An overview over
the key characteristics of the system is shown in Table 5.5 based on link budget calculations
[22, 32]. A robust and delay-tolerant transmission technique with forward error correction
should also be considered [32, 36].
Table 5.5 Characteristics of X-band nominal operation with high-gain antenna at 1.43 AU [36]
Transmit
Center
Transmit
Max.
System Bit Error
Link Antenna Frequency
Space Loss Data Rate Margin
Power
Rate
Up
Down
ø 15 m
ø2m
7.200 GHz
8.425 GHz
2 kW
4 × 70 kW
−272.7 dB
−274.1 dB
54 kbps
4 × 18 kbps
3.5 dB
3.5 dB
10−5
10−5
Due to limitations of the ground stations (see below) X- and S-band are chosen for transmission. Frequencies should be allocated early in the development phase as the classification
as a Space Research Vehicle is to be decided by ITU and critical for subsequent development
[23]. As the use of large antennas is very expensive and limited, facilities with a maximum
antenna size of 15 m are chosen for nominal operation and baseline ground segment, providing
the required permanent communication. This includes standard communication with mission
control as well as high priority emergency communication. 15 m antennas are also much more
common than the higher gain 35 m and 70 m classes, which is favorable regarding failure safety.
Most of the reviewed systems only support X- and S-band and rarely K-band communication
[43]. Even though X- and S-band provide less bandwidth, they are less susceptible to rain
attenuation [32]. Depending on space agency participation, the higher availability of 15 m
antennas has most likely the biggest impact on ground segment choices.
Apart from nominal telemetry,
Table 5.6 Major required data transmissions
tracking & control (TT&C) comData & Data Rate Timing
munication, there are two other ma- Subject
jor data sources. A summary of the TT&C
10 kpbs
Continuous
most important data to be trans- Video messages
100 MB at avail. rate Every second day
Three times a day
mitted can be found in Table 5.6. Medical monitoring 6 MB at avail. rate
Every other day, video messages
are exchanged. For a video of about 20 minutes length and 360p resolution this takes about
14 minutes at 1 Mbps. Additionally, three times a day the medical data is transmitted to
Earth for analysis (see Section 4.2).
Figure 5.4 depicts the theoretically available downlink data rates with mostly all channels
used. They are modeled using the transformed link budget equation [36]. At the beginning
and the end of the mission the data rate is limited at sufficient 1 Mbps to visualize the
possible power saving channel adjustments. After the fly-by at Mars, the data rate drops
to a minimum. Over a time of roughly 150 days the data can only be transmitted with less
than 200 kbps to 15 m ground stations. This can be mitigated by increasing the transmission
time, reducing the video quality, or using ground stations with larger antennas for high data
rate transmissions as exist in the DSN or ESTRACK for example.
In Figure 5.5 the different mission phases are depicted from the communication point of view.
At the beginning (phase 1) low delay communication is possible for a few days. In the first
part of the mission the delay is between ten seconds and about three minutes (phase 2). One
critical point prior to Mars is reached when Sun, Earth and spacecraft are in conjunction (point
3). The uplink will contain increased thermal noise, due to the strong radiance of the Sun.
20
5 Spacecraft Design
5.2 Subsystems Design
1
0.8
0.8
0.6
0.6
0.4
0.2
0
Distance
0
50
100
150
0.4
Data rate 15 m
Data rate 34 m
Data rate 70 m
200
250
300
350
400
Distance [AU]
Data rate [MBit/s]
1
0.2
450
0
500
Time [d]
Figure 5.4 Theoretically available downlink data rates with different ground station antenna ø
Earth
After that, the most important part of the trajectory
is reached, the Mars flyby (point 4). There will be
Spacecraft
a high demand for communication, but also some
Trajectory
obstacles. However, the time delay at that point will
be at about 3.5 minutes, and there will be a blackout
Sun
while Mars is between the spacecraft and Earth.
By using relay satellites of the Mars Exploration
Joint Initiative communication during flyby could
be provided. In this case, it would also be possible to
transmit higher quality video data to Earth. For this
purpose an Electra UHF system is utilized aboard
the spacecraft [54]. Another option is to record the
Relay Satellite
mission data and send it to Earth retrospectively.
Mars (at flyby)
Transmission path
After the flyby, the distance between the spacecraft
Figure
5.5
Communication
phases
and Earth will rise to a maximum. At the farthest
point (point 5) the transmission delay is at about
eight minutes and the spacecraft is about 1.43 AU away from Earth. The available downlink
data rate decreases to about 72 kbps for 15 m antennas. From this point on the conditions
improve until low delay communication is possible again before reentry. An element summary
is given in Table 5.7. Further details on the communication system, architecture and
calculations can be found in the appendix [36].
Table 5.7 Communication system summary
Mass Volume
Antenna subsystem
Electra UHF subsystem
TT&C subsystem
Payload & harness
Total
35 kg
12 kg
7 kg
96 kg
2.0 m3
0.01 m3
0.1 m3
0.3 m3
149 kg
2.4 m3
21
Power
Average
Peak
3W
19 W
11 W
122 W
17 W
74 W
148 W
798 W
154 W 1036 W
5 Spacecraft Design
5.2 Subsystems Design
5.2.3 Electrical Power System (EPS)
The electrical power system is responsible for power generation, storage and distribution.
The critical case is identified as the flyby at Mars due to the lowest solar flux and moderate
degradation of solar cells. Solar arrays cover average power and charge batteries, which
cover daily peaks as well as the Mars flyby and reentry phase. The average power of the
different subsystems is summed up while the peak power of certain elements (like oven, waste
compactor and high gain antenna) is distributed through out the day as can be seen in Figure
5.6 and is detailed in the appendix [36].
Power [W]
6000
Capacity [W h]
Average
Peak
4000
2000
0
0
2
4
6
8
10
12
14
16
18
20
22
24
Time [h]
Figure 5.6 Average and peak power distribution
Solar Arrays
During sizing of the solar arrays, several factors such as
Table 5.8 Solar array/cell
material constraints, array losses and environmental losses
losses [31]
have been considered [31]. The assumptions are summaArray resistance
0.958
rized in Table 5.8. The solar constant is assumed to be
Packing
fraction
0.85
inversely exponentially proportional to the distance to the
Tracking loss
0.996
Sun. Additionally, a 10% contingency is added to account
Radiation damage
0.976
for unforeseen eventualities.
UV darkening
0.997
As a result of the analysis, four arrays with GalliumMicrometeroid damage 0.994
Contamination
0.98
Arsenide triple junction cells (with an efficiency of 26%) are
Resistance
losses
0.98
envisioned. Since the solar arrays are sized for operation
Distribution losses
0.917
at Mars, deployment of two arrays is sufficient to cover the
Total Loss
0.69
power requirements in Earth orbit. The required diameter
of the disk-shaped arrays for the current mission is 5 m
(downscaled from 6 m with 7 kW). The specific mass of the arrays is estimated with 175 W/kg
[52]. ATK’s UltraFlex arrays are chosen due to their lightweight structure combined with
high strength and stiffness. They were successfully utilized on the Mars Phoenix Lander in
2008. The arrays are scalable up to 10 kW and a model with 5.5 m has been successfully
tested [52]. Moreover, development of the arrays is ongoing as part of Orbital’s Cygnus
module and NASA’s MegaFlex program.
22
5 Spacecraft Design
5.2 Subsystems Design
Rechargeable Batteries
To cover daily peaks of high-power eleTable 5.9 Battery assumptions [31]
ments and the reentry maneuver, convenNiH2 Li-ion NaS
tional batteries as well as fuel cells have
been considered. Regenerative fuel cells
Specific energy [Wh/kg]
60
130
132
3
are disregarded due to the relatively low
Specific density [kWh/m ]
40
160
165
Depth of discharge [%]
80
80
80
TRL [8]. The ECLSS offered the possibility
Efficiency
[%]
96
93
85
to convert water to oxygen and hydrogen
through electrolysis, thus making the use
of a conventional fuel cell possible. Although, the integration into ECLSS results in an
intolerable mass penalty. Different types of conventional batteries like NiH2 , Li-ion and NaS
are compared with properties presented in Table 5.9. Consequently, Li-ion batteries [3] are
chosen due to their high specific energy and current application on on-orbit systems as well
as the prospective use on the ISS. Depth of discharge (DoD) and efficiency are assumed to be
80% and 93%, respectively. Life time with almost constant capacity is over 1000 cycles at
this DoD [3]. Due to the use of a separate rechargeable battery for reentry with a similar
required capacity, redundancy is given.
