IUT Astronaut - Members Area

Project Title
Human Expedition on Mars
Timeline 2018
Team Name
IUT Astronaut
Institution
Islamic University of Technology (IUT)
A subsidiary organ of
Organization of Islamic Cooperation (OIC)
Board Bazar, Gazipur, Dhaka, Bangladesh
Team Members
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
MD. Wasif Zaman (Team Leader)
Fahad Al Mamun
Faisal Hussain
M. Saiful Bari
MD. Abdullah Al Joha
MD. Muhtady Muhaisin
Mohammad Abdullah Matin Khan
Muhammad Usama Islam
Raihan Uddin Ahmed
Shabab Bin Karim
Tanzil Bin Hassan
Team Coordinator
Dr. Khondokar Habibul Kabir
Assistant Professor (Department of EEE, IUT)
Ph.D. (Osaka University), M.Sc. Engg. (Osaka University), B.Sc. Engg.(IUT)
2
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Abstract
The paper is intended to provide an in brief, safety-maintained and economic design for an
expedition on MARS. The whole paper is segmented in various sections to describe
different aspects efficiently and adequately. Also Illustrative images and flowcharts are
provided to explain different ideas. Several Innovative ideas are presented in the design
and the control algorithm is chosen by weight-based comparison. Human life issues are
associated and the costing is compromised to some degree by considering safety issues.
3
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Contents
Preface
06
Chapter 1
1.1
1.2
1.2.1
1.2.2
1.2.3
1.3
1.4
1.5
1.6
Payload Transfer Maneuver
Introduction
Description of Payload Transfer Maneuver
External Tank
Solid Rocket Booster
The Orbiter
Earth to LEO Maneuver
Orbital Maneuvering System
Reaction Control System
Electrical Power Distribution System
07
07
07
08
09
10
11
11
11
Chapter 2
2.1
2.2
2.3
Launching and Landing Maneuver with In Space Trajectory
Launching Maneuver
Landing Maneuver
Time Schedule
13
17
18
Chapter 3
3.1
3.2
3.3
Mars Surface, Environment & Pressurized Rover
Landing Site
Mars Atmosphere
Pressurized Rover
19
20
21
Chapter 4
4.1
4.2
4.3
4.4
4.5
4.6
4.7
4.8
Engine and Fuel
Introduction
VASIMR Engine
Engine Subsystem
Nuclear reactor
Fuel for VASIMR Engine
Human Mission to Mars with 12 MW Input Power
Recent Development of VASIMR Engine
Observations and Recommendations
23
23
24
26
26
27
27
28
Chapter 5
5.1
5.2
5.3
5.4
5.5
5.6
5.7
Electrical Power System
Introduction
Purpose of the Power System
Design Process
Power Sources
Power Storage and Secondary Power Sources
Power Distribution
Power Regulation and Control
29
29
29
29
31
32
32
4
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Chapter 6
6.1
6.2
6.3
6.4
6.5
Control System
Introduction
Control System Architecture
Attitude and Articulation Control
Attitude and articulation control subsystem
Thruster Operations
33
34
35
35
37
Chapter 7
7.1
7.2
Space Communication
Introduction
Description
38
38
Chapter 8
8.1
8.2
8.3
8.4
Human Health Risks & Solutions
Introduction
Physical Threats
Radiation challenge
Psychological aspects
40
40
42
43
Conclusion
44
Acknowledgement
45
References
46
5
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Preface
The nature of mankind is to conquer. Since ancient time, man has always dreamt of
conquering the sky. As the days passed, the dream grew bigger. The Moon eventually
came into the grasp of mankind. But the journey didn’t stop here.
For last few decades, scientists have tried to land unmanned space rovers to the surface
of our neighboring planet, the Mars. Mars rover Spirit and Opportunity have already been
proved as projects of tremendous success. Now the time has come to conquer the red
planet. Like its name, the Roman god of war is not easy to defeat. It has built many
obstacles on the way of conquering it.
This paper proposes a full scheme of different aspects which are essential for a safe a
cost efficient round trip to Mars. In our proposal, VASIMR [2] technology was chosen for
inter planetary round trip where argon will be used as fuel. All necessary mission
appliances will be sent to the International Space Station (ISS) first. There, whole
infrastructure will be built. Finally two crew members will be sent to the ISS and they will
depart for mars. After 91 days in space journey astronauts will reach LMO and then will
descent to mars surface for 2 days. There they will explore the surface and will collect
necessary data which will help in reaching the ultimate goal i.e. establishment of human
colony on mars. Then they will come back to the mother ship and return to earth by 174
days approximately.