Power management and distribution is considered with a constant factor (150 W/kg)
depending on the peak power distributed by EPS. Mass, deployed area, (packed) volume and
technology readiness level are summarized in Table 5.10.
Table 5.10 Electrical power system summary
Mass Area/Volume TRL
UltraFlex solar arrays
128 kg
Li-ion batteries (cruise)
66 kg
Li-ion batteries (reentry)
60 kg
Power management & distribution 52 kg
Total
306 kg
37.6 m2 /0.7 m3
0.05 m3
0.04 m3
-
6-7
8
8
9
0.79 m3
-
5.2.4 Environmental Control and Life Support System (ECLSS)
ECLSS supplies the crew with necessities such as atmosphere, food and water, but also takes
hygiene, clothes and waste management into consideration. In this matter, the crew safety is
directly dependent on the reliability of ECLSS, which does not allow for any single point
failure of the system. A flowchart is provided in Figure 5.7 for an overview of ECLSS and
the synergies with other systems.
Assumptions for ECLSS Design
The ECLSS design is based on an atmospheric composition of 79% N2 and 21% O2 at a total
pressure of 101 325 Pa and a temperature of 295 K. Additionally, atmospheric humidity is
controlled by the atmosphere control system (ACS). Atmosphere for a pressurized volume of
52 m3 is ensured and leakage is considered.
Daily input and output for ECLSS is summarized in Table 5.11 and described in detail
(kg/d/p = kilogram per day per person). For the entire mission duration, dehydrated
23
5 Spacecraft Design
5.2 Subsystems Design
wastewater
SABATIER
N2
H2
CH4
CO2
kg
0K86 day
H2 O
H2
CH4
O2
SPWE
O2
H2 O
WASTE
H2 O
kg
0K43 day
H2 O
AOCS
FOOD
VPCAR
brine
O2
waste
TANKS
CO2
O2
BRINE
FECES
brine
WASTE
WATER
TANKS
H2 O
H2
waste water
N2
water
WASTECOMPACTOR
panels
food
CO2
RADIATION
PROTECTION
CO2
4BMS
leakage compensation
O2 K H2 OK N2
ACS
urine
Q̇ K water vapour
leakage
O2 K H2 OK N2 K CO2
feces
CREW
CO2
atmosphere
Q̇ K water vapour
TC
CHX
TCCS
waste water
Q̇
TCS
Figure 5.7 Flowchart of Mars18 ECLSS (including synergies with AOCS in green and radiation
shielding in red)
food is stored in the spacecraft. A regenerative life support system with implementation
of a photobioreactor for algae cultivation was not considered due to the low TRL. Food
is rehydrated and heated up by the crew. Food lockers, as used on the ISS, are replaced
by polyethylene (PE) lockers for mass reduction and radiation shielding. For the mission
duration of 500 days, mass and volume of the provision adds up to 1407 kg and 4.85 m3
(including margin). The employed water bags resemble ISS bags but contain additional
electrolytes and minerals essential for the crew.
Since volume and energy is limited, showers are not
Table 5.11 Mass assumptions in
possible. Therefore, crew members use different types
kg/d/p [19, 18]
of wipes (including wet, dry, detergent and disinfectant
Food supply (solids)
0.064
wipes) for hygienic purposes. Crew members shave
Food
supply
(water)
0.248
their hair regularly and wear wigs, if desired. Edible
Food supply (packaging) 0.270
tooth paste is used for dental hygiene. A waterless
Food supply (PE-lockers) 0.119
toilet, similar to that on the ISS is installed (toilet
Req. water (rehydration) 1.225
paper included with the wipes). Clothes are supplied
Req. water (total)
7.819
Wipes
0.206
and dispensed in the waste compactor after usage since
Clothes
0.343
no laundry is installed due to low TRL. Cotton clothes
Feces & urine
2.096
are preferred to an advanced clothing concept for higher
General trash
1.060
comfort. General trash as well as urine and feces
require processing/storage. Additional 62 kg of miscellaneous supplies (including duct tape,
latex gloves etc.) are needed during the mission [18].
24
5 Spacecraft Design
5.2 Subsystems Design
ECLSS Concept and Design
Equivalent system mass [kg]
For the life support system, different con10000
cepts are evaluated and compared. First
Open System
priority is reliability and readiness of the
Closed System (MF+CVD)
system by launch in 2018. This reduces
8000
Closed System (VPCAR)
the technology options greatly. Increasing
safety and reducing equivalent system mass
6000
(ESM) must be achieved with flight-proven
technologies. For the ESM calculations,
mass, volume, power in- and output as well
4000
as crew time are considered [33]. For an
optimal trade-off, synergies with other subsystems are formed. The two most promis2000
100
200
300
400
500
ing concepts only differ in the waste water
Time [d]
(WW) treatment. A concept corresponding
to the ISS water treatment with Multifil- Figure 5.8 Trade-off between different concepts
tration [61] (MF) and Vapor Compression
Distillation [61] (VCD) as additional urine treatment is compared to one using Vapor Phase
Catalytic Ammonia Removal [61] (VPCAR). Both concepts are compared to an open loop
concept in Figure 5.8. The VPCAR concept is chosen due to a mass reduction of 944 kg.
VPCAR is able to convert different types of waste water into potable water without the
need of pre- or post-treatment. The system works discontinuously 101 minutes per day and it
can purify up to 134.4 kg/d [61]. In this concept, a time-averaged processing rate of 9.406 kg/d
waste water is required. Downsizing would be the better option for minimizing ESM. Since it
has to be developed by June 2016 to TRL 8 to ensure sufficient testing, this risk is not taken
and an unscaled VPCAR system is used. Four-Bed Molecular Sieves [15, 39] (4BMS), Solid
Polymer Water Electrolysis [17] (SPWE), Condensing Heat Exchanger [18] (CHX) and Trace
Contaminant Control System [18] (TCCS) are chosen for the air management as these are of
TRL 9 and thus ensure a high level of reliability and safety. Other technology options, such
as Electrochemical Depolarized Carbon Dioxide Concentrator [39] (EDC) and Solid Amine
Water Desorption [15, 39] (SAWD) are excluded due to low TRL. Furthermore, SAWD has
a potential hazard of producing toxic vapor by amine degradation. A Sabatier reactor is
employed for chemical reactions of CO2 with H2 to H2 O and CH4 . From recovered H2 O, O2
can be reused again. Monitoring, atmosphere control and leakage compensation is guaranteed
by ACS [18] as already employed on the ISS. Many state of the art life support technologies
are not designed for a crew of two. Hence, 4BMS and Sabatier are scaled to optimize ESM,
with 20% of the mass fixed and thus not scaled. The remaining part is scaled linearly.
During manned missions, following waste components are produced [18]: General trash
(e.g. packaging material, clothes, hygiene wipes, food waste), feces with toilet paper, urine,
waste water and brine. Urine and waste water are recycled by VPCAR. Feces are stored. A
waste compactor (PMWC [24]) is used to compress trash and brine for volume reduction and
water extraction. Compressed trash tiles are used for radiation shielding and extracted water
is stored together with feces to support radiation shielding. Recovered water is not reused
since its quality cannot be guaranteed.
25
5 Spacecraft Design
5.2 Subsystems Design
Table 5.12 Mass flux balance. Positive values mark system output, negative values system input.
Values are in kg/d, rounded to two decimal figures. H2 O Rad is water passed to radiation shielding
tank.
CO2
Stored Air
Food
Crew
4BMS
SFWE
Sabatier
CHX
PWMC
VPCAR
Leakage
- 2.00
2.04 -2.04
-1.18
- 0.04
-
Day [kg/d]
0.86
Mission [kg] 430.99
N2
H2
CH4
- -1.67
- 1.72 0.22
- -0.22
0.01 -0.05
-0.03 -0.01
-
0.43
-
0 -0.02
0 -8.52
O2
0
0
Cycle
0.50
-7.82
-1.94
9.21
-
H2 O
Rad
Air
- 4.55
- -4.55
0.40
-
0
0.43 -0.05
0.4
0 214.18 -23.17 199.53
0
0
WW
0.96
4.55
-5.52
0
0
Average mass fluxes per day between elements are shown in Table 5.12. The system is
designed to balance out the usage and production of H2 and O2 . Excess CH4 and CO2 are
used by the AOCS and not actually stored. Leakage is assumed to be 0.625 g fluid/m3 /d,
based on empirical data from the ISS [9].
Storage and Safety Issues
ECLSS is designed to provide water and oxyTable 5.12 Mass flux balance [cont.]
gen needed during mission after Table 5.12
Urine Brine Waste Feces
and to compensate failure of any technology
for two weeks. Storage and initial fluid supply
Food
0.42
is sufficient to compensate the supplies from
Crew
3.89
1.46
0.30
PWMC
- -0.19
-1.88
any system as well as being able to store waste
VPCAR
-3.89
0.19
water, hydrogen etc. for two weeks in case of
Day [kg/d]
0
0
0
0.3
emergency. Enough lithium hydroxide (LiOH)
Mission
[kg]
0
0
0
151
cans are provided for emergency carbon dioxide filtering for two weeks [15]. In that time,
the crew has to repair the affected systems. Additional N2 and O2 are provided to replace the
atmosphere in the case of a fire. Gas is stored in high-pressure tanks with the mass estimated
per stored fluid mass (based on reference technologies and Barlow’s formula [9]). CO2 and
CH4 storage is designed to store produced gases for one day each, before the gases have to be
vented by AOCS. Waste water and urine is stored together. Brine is processed in the waste
compactor and stored for one day. Water output from the waste compactor together with
feces is stored in PE tanks, which also act as radiation shielding. Fresh water is also stored
in a separate PE tank. Storage volume and initial supply is shown in Table 5.13.
Throughout the spacecraft, IR-cameras are installed for automatic fire detection and
localization. In case of a fire, the crew wears masks connected to oxygen supply and
extinguishes the fire with portable fire extinguishers. The pressurized volume can be divided
in two parts at the connection between the trunk and the Cygnus module through a bulkhead.
26
5 Spacecraft Design
5.2 Subsystems Design
Table 5.13 Storage capacity and initial supply
O2
N2
H2 CO2 CH4 H2 O
Capacity [kg] 24.11 33.2 3.01
Supply [kg]
24.11 33.2
0
2.04
0
0.43
0
137
137
WW,
Feces,
Brine
Urine
Water
186
0
2.63
0
355
0
If the fire cannot be controlled, the crew has to evacuate the affected part and the atmosphere
of the damaged section is removed to extinguish the fire. Afterwards, the atmosphere is
renewed with the designated gas supply. Therefore, 24.7 kg N2 and 7.5 kg O2 is initially on
board to replace the Cygnus atmosphere, which is the largest possible volume to be isolated.
Additional CO2 filtering is required to ensure crew safety during reentry. For this purpose
a LiOH-filter is employed through which crew expiration is led. About 2 kg of LiOH are
needed to reduce the 1.17 kg CO2 exhaled in 14 hours (worst-case) [15]. A mass of 5 kg is
estimated for the filtering system.
Simulation with ELISSA
The concept presented above is simulated time resolved with the tool ELISSA (Environment
for Life-Support Systems Simulation and Analysis) provided by the Institute of Space Systems
at the University of Stuttgart. Simulation results prove adequate supply and storage and are
shown in the appendix [36].
Table 5.14 Environmental control & life support system summary (technologies marked with *
are scaled, shown power consumption is not continuous)
Technology Mass [kg] Volume [m3 ] Power [W] TRL
CO2 Filtration
O2 Production (electrolysis)
Thermal control and humidity
Trace contaminant control system
CO2 Reduction (get H2 O and CH4 )
Waste water treatment
Atmosphere control system
PMWC waste compactor
Waste collection system + supplies
Gas storage and tanks (incl. brine)
4BMS*
SPWE
sabatier*
CHX
TCCS
VPCAR
ACS
PMWC
69
292
23
144
18
494
169
71
122
156
0.26
1.79
0.04
0.60
0.07
1.88
0.05
0.48
2.85
0.43
381
351
0
263
13
2618
32
630
0
0
9
9
8
9
9
6
9
6
9
9
Water suppy
Food supply and packaging
Conduction oven and rehydration
Clothes
Personal hygiene
144
1407
38
360
252
0.14
4.85
0.10
2.10
0.37
0
0
1008
0
0
9
Lithiumhydroxyd cans
Fire detection system from ISS
Reentry CO2 filtering
51
10
5.25
0.26
0.04
0.01
0
0
0
9
9
9
Total
3825
16.3
5294
27
5 Spacecraft Design
5.2 Subsystems Design
5.2.5 Propulsion System
The PM is responsible to inject the MTV into the determined free-return trajectory on
January 4, 2018. As already mentioned in Chapter 3.2, the PM is closely related to the
launch concept. This chapter, however, will focus on the PM only.
Design Process & Requirements
The PM needs to be capable of injecting roughly 15 000 kg into the Mars transfer trajectory.
Therefore, a total ∆v of 4841 m/s is required. To account for losses due to thrust inaccuracies
and the Oberth Effect, a margin of 5% is added to the ∆v. However, the g-load must not
exceed the maximum load rated for humans. All upper stages and their respective engines as
well as solid and electrical engines were evaluated in-depth to get a comprehensive overview.
In order to keep the propellant mass of the PM as low as possible, engines with a high specific
impulse (Isp ) are compulsory.
Thus, solid engines were disregarded and the analTable 5.15 Engine trade-off
ysis shows that only a handful of engines actually
Engine
Isp [m/s] Upper stage
support an adequate Isp for such a mission, as shown
in Table 5.15. The respective data sheets can be
RL-10 B2
4566
Centaur- & Delta
found in the appendix [36]. The arcjet engine TIHIV Second-Stage
Vinci
4561
ESC-B
TUS developed by the Institute of Space Systems in
J2-X
4395
Earth
Departure
Stuttgart seems promising due to its extremely high
Stage
Isp and thrust for an electric engine of 100 N but
TIHTUS
10 000 is disregarded because of its high electrical energy
consumption of 500 kW. The J2-X and Vinci as well as their respective upper stages on the
other hand will not be sufficiently tested until 2018. In contrast, the RL-10 B2 proved to
be a very reliable engine in a multitude of missions. It is one of the most used upper stage
engines and provides the highest Isp due to its extendable nozzle.
Additionally, many unconventional concepts have been considered like using cross-feeding
between different propulsion stages, using alternative propellants like lithium and fluorine
to get a higher Isp or launching empty rockets to use the remaining upper stage as the PM.
However, only two concepts prove to be satisfying. The chosen concept is described in the
next chapter and the alternative is explained in the appendix [36].
Propulsion Module Concept
The final PM consists of two Delta IV 4-m Second-Stages with one RL-10 B2 engine per
stage [5]. It uses liquid oxygen (LOX) and liquid hydrogen (LH2 ) for propellant and offers a
5% ∆v margin. In order for the PM to fit into the Asymmetric Payload Fairing (APLF),
the two stages are connected over four hinged telescope beams which hold them in their
determined position. In launch configuration, the stages are positioned side by side as shown
in Figure 5.9. After the PM reached parking orbit, the hinged telescope beams are extended
electrically. Afterwards, the second stage moves backwards until final TMI-configuration
is attained. It will then be safely locked in its position. Due to this construction concept,
another docking maneuver is not necessary and the mission complexity as well as the fuel
mass for the AOCS can be reduced enormously.
28
5 Spacecraft Design
5.2 Subsystems Design
Figure 5.9 Configuration change of PM after launch
After the configuration change, the PM turns with an additional mounted magnetorquer
into a gravity-gradient stabilized position to reduce the fuel consumption of AOCS until it
docks with the MTV. For this docking maneuver, the second stage of the PM is equipped with
a modified payload adapter with a reinforced Soyuz probe and drogue docking port, which
will connect to the docking port of the Cygnus. The PM, additionally, has an enhanced AOCS
with a reflector system, a GPS module and a KURS docking system to be able to perform
this rendezvous. Due to the fact that in the worst-case the PM has to stay in the parking
orbit for twelve days without the electrical supply of the MTV, it is also equipped with body
mounted solar cells that provide the electric energy for the magnetorquer and the avionic.