In chapter 1, payload maneuver system is discussed elaborately along with distribution of
payload components in three separate launches. In chapter 2, landing and launching
maneuver with in time schedule is discussed. After that comes the discussion of mars
surface and pressurized rover. In the nest chapter VASIMR engine with RMBLR reactor
and fuel is mentioned with working principle. In chapter 5, electrical power system is
discussed which is followed by Control system in chapter 6. In the later portion, Space
communication is focused in chapter 7 and most importantly in chapter 8, all human
concerns i.e. space radiation, hazard and their probable solutions are discussed
elaborately.
It is to mention that, most of the technological references mentioned in this paper are
practical. But the concepts of VASIMR engine and RMBLR reactor are still in the
laboratory research level. If proper efforts are given, they can be made functional for mars
mission within 2018.
6
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Chapter 1
Payload Transfer Maneuver
1.1. Introduction
The first topic that is going to be discussed is payload transfer maneuver as it’s necessary
to send all the mission equipment to the International Space Station (ISS) first.
1.2. Description of Payload Transfer Maneuver
Various required parts of the mission will be sent to the ISS separately where they will
rendezvous and begin their space travel to mars. 3 separate launches will be needed to
send all the parts to the ISS. The payloads will contain:
 Mars Travel Vehicle(MTV)
 VASIMR engine with fuel
 Power supply reactor for VASIMR
 Control Room
 Landing Module
 Food, medicine
 Human
Current convention for transferring payload to ISS will be used where the orbiter along with
external tank and SRB will launch from earth and transfer the payload to ISS.
In the payload transfer procedure there are three main parts. They are
 External Tank
 Solid Rocket Booster(SRB)
 The Orbiter
1.2.1 External Tank
The three main components of the External Tank are an oxygen tank, located in the
forward position, an aft-positioned hydrogen tank, and a collar-like intertank, which
connects the two propellant tanks, houses instrumentation and processing equipment, and
provides the attachment structure for the forward end of the solid rocket boosters. The skin
of the External Tank is covered with a thermal protection system that is a 2.5-centimeter
(1-inch) thick coating of spray-on polyisocyanurate foam.
The External Tank includes a propellant feed system to duct the propellants to the Orbiter
engines, a pressurization and vent system to regulate the tank pressure, an environmental
conditioning system to regulate the temperature and render the atmosphere in the intertank area inert, and an electrical system to distribute power and instrumentation signals
and provide lightning protection as shown in Figure 1.1.
7
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Figure 1.1: Cross section of external tank
The tank’s propellants are fed to the Orbiter through a 43-centimeter (17-inch) diameter
connection that branches inside the orbiter to feed each main engine. 234,265 lbs of liquid
hydrogen and nearly 1.4 million lbs of liquid oxygen stored in the external tank.
1.2.2. Solid Rocket Booster
Two solid rocket boosters provide the main thrust to lift the space shuttle off the pad. They
provide thrust up to an altitude of about 150,000 feet each booster producing a thrust of
3,300,000 pounds at lunch. Each is 149.16 feet long and 12.17 feet in diameter. Each SRB
weighs approximately 1,300,000 pounds at launch. The propellant for each solid rocket
motor weighs approximately 1,100,000 pounds. The fuel the SRBs burned is called
Ammonium Perchlorate Composite Propellant (APCP). It consists of ammonium
perchlorate (oxidizer, 69.6% by weight), aluminium (fuel, 16%), a polymer (12.04%), an
epoxy curing agent (1.96%), and iron oxide (0.4%). The aluminum was quite powerful as
a fuel but difficult to accidentally ignite. The two SRBs provide 71.4 percent of the thrust at
the lift off and during the first stage ascent as shown in Figure 1.2.
Figure 1.2: Cross section of SRB
8
Human Expedition on Mars Timeline 2018 by IUT Astronaut
The SRB Hydraulic system is to supply the required hydraulic flow and pressure to extend
and retract the actuator piston. The end of the piston is attached to the nozzle of the solid
rocket motor to provide thrust vectoring during the mission. This system is called Thrust
Vector Control (TVC), and it provides 80% of steering for the integrated vehicle during
ascent. A similar system vectors the main engine nozzles, providing the other 20% of the
steering control.