Moreover, as the PM will be in the orbit 14 days before
Table 5.16 Propulsion module
the TMI, the boil-off of the cryogenic propellant of the
1. Stage 2. Stage
PM must also be carefully examined. The propellant
tanks are equipped with 50 layers of double aluminized Structure
2850 kg
2850 kg
Kapton insulation (MLI) which decreases the boil-off to Propellant
20 410 kg 20 410 kg
a tolerable level of 1.3% of the total propellant mass for Add. structure
142 kg
580 kg
these 14 days [12]. All these modifications are shown in & EPS
AOCS
35 kg
152 kg
Table 5.16 and the calculation of the TMI with the above MLI insulation
334 kg
334 kg
mentioned worst-case boil-off of 272 kg of each stage is
Total at TMI 23 771 kg 24 326 kg
shown in Table 5.17. For more detailed information see
the appendix [36].
Table 5.17 Propulsion module before TMI (including losses due to boil-off)
∆v
Isp
1. Stage 1774 m/s 4566 m/s
2. Stage 3310 m/s 4566 m/s
Total ∆v 5084 m/s
Dry Mass
Propellant Stage Mass
3361 kg
3916 kg
20 138 kg
20 138 kg
Total Wet Mass
47 553 kg
61 349 kg
37 850 kg
5.2.6 Radiation Protection
Space radiation mainly consists of particle radiation, which is very dangerous for humans
if certain dose equivalent rates are exceeded. If primary particle radiation impacts matter,
secondary radiation is produced, which can be even worse for humans. There are different
sources of radiation in space. Trapped radiation, as it can be found in the Van Allen radiation
belts, can be neglected. Also electromagnetic radiation does not cause problems for manned
missions. Galactic cosmic rays (GCR) are highly energetic and consist mainly of protons
with energies up to a few TeV [59] at low flux densities. GCR are especially critical for
manned long-term missions. Because of high energies, GCR cannot be shielded effectively
29
5 Spacecraft Design
5.2 Subsystems Design
with available technology. Another radiation danger is caused by solar activity such as flares
and coronal mass ejections. SPEs increase the particle flux, mainly protons, significantly at
an energy of a few hundred MeV [59]. This kind of radiation must be shielded since it can
produce lethal radiation doses in a few hours. Therefore, storm shelters are used. Frequency
and strength of SPEs depend on the solar cycle. Since the mission is planned to be during a
declining Sun cycle near a solar minimum, only a few SPEs are expected [60]. The strength
of SPEs is evaluated based on its spectrum of energy and flux. A strong event can neither be
predicted nor guarded against at the present time, but these are rare [50, 51]. For example,
during solar cycle 21, no unusually large flare occurred [38]. Nevertheless, the assumption
that SPEs do not take place during a solar minimum cannot be made.
Alternatives to mass shielding are for example electrostatic shielding, magnetic shielding
or biomedical solutions. So far, only concepts and proposals have been found whose power
demand and system weight are very high and/or whose TRL are too low to be ready in 2018
without making large investments [45]. For these reasons, the proposed solution for GCR
and SPE radiation shielding is the usage of synergistic effects with other systems, smart
positioning of actual load, adjustment and replacement of some subsystem material and
personal shielding by vests and blankets.
Sensor Systems and Early Warning for SPEs
In order to collect data for future missions, as it was done by the Radiation Assessment
Detector (RAD) on the Mars Science Laboratory [62], and to record the received doses for
aftercare reasons, astronauts are equipped with personal active dosimeters and two radiation
sensors are mounted on the spacecraft. This poses a scientific experiment since body dose
values are received during a long-term mission which could not be done before.
Since passive dosimeters would become saturated during long-term missions, a new generation of active dosimeters, being developed by the German Aerospace Center (DLR) in
Cologne, with the goal to fly at the end of 2015 is employed. Furthermore, two Timepix
sensors developed at CERN and used on the ISS are mounted on the spacecraft [46]. One is
placed on the inside of the Cygnus, which is the preferred habitat for the astronauts. The
other one is placed on the outside facing the Sun to detect SPEs. In case of an SPE detected
by an increase of X-Rays, the crew is warned and has to move to their storm shelter. For
redundancy reasons, the spacecraft also communicates with the satellites GOES and SDO
via ground stations, which are capable of announcing SPEs as well.
Proposed Strategies for Dose Equivalent Rate Reduction
Dose equivalent rates caused by GCR cannot be shielded Table 5.18 Overview of different
effectively without a large increase of system mass. An assembly zones
increase of the shielding from 10 g/cm2 to 40 g/cm2 aluArea Weight
minum reduces the dose equivalent rates for near Earth
2
Heat shield
9.8 g/cm
interplanetary space only by 2% during a solar minimum
2
Dragon
12.9 g/cm
2
[42]. Therefore, the system is designed to provide adequate
Dragon trunk
9 g/cm
2
shielding for regular SPEs. To minimize system weight,
Cygnus trunk
31.9 g/cm
only Cygnus is equipped with additional shielding, reduc-
30
5 Spacecraft Design
5.2 Subsystems Design
ing Dragons reentry mass. A mean area weight for the different areas in the assembly is
determined, as shown in Figure 5.10 and Table 5.18. Material for the first layer of radiation
shielding is assumed to be aluminum.
Figure 5.10 Definition of different shielding zones
20
Structure & systems
Water from food
Food & packaging
Feces, tiles/water from waste
Total
15
10
5
0
0
100
200
300
400
h
i
Area weight g/cm2
h
i
Area weight g/cm2
Since Cygnus becomes the main habitat that is only left for privacy time spent in Dragon
(1.5 h/d) and repair work in the Dragon trunk (1 h/d), additional arrangements are made.
Cygnus is equipped with two tanks made of PE, which is superior to aluminum for shielding
purposes due to its high hydrogen content, lower density and less production of secondary
radiation [59, 60, 6]. Each tank covers half of the Cygnus’ outer area. The tank on Cygnus
side 1 is empty at the begin of lifetime (BOL) and is filled with feces during the mission
and the water recovered by the waste compactor [24]. The tank on Cygnus side 2 contains
process water from ECLSS at BOL. In addition, food, wet wipes and disinfectant wipes are
arranged on the inside of the capsule to produce a protective curtain whose dose reduction
effect was proven in [27, 47]. The waste compactor generates a plastic tile of 0.41 m × 0.41 m
per day, which can be used to improve radiation shielding. The development of the shielding
area weights are shown in Figure 5.11.
20
25
500
Time [d]
15
Structure & systems
Processed water
Wet & disinfectant wipes
Tiles from waste
Total
10
5
0
0
100
200
300
400
500
Time [d]
(a) Side 1
(b) Side 2
Figure 5.11 Area weight over time of Cygnus module surface area
Since SPEs can be predicted neither in occurrence nor in direction, additional measures
must be taken, based on the development of the shielding strength. In Figure 5.11 the
critical phases can be seen: for Cygnus side 1 it is the end of life (EOL) since the total
31
5 Spacecraft Design
5.2 Subsystems Design
area weight decreases over time and for Cygnus side 2 it is BOL (increasing area weight
over time). The thickness of the tank walls, made of polyethylene, is chosen to match the
area weight of the TransHab Radiation Shield Water Tank of 5.74 g/cm2 [2]. This results
in a wall thickness of 1 mm for Cygnus side 1 and 10 mm for Cygnus side 2. In order to
achieve a high shielding/weight ratio, it is effective to implement any shielding as close as
possible to the human body. This personal shielding will consist of garments and sleeping
covers made from PE, which has proven to be highly effective in absorbing particle radiation.
Breathable, sleeveless vests will be made in a shielding quality of 1 g/cm2 . This will protect
blood-forming organs while not being uncomfortably hot for the astronauts. They should
be worn as often as possible, especially while staying in the lesser shielded Dragon capsule.
Another measure to increase shielding will be sleeping bags made from PE, again with an
area weight of 1 g/cm2 . Because of the lack of natural convection, cylindrical bags with a
large diameter of 0.5 m will prevent overheating. In both cases the fabric will be made from
extruded PE fibers in different weights/diameters depending on application. Suggested fiber
weights are 500 dtex for the shell of the sleeping bags and 100 dtex for the fill material as well
as the woven fabric for the vests. While these measures do not apply throughout the whole
day, it still decreases the statistical radiation dose and should be seen as a bonus on top of
the aforementioned spacecraft shielding.
When an approaching SPE is detected, the astronauts put on the vests and wrap in their
sleeping bags. Until the sensors register the end of the SPE, the astronauts have to stay in
the radiation shadow zone between the blue ring in Figure 5.10 and the Cygnus wall. The
ring provides radiation protection from less shielded areas than Cygnus (paths through the
heat shield, Dragon trunk or Dragon capsule). Due to its chemo-protective effect, amifostine
is used for radioprotection at a radiation exposure to achieve a reduction of the hematocrit
toxicity. This drug was developed for radiation therapy and may also be helpful to prevent
physical damage caused by SPEs. A one time application of 740 mg/m2 body surface seems
reasonable [10].