1.2.3 The Orbiter
The orbiter is the manned spacecraft of the Space Shuttle’s three main components. It can
transport into near earth orbit (115 to 690 miles from the earth’s surface) cargo weighing
up to 56,000 pounds, and it can return with up to 32,000 pounds. This cargo, called
payload, is carried in a bay 15 feet in diameter and 60 feet long. As shown in figure 1.3,
the major structural sections of the orbiter are:
 the forward fuselage, which contains the pressurized crew compartment
 the mid fuselage, which contains the cargo bay
 the aft fuselage, from which the main engine nozzles project
 the vertical tail, which serves as a speed brake used during entry and landing
Figure 1.3: Cross section of Orbiter
The Space Shuttle orbiter has three main engines weighing 7,000 pound each. They are
very sophisticated power plants that burn liquid hydrogen with liquid oxygen, both from the
external tank (ET). The main engines are located in the aft (back) fuselage (body of the
spacecraft). They are used for propulsion during launch and ascent in to space with the aid
of two powerful solid rocket boosters (SRBs). The main engines provide 29% of the thrust
needed to lift the shuttle off the pad and into orbit. Each engine can generate almost
400,000 pounds of thrust at liftoff. MMH (Monomethyl-Hydrazine) and N2O4
9
Human Expedition on Mars Timeline 2018 by IUT Astronaut
(Nitrogentetroxide) for the OMS engines and the RCS. This is used for maneuvers in orbit
and for the attitude control system. MMH and N2O4 have the advantage that both react on
contact with each other and engines using them can be started and more important
restarted easier. But also, both chemicals are pretty toxic.
1.3. Earth to LEO Maneuver [3]
The Shuttle’s three main engines (SSMEs) are sequentially started at approximately the T7 second mark. When the engine controllers indicate that they are all running normally, the
twin solid rocket boosters (SRBs) are ignited at the T-0 mark. A sequence of events occurs
within a few seconds before launch, leading up to SRB ignition and liftoff.
Terminal Countdown
Arm Solid Rocket Boosters
Auto Sequence Start
Main Engine Start
SRB Ignition
Liftoff
-9.00.0
-5.00.0
-0.31.0
-0.06.0
0.00.0
0.00.3
The three Space Shuttle Main Engines, in conjunction with the Solid Rocket Boosters,
provide the thrust to lift the Orbiter off the ground for the initial ascent. The main engines
continue to operate for 8.5 minutes after launch, the duration of the Shuttle’s powered
flight. The SRBs together burned 2.2 million lbs of fuel during the first 2 minutes and 13
seconds of flight before falling away.
SRB separation is initiated when the three solid rocket motor chamber pressure
transducers are processed in the redundancy management middle value select and the
head-end chamber pressure of both SRBs is less than or equal to 50 psi. A backup cue is
the time elapsed from booster ignition.
The separation sequence is initiated, commanding the thrust vector control actuators to
the null position and putting the main propulsion system into a second-stage configuration
(0.8 second from sequence initialization), which ensures the thrust of each SRB is less
than 100,000 pounds. Orbiter yaw attitude is held for four seconds, and SRB thrust drops
to less than 60,000 pounds. The SRBs separate from the external tank within 30
milliseconds of the ordnance firing command. Exactly 295 seconds after they separate
from the vehicle, both SRBs fall into the Atlantic Ocean, where they are recovered for
reuse as shown in Figure 1.4.
After the solid rockets are jettisoned, the main engines provide thrust which accelerates
the Shuttle from 4,828 kilometers per hour (3,000 mph) to over 27,358 kilometers per hour
(17,000 mph) in just six minutes to reach orbit. They create a combined maximum thrust of
more than 1.2 million pounds.
The main Orbiter carries the external tank piggyback to near orbital velocity, approximately
113 kilometers (70 miles) above the Earth. The now nearly empty tank separates and falls
in a preplanned trajectory with the majority of it disintegrating in the atmosphere and the
rest falling into the ocean. Then when the orbiter reaches the orbit the Orbital Maneuvering
System takes on the control
10
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Figure 1.4: Payload Attachment Scheme
1.4. Orbital Maneuvering System
The Orbital Maneuvering System (OMS) is one of two systems that allow the shuttle
orbiter to maneuver once it is in space. This system is composed of two small rocket
engines that are located in the aft fuselage, just either side of the upper main engine. The
OMS is vital to getting the orbiter into its correct orbit (includes putting the orbiter into its
proper orbit before reentry). Approximately 10.5 minutes after liftoff (about 1 1/2 minutes
after the external tank is jettisoned), the OMS rockets are fired to help put the orbiter into a
low orbit. Around 45 minutes after liftoff, they are fired again to elevate the orbiter into its
mission orbit, roughly 250 miles above the surface of the earth. NASA uses special liquid
fuel that does not require a spark igniter. Liquid nitrogen tetroxide produces an explosion
when mixed with liquid monomethyl hydrazine. This fuel and oxidizer combination allows
these two rockets to produce 6,000 pounds of thrust each. This is enough thrust to change
the orbiters acceleration by 2 ft per second squared or change its velocity by 1000 feet per
second.
1.5. Reaction Control System
The reaction control system (RCS) on the orbiter is very similar to the OMS. It helps to
maneuver the orbiter in more delicate situations. Two prime examples of when the RCS is
used is when the orbiter is docking with the International Space Station (ISS) or capturing
a satellite to be repaired. The RCS consists of 44 small nozzles that are fueled by the
same liquid nitrogen tetroxide and monomethyl hydrazine combination as the OMS. With
the help of OMS and RCS the orbiter fixes its trajectory towards the ISS and delivers the
payload there. Then the Orbiter again returns back to the earth for reuse.
1.6. Electrical Power Distribution System[8]
The EPS consists of three subsystems: power reactant storage and distribution, fuel cell
power plants (electrical power generation) and electrical power distribution and control.
The PRSD subsystem stores the reactants (cryogenic hydrogen and oxygen) and supplies
11
Human Expedition on Mars Timeline 2018 by IUT Astronaut
them to the three fuel cell power plants, which generate all the electrical power for the
vehicle during all mission phases. In addition, cryogenic oxygen is supplied to the
environmental control and life support system for crew cabin pressurization. The hydrogen
and oxygen are stored in their respective storage tanks at cryogenic temperatures and
supercritical pressures. The storage temperature of liquid oxygen is minus 285 F and
minus 420 F for liquid hydrogen.
The EPDC subsystem distributes the 28 volts dc generated by each of the three fuel cell
power plants to a three-bus system that distributes dc power to the forward, mid-, and aft
sections of the orbiter for equipment in those areas.
The EPDC subsystem controls and distributes electrical power (ac and dc) to the orbiter
subsystems, the solid rocket boosters, the external tank and payloads. Power is controlled
and distributed by assemblies.
12
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Chapter 2
Launching and Landing Maneuver with
In Space Trajectory
2.1. Launching Maneuver
2.1.1. Introduction
In our proposal, for sending humans to Mars, the crucial first step is launching the
spacecraft into a low Earth orbit (200 to 500 kilometers up). The basic problem is that any
manned craft using present-day propulsion technologies (chemical propulsion technology)
will need a huge supply of propellant to get to Mars and hence will be extremely heavy:
Possibly 250 metric ton. VASIMR engine can be used only after reaching the Lower Earth
Orbit. No launch vehicle now in use can lift that much mass into orbit. So Mars craft would
have to be launched in stages and then assembled in orbit, preferably through docking
maneuvers that could be controlled from the ground. So we propose for two spacecraft: A
crew transfer vehicle (CTV) and a descent/ascent vehicle (DAV). The CTV carries the
astronauts to Mars orbit along with DAV. The astronauts will then board on DAV, descend
to the surface, stay for 2 days (optimal) and return to the CTV parked in the orbit. The
CTV, which has been waiting in orbit, brings them home. Chemical Trans Mars Injection
(TMI) will be launched from KSC and assembled in LEO.
Assembling the craft at the International Space Station would be inefficient because the
launching pad has an inclination of 51.6 degrees; from the launch facilities at Cape
Canaveral, Fla., it is easiest to boost payloads into an orbit with a 28.5-degree inclination.
The space shuttle could transfer the crew to the Mars craft once it was completed. To
simplify the assembly, the number of launches and orbital rendezvous would have to be
minimized. We must use an Expandable launching system (ELS) for carrying our payloads
to LEO. The ELV (expandable launching vehicle) will be human-rated.
2.1.1. Heavy Lift Launching
The SLS heavy-lift launch vehicle is essential to NASA’s deep-space exploration
endeavors. The system will be flexible and include multiple launch vehicle configurations.
The SLS will carry crew, cargo and science missions to deep space. The 70-metric-ton(77 ton) configuration will lift more than 154,000 pounds and will provide 10 percent more
thrust than the Saturn V rocket while the 130-metric-ton-(143 ton) configuration will lift
more than 286,000 pounds (130 mT) and provide 20 percent more thrust than the Saturn
V.
13
Human Expedition on Mars Timeline 2018 by IUT Astronaut
SLS have launching capability twice a year with proper funding and fuel supply which is
shown in Figure 2.1.1.
SLS 130-metric-ton
Evolved Configuration
Figure 2.1.1: Heavy Lift Launch Vehicle SLS
We propose to send the whole project dividing in 2 parts and use the 130 mT configuration
SLS.
1. 1st launch will include MTV, DAV, Rocket structure, Thruster radiator, Reactor
radiator making 125 mT of cargo.