Verification of Shield Design via Simulation
To get the equivalent dose rates for the asTable 5.19 Overview of shielded zones and received
tronauts and to prove the radiation shielddose rates by GCR (mission total)
ing concept, simulations with SPENVIS
Ionizing dose Dose equivalent
and the Geant4 tool MULASSIS have been
performed. MULASSIS calculates the re- Cygnus side 1
0.133 Gy
0.532 Sv
Cygnus side 2
0.135 Gy
0.540 Sv
sulting ionizing doses of a multi-layered
0.133 Gy
0.533 Sv
shield design. Table 5.19 shows the equiv- Cygnus trunk
Dragon trunk
0.154 Gy
0.617 Sv
alent dose rate for blood-forming organs
Dragon
0.165 Gy
0.661 Sv
behind planar slabs of different composi- Heat shield
0.169 Gy
0.677 Sv
tions with a mean quality factor of four
as it has been found for a Mars transit for 500 days [62]. The GCR environment is calculated
for the mission time frame and doses for Cygnus side 1 and 2 are a mean value for BOL and
EOL. Weighted by duration of stay and percentage of slabs on the total surface, this gives a
total equivalent dose of 0.562 Sv (detailed calculation can be found in the appendix [36]).
32
5 Spacecraft Design
5.2 Subsystems Design
Maximum dose for a single SPE plus GCR in Table 5.20 Overview of shielded zones and
30 days is 0.25 Sv [60]. Since SPEs can produce received dose rates by a single SPE
a high equivalent dose in a very short time, it
Dose equivalent
has to be guaranteed that normal SPEs do not
0.0254 Sv
exceed this dose. A measured SPE from October Cygnus side 1 EOL
Cygnus
side
2
BOL
0.0303 Sv
24, 1989 (near solar maximum) is simulated with
Cygnus trunk
0.0166 Sv
the given shielding. Its spectrum can be found Dragon trunk & shelter
0.0328 Sv
in the appendix [36]. The doses presented in
Table 5.20 are for a SPE duration of one day, the used quality factor is one [62] and values
are given for the most critical paths.
The design of a mass-based radiation shielding concept for a 500-day, manned Mars mission
is exemplified and its effectiveness, at minimal additional mass, is proven by simulations.
The dose equivalent rate by GCR for the whole mission is found to be 0.562 Sv and the doses
by regular SPEs clearly do not exceed the 30-day equivalent dose rate of 0.25 Sv. A summary
of the elements is given in Table 5.21.
Table 5.21 Radiation protection summary
Mass
Volume TRL
Radiation sensor (Timepix/Medipix) 0.1 kg 0.044 m3
Active dosimeters (DLR Köln)
0.6 kg 0.0002 m3
2
PE-blanket & vests (with 1 g/cm )
138 kg 0.174 m3
Feces/PMC/processed-water tank
352 kg
0.37 m3
Shelter-ring
240 kg
0.26 m3
9
5
-
0.9 m3
-
Total
731 kg
5.2.7 Thermal Control System (TCS)
The Thermal Control System (TCS) is responsible for maintaining the cabin and all electronic
devices at an appropriate temperature. The sizing is achieved with a simplified energy balance
formula [40]. Nevertheless, all heat sources are considered, such as solar radiation Qs , albedo
radiation Qa , IR radiation Qir , and dissipation heat Qdiss :
Qs + Qa + Qir + Qdiss − Qout = 0.
(5.1)
The emission term Qout is defined by the thermal control system. It is designed to reject the
critical heat load amount of 11 430 W encountered in Earth orbit. Other considered cases
include the crossing of Venus’ orbit and the dark side of Mars.
The TCS consists of two main loops and is depicted in Figure 5.12. The external thermal
control system (ETCS) transports heat load from the heat exchanger to the radiators by the
means of a water-glycol mixture. The second loop is the internal thermal control system
(ITCS), which has a fully-independent fluid loop and uses water as working fluid. Since no EVA
is planned, reserve pumps are installed inside the cabin in case of an emergency. Consequently,
a non-toxic mixture of water and propylene-glycol [19] is chosen as a working fluid, which
has a low freezing point (225 K). Furthermore, MLI of aluminized polytetrafluoroethylene is
utilized on areas facing the Sun. White paint is used on the remaining structure to further
reduce radiator size.
33
5 Spacecraft Design
5.2 Subsystems Design
Radiator
Radiator
Radiator
ETCS
Radiator
ITCS
Pumps
(redundant)
Temperature
Control
Pumps
(redundant)
Heat Exchanger
Subsystems
Electronics
Environmental
Sources
Waste
Heat
Figure 5.12 Flow chart of the thermal control system
In order to increase flexibility and redundancy, four deployable carbon-carbon radiators
with an area of each 7.5 m2 have been envisioned. For a cold-case situation, e.g. on the dark
side of Mars, only two radiators are active, while two stay redundant on standby. Additionally,
the temperature of the working fluid is adapted by the mass flow rate to avoid freezing.
Carbon-carbon radiators improve the overall mass by reducing 50% mass towards common
aluminum radiators [4]. The increased development cost is included in the budget.
During the prolonged reentry due to the aerobraking maneuver, TCS has to be active.
Before the separation, approximately 137 kg water from the Cygnus is pumped into the
Dragon capsule. The total water quantity of 221 kg is used as a heat sink [31] and all heat
generated in Earth orbit and residual heat from reentry is absorbed and released as vapor
into the atmosphere. An overview of TCS is given in Table 5.22.
Table 5.22 Thermal control system summary
Mass Area/Volume TRL
Deployable radiators
232 kg
MLI insulation
58 kg
Pumps, valves, heat exchanger, etc. 153 kg
Working fluid (water & glycol)
97 kg
Heat sink reentry
88 kg
Total
628 kg
34 m2 /3.7 m3
0.62 m3
0.09 m3
0.08 m3
7
9
9
9
9
0.79 m3
-
5.2.8 Thermal Protection System (TPS)
The worst-case scenario occurs for a reentry with three passes as opposed to the nominal
reentry trajectory with two passes presented in Section 3.3. The maximal thermal stresses, as
shown in Figure 3.4, occur for a trajectory with a perigee altitude of 60 km. The computed heat
flux peak is qmax = 1878 W/cm2 and the total heat load is Qtot = 1165 MJ/m2 . Furthermore,
the TPS is responsible for a soft-landing of the capsule.
Heat Shield
To protect the spacecraft during reentry, the improved phenolic impregnated carbon ablator
(PICA-X) is chosen, which is a further development of NASAs PICA material. It is designed
34
5 Spacecraft Design
5.3 Scientific Payload
for reuse, making it suitable for multiple passes. Furthermore, it is ten times less expensive in
production than the original material. Due to proprietary issues, actual values about PICA-X
are not available. A conservative estimate for the vaporization heat is 26.516 MJ/kg which is
equal to the vaporization heat of the Apollo heat shield material Avcoat [26]. Since it is a
fairly old material, it is assumed to be a conservative worst-case of what a modern ablator
like PICA-X is capable of.
To estimate the thickness of the heat shield, a virtual ablation analysis was performed [26].
86% − 88% of the total heat load occur at a temperature above the ablation temperature of
the heat shield material [26]. With this assumption and the vaporization heat, the thickness
of the ablated layer can be calculated, which ended up being 162 mm. Using the heat shield
area and the density of the material, the mass can be determined. The heat shield has
to be protected against impacts of micrometeorites, therefore an aluminum-cover with a
jettison-mechanism is mounted on top.
Soft-Landing System
For a mass assumption of the parachute system,
Table 5.23 Thermal protection system suma design table [31] was used with a sink rate and
mary
the total mass of the return capsule. The sink
Mass Volume TRL
rate is chosen to be 11 m/s; similar to the Apollo
missions [57]. For the water landing, the system Heat shield & cover 594 kg
7
0.4 m3
9
needs to provide a stable floating position. This Parachute system 156 kg
3
Overwater
recovery
65
kg
0.2
m
9
will ensure that the crew can open the hatch
and leave the capsule. The mass of the overwater Total
815 kg
0.6 m3
recovery system is estimated with an extrapolated
assumption [37]. Due to the capability of changing the bank angle after the first pass of the
aerobraking maneuver, the landing site can be determined. An overview of the elements of
TPS is given in Table 5.23. Further simulation results are presented in the appendix [36].
5.3 Scientific Payload
Additional scientific missions do not only result in a greater benefit for the mission but will
also help to keep the astronauts busy. They could serve as something the crew feels in control
of and thus stabilize their mental state. Therefore a number of experiments was deviced that
can be implemented with little effort but still produce interesting results. It is difficult to
actually estimate space, mass and power requirements of all these experiments. However, they
were chosen as to not exceed a total of 300 kg (i.e. total habitat mass below 15 t), require a
small volume and use little electric power (thus, covered by 20% system margin).
Communication: After the succesful test of a lunar laser downlink, communication over
a distance of 1.43 AU during this mission would be a good opportunity to qualify the system
for deep space communication and allow future missions to send and receive much larger
amounts of data. Based on the LADEE spacecraft the payload weighs ∼ 30 kg and requires
an additional ∼ 50 − 140 W of electrical power. Since the Laser Communications Relay is
scheduled for 2017 it can be expected that the TRL of this technology will be sufficient.
35
5 Spacecraft Design
5.4 Systems Engineering & Budgets
Biological: Investigation of algae growth in deep space could further the use of algae
as a regenerative ressource for future missions (production of oxygen and food). The most
interesting parameters are algea growth rates and response to radiation as well as signs of
mutation. Since algae can be cultivated in completely enclosed tanks there are no interfaces
to the actual ECLSS required and measurements can be made directly at the experimental
device. More precisely a so called panel reactor which is already being researched could be
used. The size of these devices is highly scalable and the only additional electrical power
required is for a water pump and artificial lighting.
Medical: With the already issued medical equipment a number of experiments can be
conducted over the course of the mission. The words in bracktets are the names of similar
experiments aboard the ISS. Cardiac activity during exercises (Cardio ODNT) with monitoring
equipment, efficiency of drugs in micro gravity (Farma) and changes in the hematokrit
(Gematologija) using Microflow and chromosomal aberations in blood lymphocytes back on
Earth. Furthermore, the results of active dosimeters in combination with measured blood
values indicate the actual dosis the astronauts were personally exposed to. While some of
these experiments have been conducted before, this mission presents a completely new frame
to study long term effects of the deep space environment on the human body resulting in
invaluable information for more effective counter measures.
Social: Analyzing the personal diaries of the crew will give insight into mental effects of
such long term missions including actual isolation from the rest of mankind. Therefore, the
participants have to give an evaluation team the permission to access their diaries after the
mission, maybe under certain conditions regarding the publishing of private information.
Optical: Although other missions already exensively cartographed Mars the astronauts
should have access to a camera and telescope system as this might be a big motivational
factor. Thus, pictures of the flyby can be taken as well as during the mission of motives the
crew deems interesting. An interesting candidate, unless it proves to be too heavy, together
with this mission’s high gain system would be the ISS tested high resolution RAL space
camera.
5.4 Systems Engineering & Budgets
During design of the spacecraft, a systems engineering approach is applied [31]. From a
mission statement, top-level and subsequently system- and subsystem-level requirements are
derived. Trade-offs between different concepts are judged based on mass, cost, complexity
and technology readiness. Element margins are applied based on the following: 5% for
off-the-shelf items with only minor modifications, 10% for major modifications and 20% for
new developments and drastically modified elements. Furthermore, a 20% system margin is
applied to mass, power and volume of the MTV.
Table 5.24 provides budgets for the MTV, PM and mission total. The volume is divided
into pressurized, unpressurized and packed volume. The latter describes elements such as
solar arrays, radiators and other elements which are deployed after launch but have to fit
the launcher fairing. Power is differentiated based on the average and peak power as well as
waste heat. The total mass of the MTV shortly after TMI is roughly 14.5 t. The total wet
mass of the system is approximately 63 t.
36
5 Spacecraft Design
5.4 Systems Engineering & Budgets
Table 5.24 Mars18 mission summary by subsystem
Mass [kg]
Structure
AOCS
Communications
EPS
ECLSS
Radiation protection
TCS
Human factors
TPS
Scientific payload
Volume [m3 ]
Power [W]
Press. Unpress. Pack. Sum Average Peak Waste
3602
1871
149
306
3825
731
628
423
815
300
2.2
1.2
0.4
0.0
15.9
0.8
0.8
3.3
0.0
0.0
0.4
2.3
0.0
0.1
0.4
0.0
0.1
0.0
0.6
0.0
0.0
0.0
2.0
0.7
0.0
0.0
3.7
0.0
0.0
0.0
1.3
3.5
2.4
0.8
16.3
0.9
4.6
4.6
0.4
0.0
0
634
154
0
1474
2
422
367
0
0
144
2551
1036
0
5294
2
422
378
0
0
0
144
519
398
1819
0
156
535
0
0
MTV Total
MTV Total + Margin
11 899
14 988
24.9
29.5
4.0
4.8
6.4
7.6
35
42
2696
3236
7452
8943
3572
4286
Propulsion module [kg]
Mission total [kg]
48 097
63 085
Moreover, a budget per vehicle is available in Table 5.25 to determine the respective mass
and volume distribution. The propulsion module is excluded from this budget, although it can
be regarded as an autonomous spacecraft. Moreover, propellant for the docking maneuvers is
excluded as well. The living space is evaluated with NASA guidelines for manned missions.
Minimal living space at the tolerable limit is approximated with 5 m3 per person [31]. The
total available living space at TMI results in 15.7 m3 , which increases in correlation to the
mission duration and is well above the minimal required 10 m3 .
Table 5.25 Mars transfer vehicle summary by spacecraft (before TMI)
Dragon
Cygnus
Capsule
Trunk
Hab
Trunk
Mass [kg] Vol. [m3 ] Mass [kg] Vol. [m3 ] Mass [kg] Vol. [m3 ] Vol. [m3 ]
Structure
AOCS
EPS
TPS
Communications
Radiation protection
ECLSS
TCS
Human factors
Scientific payload
1215
664
78
736
114
0
5
98
219
300
0.2
1.8
0.0
0.6
0.4
0.0
0.0
0.09
0.6
0.0
579
0.0
147
35
0.0
230
254
13
0.0
0.3
0.0
0.1
2.0
0.0
1.0
0.4
0.2
0.0
1534
457
81
0
730
3590
275
467
0
0.6
0.6
0.0
0.0
0.8
15.3
0.4
3.8
0.0
0.2
1.1
0.3
0.0
0.0
0.0
0.0
0.0
0.0
Total
Total + margin
3562
4274
4.0
4.8
1259
1510
3.8
4.6
7132
8559
21.6
25.9
1.6
2.0
37
6 Programmatic Issues
In addition to the technical challenges, a mission of this scope requires thorough cost
estimation, planning and scheduling. In the following, these points are discussed. The chapter
concludes with a risk assessment including mitigation techniques.
6.1 Cost
Nowadays, spacecraft engineering is mostly driven by the influence of budgets. Back in the
1960s and 1970s with governments providing substantial budgets, cost was not as important as
it is today. The paradigm shift came in the early 1990s. A new business model in spaceflight
history became prominent: the public-private ownership. Private companies are mostly
focused on their return on investment. Governmental organizations, such as NASA and ESA,
try to reduce costs by outsourcing parts of their development and production structures.
These points make cost estimation very important but also very challenging.
Most of the common parametric cost estimating tools are driven by weight. Nevertheless,
there are also other factors such as mission difficulty. Despite that, the biggest problem is
access to these cost models. Cost databases such as CADRe from NASA are not available
to third persons. Others, such as PRICE-H, are software-based and not open source.
Notwithstanding, the following cost estimation is based on proven methods that will give a
detailed breakdown of the mission cost for Mars18.
6.1.1 Introduction of Applied Cost Estimating Tools
Cost estimating tools can be subdivided into two main categories. The first category is called
the commercial off-the-shelf (COTS) method. The second is called government off-the-shelf
(GOTS) method. Among other, COTS includes [56]:
ˆ TransCost Model
ˆ Unmanned Space Vehicle Cost Model (USCM)
ˆ Small Satellite Cost Model (SSCM)
ˆ PRICE-H
ˆ SEER-H.
Among other methods, GOTS estimating tools include:
ˆ NASA/Air Force Cost Model (NAFCOM)
ˆ Aerospace Launch Vehicle Cost Model (LVCM)
ˆ Advanced Missions Cost Model (AMCM).