2. The next launch will include Reactor, Thruster and fuel Tanks Along with the E-M &
M-E propellant. This launch will carry a little more than 130 mT.
3. And finally mission crew with necessary medical and food supplies. To minimize the
cost we can send the mission crews along with regular ISS transportation vehicle.
2.1.2. Conjunction Class [3]
For high-thrust rockets, the most fuel-efficient way to get to Mars is called a Hohmann
transfer. It is an ellipse that just grazes the orbits of both Earth and Mars, thereby making
the most use of the planets’ own orbital motion. The spacecraft blasts off when Mars is
ahead of Earth by an angle of about 45 degrees (which happens every 26 months). It
glides outward and catches up with Mars on exactly the opposite side of the sun from
Earth’s original position. Such a planetary configuration is known to astronomers as a
conjunction. To return, the astronauts wait until Mars is about 75 degrees ahead of Earth,
launch onto an inward arc and let Earth catch up with them. Each leg requires two bursts
of acceleration. From Earth’s surface, a velocity boost of about 11.5 kilometers per second
breaks free of the planet’s pull and enters the transfer orbit. Alternatively, starting from low
Earth orbit, where the ship is already moving rapidly, the engines must impart about 3.5
kilometers per second. At Mars, retrorockets or aero braking must slow the ship by about 2
kilometers per second to enter orbit or 5.5 kilometers per second to land. The return leg
reverses the sequence. The whole trip typically takes just over two and a half years: 260
days for each leg and 460 days on Mars. In practice, because the planetary orbits are
14
Human Expedition on Mars Timeline 2018 by IUT Astronaut
elliptical and inclined, the optimal trajectory can be somewhat shorter or longer. But since
a long stay on mars is risky for human life opposition class/ heliocentric-transfer orbit is
much more favorable.
2.1.3. Opposition/Heliocentric-transfer [5]
To keep the trip short, various planners traditionally considered opposition-class
trajectories, so called because Earth makes its closest approach to Mars—a configuration
known to astronomers as an opposition—at some point in the mission choreography. Mars
is near opposition at the midpoint of a short duration mission, and Mars passes through
conjunction in the middle of a long stay time mission as shown in Figure 2.1.3. Since our
proposal is an orbit to orbit transfer, it can also be called as a Stopover mission.
For stopover missions, the crew begins on Earth, but the transfer vehicle begins in Earth
orbit. After an in-orbit rendezvous the crew and transfer vehicle leave for Mars, where the
crew lands and the transfer vehicle is placed in orbit. Following another in-orbit
rendezvous, the crew and transfer vehicle depart for Earth, where the crew lands and the
transfer vehicle is placed in Earth orbit for reuse.
Figure 2.1.3: Opposition/ Heliocentric-Transfer Orbit
To reach an outer planet i.e. Mars a space craft must be launched from earth at a velocity
greater than the planetary escape velocity Ve. This extra velocity changes the speed of the
space craft while in the heliocentric orbit around the sun. Given the proper velocity a body
goes into a heliocentric orbit that carries it to the destination planet along its new elliptical
path. To reach the outer planets a space craft must be launched fractionally faster than the
planetary escape velocity and must be launched in the same direction the earth moves
around the sun. The extra fractional velocity (dV) then adds to the 29.73 km/s the space
craft has because of the earth’s heliocentric motion around the sun V cs. The space craft
finally speed around the sun (V1 = dV + Vcs) causes it to coast outward until it reaches the
outer planet. For small body orbiting another, very much large body (such as satellite
orbiting the earth) the total energy of the body is just the some of its kinetic energy and
potential energy, and its total energy also equals half the potential at the average distance
a (the semi major axis):
15
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Solving these equations for velocity results in the orbital energy conversion equation.
(
)
Where:
is the speed of orbiting body
is the standard gravitational parameter of the primary body
is the distance of orbiting body from the primary focus
is the semi major axis of the body’s orbit
Therefore the delta-v required can be computed as follows
√ (√
√ (
)
√
)
Where
and
are respectively the radii of the departure and arrival circular orbits, the
total delta-v in then
By Kepler’s third law, the time taken to transfer between the orbits is:
√
(
√
)
2.1.4. Timeline for Chemical Engine
These trajectories involve an extra burst of acceleration, administered en route. A typical
trip with liquid chemical propulsion takes one and a half years: 220 days getting there, 30
days on Mars and 290 days coming back. The return swoops toward the sun, perhaps
swinging by Venus, and approaches Earth from behind. The sequence can be flipped so
that the outbound leg is the longer one.