The cost approximation for the Mars18 mission focuses on the TransCost Model, USCM,
AMCM and Analogy/Build-up.
38
6 Programmatic Issues
6.1 Cost
TransCost Model is designed to be applied in the early mission phases for cost estimation
of the development and production of space vehicle systems. TransCost uses a parametric
methodology with cost estimating relationships that are deduced from costs for European
and US space vehicles of the last 50 years. Moreover, TransCost supports the calculation of
operating costs at the ground during the mission. It will be used to approximate the ground
and flight operation cost of Mars18. The estimation is based on the fourth and final edition
of TransCost [28].
USCM is a parametric handbook and cost model for estimating costs for unmanned space
vehicle missions. The model is based on a NASA, military and commercial satellite database.
USCM facilitates the calculation of recurring and non-recurring mission costs. The model is
mostly driven by weight and was established by the US Air Force because of its possibility
to estimate the costs of communication subsystems [56]. Therefore, USCM will be used to
calculate, among other things, the communication subsystem.
AMCM is developed by the Johnson Space Center. Compared to other cost models such
as USCM, it does not solely focus on weight. Moreover, it includes five additional input
parameters: quantity of developed and produced units; specification factor for mission level;
year of first operation; level of design heritage (block number) and difficulty [31]. The model
allows calculation of the costs of the flight hardware for development and production phase.
The model is used to determine costs of those subsystems that could not be calculated by
the analogical methods.
Analogy/Build-up is used to determine costs on the basis of historical data such as work
hours and bills of material. Analogy is applied to adjust or extrapolate data. Estimations on
Analogy/Build-up are used for Mars18, when information is available and can be adjusted or
extrapolated. This estimation method is mainly based on literature analysis.
6.1.2 Overview of Estimated Costs
In the following, cost approximation is presented. A more detailed description of how the
costs are composed and cost drivers are shown in the appendix. Table 6.1 specifies the
development and production costs for all different subsystems.
Table 6.1 Development and production costs using applied cost estimation tools
Subsystem
Structure
AOCS
EPS
Propulsion
TPS
Communication
Radiation
ECLSS
TCS
Human factors
Sum
Costs [$] Applied cost estimation Share [%]
464,342,590
318,128,143
15,960,936
684,244,933
5,423,893
39,604,088
3,004,707
1,652,027,918
715,027,857
36,027,900
Analogy/build-up
AMCM
Analogy/build-up & USCM
Analogy/build-up
Analogy/build-up
USCM
Analogy/build-up
AMCM
AMCM
Analogy/build-up
3,933,792,966
10.8
7.4
0.4
15.9
0.1
0.9
0.1
38.5
16.6
0.8
91.6
39
6 Programmatic Issues
6.1 Cost
Cost estimations of the subsystems differ significantly. The ECLSS, for example, causes
the highest costs due to high complexity and necessary reliability. Cost of ground and flight
operation as shown in Table 6.2 claim only 8.4% of the total costs, which may seem small
but results from the knowledge of a definite end. Operation costs for missions with a definite
end are less than 20% [31]. Finally, the total costs are given in Table 6.3.
Table 6.2 Ground and flight operation costs calculated by TransCost
Costs [$] Share [%]
Term
Prelaunch ground operation cost
Crewed vehicles mission cost
Launch, ascent and descent operation cost
Recovery operation
Launch site user fee
Commercialization cost assessment
Sum
21,027,201
294,232,708
16,128,000
84,807
347,779
28,800,000
0.5
6.9
0.4
0.0
0.0
0.7
360,620,495
8.4
Table 6.3 Total costs of Mars18 mission
Cost [$] Share [%]
Term
Development and production 3,933,792,966
Ground and flight operation
360,620,495
Sum
4,294,413,461
91.6
8.4
100.0
6.1.3 Life Cycle Costs
4000
150
Monthly
Total
3000
100
2000
50
0
1000
Phase 0/A Phase B
Phase C
Phase D
Figure 6.1 Life cycle costs of Mars18
40
Phase E
0
Total cost [M$]
Monthly cost [M$]
Costs per phase vary from project to project. To portray an overview of cost development
throughout the project, a cost diagram is provided in Figure 6.1. Cost trend is displayed by
an upward curve which is determined by the accumulated cost throughout the project. Cost
distribution of each phase is provided with aid of a bar chart. In this case, costs per phase
are calculated on a monthly basis which consists of 21.66 work days.
6 Programmatic Issues
6.2 Roadmap & Schedule
6.2 Roadmap & Schedule
The most critical parts of the mission are technologies which require an increase in TRL
and launch preparation and scheduling. A detailed plan to develop all required launch
technologies and gain the necessary experience within the given time frame is presented in
Table 6.4. The overall imperative for the development of this mission was to use as much
flight-rated equipment as possible. Nonetheless, in some cases systems have to be employed
that still require a certain amount of research and development to make the whole mission
feasible. Care was taken to ensure that only a few of these technologies are currently below
an estimated TRL 6, making it realistic that they can be fully developed by the first launch
in 2018. The Asymmetric Payload Fairing (APLF), which has a TRL of 4, is used due to
its advantages for the mission, which are explained more detailed in the appendix [36]. To
ensure that it is verified and tested until 2018, all measures are undertaken and even a test
flight is foreseen. In addition, the beginning of development is scheduled as soon as possible,
to have an adequate amount of time to solve unexpected problems.
One of the biggest problems arises from ECLSS, specifically with the development of
VPCAR/PMWC and life-cycle testing. Also, the development and integration of the novel
AOCS software will consume a considerable amount of time. Moreover, the heavy modifications
on the Dragon trunk and the development of the personal dosimeter through the DLR are
crucial elements that have to be completed in time.
In the case of VPCAR and PMWC, to achieve the necessary increase in TRL scaling of
the technology was foregone. Thus, the resulting machine will be slightly oversized but more
likely ready for operation. Usually the complete ECLSS has to be tested with two life-cycles.
To stay within the given time frame, it was decided to conduct only one complete life-cycle
test. This can be justified with two arguments: first, for the most parts (12 out of 14) proven
technologies are used and only slightly modified. Second, should there be modifications
necessary during the testing, they will be implemented directly and the results extrapolated.
Since MPC has been used extensively in large and complex systems already the adaption
to the spacecraft within the given time is possible. Furthermore, as mentioned before, remote
access to the code from Earth will allow for in-flight debugging and increased safety should
something actually go wrong. Also, the usage of waste products for spacecraft control and a
backup PID system can contain an absolute failure of the new software.
The overall schedule for the development is presented in 6.2. A more detailed timetable can
be found in the appendix [36]. In this copy, only the rough outline and the already mentioned
critical parts of the subsystems are presented. Development periods have been estimated
based on experience and heritage. It is clear that a launch in 2018 is very ambitious but with
proper planning and collaboration between all involved parties it should be possible.
The launch manifest in 6.4 has been developed based on current research programs and
requirements identified in this report. The entire plan integrates well into existing schedules
of agencies and companies. Therefore, crucial scientific/engineering knowledge and flight
qualifications with regard to space transportation can be obtained. At the same time, the
overall cost is kept low and already arranged plans are not majorly interrupted.
41
6.2 Roadmap & Schedule
Figure 6.2 Critical technologies
6 Programmatic Issues
42
6 Programmatic Issues
6.3 Risk Management
2016
2015
2014
Table 6.4 Launch system development, integration & manifest
3. Q Contract & order of Asymmetric Fairing development
Contract & order for KURS integration
4. Q
Falcon Heavy testflight
Search partner for asymmetric fairing launch test
1. Q
Begin of development of asymmetric fairing
COTS-6 with KURS dock (Dragon)
2. Q COTS-7 with KURS dock (Cygnus)
3. Q
Rendezvous & docking of Dragon and Cygnus
Falcon Heavy first launch Cape Canaveral
4. Q
Atlas V551 man-rating - Dreamchaser first launch
Asymmetric fairing - first test launch
2017
1. Q Atlas V551 man-rating - CST-100 first launch
2. Q Asymmetric fairing - first cargo test with partner
3. Q Integration of payload and launcher
2018
4. Q
Launch PM with Falcon Heavy
Launch MTV with Atlas V551
Rendezvous & docking
1. Q In-space final system test
Trans-Mars injection
6.3 Risk Management
A manned Mars flyby poses significant challenges and risks. Some of them are referenced
by numbers in brackets and can be found in tabelle riskmatrix. One of the most prominent
risk in spaceflight is the chance of a single point of failure (SPOF). The risk of a SPOF is
present in every technical aspect of the mission. More specifically SPOFs could result in a
LOM or LOC of which the latter is unacceptable. Furthermore, a cancellation of the mission
due to financial reasons is a possibility (16). This is especially a threat, if total mission
costs are gravely underestimated, as it is often the case in such large endeavors. Lack of
public attention and/or support could also endanger the entire execution of the mission, i.e.
commencement of the mission and continued ground segment operations. Lastly, missions
executed by humans are always endangered by human failure.
6.3.1 Technological Risks
In general, SPOF can be avoided through redundancies, heritage and testing. This is the most
simple approach and has been implemented for all subsystems of the spacecraft. However,
there are several facets of the mission that can not be treated that way and some of them
will be presented in the following. Table 6.5 shows a more comprehensive summary of risks
43
6 Programmatic Issues
6.3 Risk Management
during the mission. Whereever applicable bracketed numbers in the upcoming paragraphs
will reference the content to the table. A complete breakdown of the numbers can be found
in the appendix [36].
During TMI a failure of either PM would lead to a LOM and, if not designed carefully, to a
LOC during the second propulsive maneuver (7). A LOM is unlikely since the PM is regarded
as highly reliable and has been tested in many missions. Before reaching hyperbolic velocities
during the second stage firing the crew has ∼ 490 s of time to trigger an emergency shutdown
and remain in a highly elliptical orbit around Earth. Thus, a LOC has been avoided and a
rescue mission is still possible.
Changes in the requirements regarding heat control due to a varying solar constant and
albedo radiation have to be compensated for by the TCS. Failure of this system could result
in a LOC. As a result, the radiators have been designed to allow for a large operating range
regarding excess heat without failure by frozen cooling fluid. Since the spacecraft will return
to Earth with a high hyperbolic velocity special precautions have to be taken for reentry.
The heat shield has to be in an acceptable condition after being exposed to interplanetary
space for the entire mission duration (27). The structure has to be able to withstand the heat
and mechanical loads and posses the right aerodynamic properties (18,19). A malfunction in
either of those systems would result in a LOC. Therefore, it has been decided to perform an
aerocapture and multiple aerobrakes before reaching the ground to reduce the peak heat and
mechanical loads. Before performing the Earth-return manoveurs the heat shield is covered
by an aluminum shell.
To ensure that no LOM or LOC occurs due to failure of minor parts a 3D printing
system will be part of the mission (28). Thus, the crew can reproduce and replace small
malfunctioning parts.
The mission risk and uncertainty is actually very low. In-orbit assembly (8,9), gravity
assists (26) and aerobrakes (24,25,28) have been conducted in the past several times without
failure. Moreover, the size of the Dragon capsule puts it in the lowest risk class for an
aerocapture. Moreover, the risk of failure associated with the total habitat mass is considered
to be below 10% based on heritage data [49].
6.3.2 Programmatic risks
A big, but often neglected, risk are programmatic issues, such as public support and finances
(16,17). While the general opinion is that interest in space exploration has been declining,
the experience gathered during this project suggests othwerwise. However, if a focused effort
is made, it should be possible to convince large parts of the population of the usefulness
(scientific progress, jobs, competition, long term investments) of a manned Mars mission.
This assessment stems from the evalutation of the efforts made by Mars18 to gain public
attention and support. Without spending any money, the Mars18 website had ≈ 429 visits
per month since November 2013. The Facebook page currently has 173 followers and a reach
of ≈ 212 per post. While these numbers are certainly not representative they show that
people are still fascinated by space exploration and its endless possiblities.
To mitigate the danger of a financial overkill in a late project phase a conservative hybrid
approach was employed in the cost estimate of this report, utilizing up-to-date information,
heritage and cost models. This method ensures that unknown factors are accounted for
44
6 Programmatic Issues
6.3 Risk Management
realistically while not overshooting the known costs. The hope is to keep the actual costs
after the program has been completed below the number given in this report. This might
be possible through current developments in the commercial space sector. In general, these
developments will probably lead to an overall increased launch frequency and decreased costs.
Additionally, it is absolutely necessary to make the financial decisions and necessities as
transparent to the public as possible to avoid frustration. Thus, it is more likely that the
program budget will not be cut severely after the initial planning phase. Furthermore, the
financial participation of stakeholders should be considered. For example, selling slots to
perform scientific experiments during the mission or broadcasting rights can help to cover
some expenses.
6.3.3 The Human Factor
Every man-operated system is subject to human failure. This factor can not be neglected in
the risk assessment. Several different approaches exist to mitigate this danger: the first and
most obvious one is a well-trained and selected team. This is ensured by the selection and
training process presented in Section 4.1 (5). A second approach is the implementation of
plans that take effect when critical situations emerge, e.g. outbreak of a fire (11,15), short
circuits, gas leakage or malfunctions (12,13,22). The risks and their effects have to be rated
by timeframe, likelihood, severity (e.g. impact on other system, LOM, LOC) and uncertainty.
Furthermore, the ground segment has to have a possiblity of overriding command inputs
given to the spacecraft by the astronauts. This could become necessary when unexpected
behavior caused by psychological changes or extreme situations make the astronauts choices
untrustworthy.
Unfortunately there are aspects like very hard SPEs (6), high energetic micro meteorites or
space debris impacts (18) that can neither be foreseen nor can their effects be avoided. These
events could lead to the death of part or the whole crew. This means that both astronauts
have to be aware of the possibility that they might have to deal with a deceased partner
during the mission (4). If that was the case, a nontransparent vacuum bag is provided in
which the body can be stowed. All mentioned aspects have to be ranked in likelihood and
severity to allow for a justifiable decision whether the risk is acceptable.
Point (23) represents a case in which the Mars mission will fail. This point had to be
included since a realistic view at the endeavor shows that with such a tight schedule failure
due to development problems is a possibility. Thus, rigorous project management, more
feasability analysis and well worked out schedules are an absolute must for this mission.
Table 6.5 Risk matrix
very likely
likely
possible
unlikely
rare
16,24
18
14
1,10
minor moderate
17,26,28
5
insign.
45
6,15,27
4,9,20
major
23
7,22,25
2,3,8,11,12,13,19,21
catastrophic
7 Conclusion
The foregoing review shows the feasibility of a manned Mars flyby for the year 2018 by
presenting a concept that accounts for the following properties: Requirements are identified
and implemented. Risks are discussed, rated and minimized. At the same time the total cost
is kept as low as possible by using or modifying existing systems and reducing the overall
habitat weight. Synergies are employed wherever it was deemed possible to decrease system
weight and complexity. Reasonable conservative estimates were assumed throughout the
whole mission to account for all avoidable worst case scenarios. The mission has a duration
of 501 days and offers space for two astronauts, a man and a woman. A modified Dragon
and Cygnus are assembled in LEO with the PM before TMI takes place. The trajectory and
several key moments during the mission can also be followed in a video in the appendix on
the Mars18 homepage [36]. Returning to Earth, an aerobreak maneuver is performed.
The final results of this report also show that a mission to Mars in 2018 is an extremely
challenging idea. In order to increase the likelihood of success and decrease the risks without
postponing the mission too much a Mars flyby in 2021 could be considered. Thus, there is
more time available for testing, development and the qualification of additional heavy-lift
launch systems like the SLS.
Acknowledgements
It is no small feat to design a manned mission to Mars in such a short time frame even with
a large group as ours. With that said, none of this would have been possible, without the
support and help from many companies, organizations, societies, institutes and individuals.
We would like to thank both Brainlight and Airbus Defense and Space for their professional
guidance and support. In addition to contributing their time and knowledge, Astos Solutions
also provided software licenses. We also had a great deal of support from the University of
Stuttgart community.
The importance of the knowledge gained through exchanges with the Institute of Space
Systems cannot be overstated. In particular, we would like to sincerely thank Tilman Binder,
Emil Nathanson and Christine Hill. We also greatly appreciated the encouragement and
interest of our peers from Constellation and the aerospace society DGLR-BG Stuttgart.
Finally, it behooves us to recognize and thank a handful of individuals whom critically
facilitated our work. We are grateful for the knowledge and assistance from Thomas Berger
and Daniel Matthiä from the German Aerospace Centre (DLR) as well as Jürgen Herholz,
Karin Schlottke and Patrick Wang.
46
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