2.1.5. Trajectory for VASIMR [5]
The Earth-Mars heliocentric transfer, which takes 91 days and utilizes 36 mT of propellant,
Isp schedule in the range of 4,000 to 30,000 s, which delivers the specified payload with
minimum propellant in the required time. At Mars arrival, the relative velocity is 6.8 km/sec,
the DAV (61 mT) and empty propellant tanks (4 mT) are separated from the crew transfer
vehicle (CTV). The DAV descends directly to the surface, as the crew transfer vehicle
continues orbiting Mars (geostationary orbital parking) without a crew and rendezvous with
Mars after 2 days with 42 mT of propellant for return mission. The separation of the DAV
from the propulsion system at Mars arrival and its direct entry are operationally
reasonable. The delay in achieving orbital insertion of the propulsion module at Mars
results in considerable fuel and time savings.
16
Human Expedition on Mars Timeline 2018 by IUT Astronaut
In case of Mars-Earth heliocentric-transfer, duration is 174 days and propellant used in
our case is 42 mT as shown in Figure 2.1.5.
Figure 2.1.5: Heliocentric-Transfer orbit for VASIMR
2.2.
Landing Maneuver [2]
Descent Ascent vehicle (DAV) launched from KSC with Heavy Lift Launch Vehicle (HLLV)
and placed in LEO.A chemical Trans Mars Injection (TMI) is launched from KSC and
placed in LEO. It is then assembled with DAV. After vehicle verification, engines are
ignited and then TMI and DAV depart LEO for Mars. The DAV aero brakes descends and
lands on Mars. The CTV will be assembled in LEO. This will require several HLLV
launches. Components include, the VASIMR and its nuclear power plant, propellant, a
structural system for connection of all components to the DAV as shown in Figure 2.2.
Figure 2.2: Landing the Descent/Ascent Vehicle
The DAV is detached from CTV when the astronauts arrive in the crew transfer vehicle.
The DAV will aero brake and enter the lower Mars orbit after the astronauts aboard the
DAV, it descends much like the space shuttle, with its nose tilted upward. By rolling the
spacecraft to the left or right, the pilot can steer it toward the landing site. Parachutes slow
its descent, and then the retrorockets fire, enabling the pilot to set the craft down at exactly
17
Human Expedition on Mars Timeline 2018 by IUT Astronaut
the right spot. When the exploration in Mars will be finished the astronauts will then aboard
on DAV that blasts off the surface to an orbital rendezvous with the crew transfer vehicle,
which then brings the astronauts back to Earth.
2.3.
Time Schedule [1]
2.3.1. For Chemical Propulsion







1st Launch (Cargo Supply 1): 2nd October, 2017
2nd Launch (Cargo Supply 2): 17 November, 2017
3rd launch (Crew): 2nd January, 2018
Final launch From ISS: 4th January, 2018
Land on Mars: 11 August, 2018
Return Launching from Mars: 10 September, 2018
Return on Earth: 17 June, 2019
Figure 2.3: Mission overview.
2.3.2. For VASIMR Propulsion [1]
 1st Launch (Cargo Supply 1): 10 February, 2018
 2nd Launch (Cargo Supply 2): 27 March, 2018
 3rd launch (Crew): 10 May, 2018
 Final launch From ISS: 12 May, 2018
 Land on Mars: 11 August, 2018
 Return Launching from Mars: 13 August, 2018
 Return on Earth: 21 January, 2019
18
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Chapter 3
Mars Surface, Environment &
Pressurized Rover
3.1. Landing Site[6]
MGM2011 is a gravity field model for Mars that resolves features down to km-scales.
The model is constructed as a composite of a normal gravity field, a recent spacecollected gravity field and Newtonian gravity forward-modelling as shown in Figure 3.1.
 The normal gravity field approximates Mars's gravitational attraction as rotating
mass-ellipsoid,
 space-collected gravity (MRO110B2, Konopliv et al. 2011) delivers anomalies of the
Martian gravity field down to scales of ~125 km, and
 Newtonian forward-modeling and high-resolution Mars topography data from laser
altimetry (MOLA, Smith et al. 2001) is used to derive topography-implied gravity
(MRTM85) at spatial scales of ~125 km down to ~3km.

Figure 3.1: MGM2011 surface gravity accelerations over Mars' surface, unit is m s -2,
Mollweide projection centered to 0° longitude. Meridians and parallels are 30° apart.
So favorable landing places on Mars are the places colored with blue, sky blue or greenish
blue. Considering the surface condition the landing site should be a smooth one where
reduced level is almost same. Based on this aspect our proposed landing site is Quad 51
(nicknamed Yellowknife) of Aeolis Palus in Gale Crater. The landing site coordinates
are: 4.5895°S 137.4417°E. Gale crater, an estimated 3.5 to 3.8 billion-year-old impact
crater, is hypothesized to have first been gradually filled in by sediments; first waterdeposited, and then wind-deposited, possibly until it was completely covered.
Wind erosion then scoured out the sediments, leaving an isolated 5.5 km (3.4 mi) high
mountain, Aeolis Mons ("Mount Sharp"), at the center of the 154 km (96 mi) wide crater.
Thus, it is believed that the rover may have the opportunity to study two billion years of
Martian history in the sediments exposed in the mountain. Additionally, its landing site is
near an alluvial fan, which is hypothesized to be the result of a flow of ground water, either
19
Human Expedition on Mars Timeline 2018 by IUT Astronaut
before the deposition of the eroded sediments or else in relatively recent geologic history.
Figure 3.2: Inside planet mars
3.2
Mar’s Atmosphere
3.2.1 Chemical Composition
The atmosphere of Mars is about 100 times thinner than Earth’s, and it is 95 percent
carbon dioxide. Here’s a breakdown of its composition:
Also, minor amounts of: water, nitrogen oxide, neon, hydrogen-deuterium-oxygen, krypton
and xenon.
20
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Viking atmospheric measurements
Composition
Surface pressure
3.3
95.32%
2.7%
1.6%
0.13%
0.07%
0.03%
trace
carbon
nitrogen
argon
oxygen
carbon
water
neon,
krypton,
ozone, methane
1-9 millibars,
average 7 mb
depending
on
dioxide
monoxide
vapor
xenon,
altitude;
Pressurized Rover
Probably the most important activity for humans on mars is using the pressurized rover to
explore the surface. This rover enables the crew to explore up to 400 Km from the landing
point of DAV. This enables them to explore a potentially vast area of the Martian surface.
Design Features:
The rover will consist of following parts:
 Metal wheels
 Crew compartments
 Inflatable connectors
 Internal Combustion Engine
 Fuel Tanks
 Swivel collars
 Windows and Headlights
 Radiator fins
 Cranes
Metal Wheels
The rover will have metal wheels, since
the rocky surface of mars would be too
rough on tires. These wheels will have
special springs built into them to help
cushion the ride along with specially
designed suspension to swivel widely.
21
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Crew Compartment
The crew compartment will be specially protected from the stresses and strains of the
journey by a waffle-like shell around it. This protection is crucial, since any damage to the
shell of the rover could cause it to lose pressure. It should also be protected from
radiations and UV rays and carry food and medicine for the crews.
Inflatable Connectors
These connect the DAV to the rover.
Internal Combustion Engine
Most energy-efficient way of travelling across the Martian surface is internal combustion
engine. Rover will have two such engines, each of which burns methane (CH4) and
oxygen (O2). The rover will have an electrical transmission system. The engines will create
electricity that is transmitted to wheels by wires. An electric motor in each wheel then
propels the craft which is very resistant to breakdowns.
Fuel Tanks
Two fuels methane and oxygen are mixed in rover’s engine. Both of them are made from
the chemical reaction of CO2 with H2. CO2 will be extracted from atmosphere and O2 will
be brought from mars.
Swivel collar
These collars are directly attached to front and rear wheel assemblies. These collars
enable the wheel assemblies to rotate around the long axis of rover independently.
Radiator Fins
These fins help to control the temperature in the rover’s interior.
Crane
A multi-purpose folding crane is mounted on rails on the top of the rover. It will be used for
drilling and lifting purpose.
22
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Chapter 4
Engine and Fuel
4.1. Introduction
The history of man’s voyage to the space not brief. All the missions carried out till now has
been done using chemical propulsion engines. But certainly question arises during
considering the effectiveness of this engine in long term space travels, as the fuel
consumption of the engine is pretty bit high and the Isp value is low, which will result in a
longer voyage duration. Besides, longer exposure of the astronauts in the outer space
radiation is lethal. So, the question of finding a fast and fuel efficient engine certainly
comes in consideration. Recent development in the field of rocket science gives us many
options to choose [1]. Among them, we are considering the VASIMR Engine.
4.2. VASIMR Engine [1]
4.2.1. Composition
VASIMR stands for “Variable Specific Impulse Magneto plasma Rocket”. Actually, it is a
plasma rocket, which is a precursor to fusion propulsion. Currently under development the
leading advancement of a high power, electro thermal plasma rocket. Its design
incorporates low cost by utilizing hydrogen or inert gas propellant. The design also
provides high and variable specific impulse putting VASIMR at the forefront of any
propulsion system available today by NASA. It creates plasma under extremely hot
conditions and then expels that plasma to provide thrust. There are three basic cells in the
VASIMR engine, which are shown in Figure 4.2.1.
Forward cell - The propellant gas, typically hydrogen, is injected into this cell and ionized
to create plasma.
Figure 4.2.1: VASIMR structure
23
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Central cell - This cell acts as an amplifier to further heat the plasma with electromagnetic
energy. Radio waves are used to add energy to the plasma, similar to how a microwave
oven works.
Aft cell - A magnetic nozzle converts the energy of the plasma into velocity of the jet
exhaust. The magnetic field that is used to expel the plasma also protects the spacecraft
because it keeps the plasma from touching the shell of the spacecraft. Plasma would likely
destroy any material it came in contact with. The temperature of the plasma exiting the
nozzle is as hot as 180 million degrees Fahrenheit (100 million degrees Celsius). That's
25,000 times hotter than gases expelled from the space shuttle. The magnetic field ties all
the parts together and transmits the exhaust force that propels the ship.
4.2. Working principle [1]
The Variable Specific Impulse Magneto plasma Rocket bridges the gap between high- and
low-thrust systems. The propellant, generally hydrogen, is first ionized by radio waves and
then guided into a central chamber threaded with magnetic fields. There the particles spiral
around the magnetic-field lines with a certain natural frequency. By bombarding the
particles with radio waves of the same frequency, the system heats them to 10 million
degrees. A magnetic nozzle converts the spiraling motion into axial motion, producing
thrust. By regulating the manner of heating and adjusting a magnetic choke, the pilot can
control the exhaust rate .The mechanism is analogous to a car gearshift. Closing down the
choke puts the rocket into high gear: it reduces the number of particles exiting (hence the
thrust) but keeps their temperature high (hence the exhaust speed). Opening up
corresponds to low gear: high thrust but low efficiency.
Figure 4.2: VASIMR structure concept
4.3. Engine Subsystems [1]
Following subsystems are there in the total working process of the VASIMR engine.
1. Injection Stage (Helicon discharge)
A helicon is a low frequency electromagnetic wave. A helicon discharge can be defined as
an excitation of plasma by helicon waves induced through radio frequency heating. The
24
Human Expedition on Mars Timeline 2018 by IUT Astronaut
pressure of the magnetic field creates a helicon mode of operation with higher ionization
efficiency and greater electron density. VASIMR uses radio antenna to heat the plasma.
Two wave processes come into process. First, neutral gas in the injector stage becomes
dense and comparatively cold (60000 Kelvin) plasma through the action of helicon waves.
These are electromagnetic oscillations of frequencies of 10 to 50 MHz, which in a
magnetic field, energize free electrons in a gas. The electrons multiply rapidly by liberating
other electrons from nearby atoms in a cascade of ionization.
2. The Nozzle
While the cyclotron heating process results in approximately thermalized ion energy
distribution, the non-linear absorption of energy in the single-pass process produces a
boost, or displacement of the ion kinetic energy distribution. The ions are immediately
ejected through the magnetic nozzle before the ion distribution has had the time to
thermalize. A diverging magnetic field is used to convert plasma’s thermal energy into
direct kinetic energy.
3. Helicon Stage
The helicon stage is a stage where actual propulsion starts. It gives high temperature to
the gas passing in the two ends of the coil. The helicon is a helical antenna with width of
11 cm spread over the length of 16 cm.
Fig 4.3: Steady state Helicon discharge
A very high quality glass tube is inserted in the helicon antenna which can bear high
temperature. The glass tube in between is made of 4 cm in diameter as it gives the
possibility of the lowest operating temperature.
4.4
Nuclear reactor
A nuclear reactor capable of producing multi-megawatts of electric power is necessary to
allow the VASIMR engine to provide optimal thrust. It is suggested to employ a reactor
based on Battelle’s Rotating Multi-Megawatt Boiling Liquid Reactor (RMBLR). The
employs a fast energy spectrum, UN/molybdenum alloy cermet fuel reactor cooled by
boiling potassium. A direct Rankin cycle is used for energy conversion. The fuel is
fabricated in the form of blocks with coolant channels. Potassium flows through the reactor
in an inward radial flow to reduce thermal stress, leaving the reactor as vapor at a
temperature of 1440 K. With a bubble membrane radiator the specific mass at 20 MWe is
estimated to be 1-2 kg/kWe. The radial coolant inflow results in a reduced reactor vessel
operating temperature, which has potential benefits in operations and safety. This is a very
lightweight, advanced concept, and substantial development is required. Battelle
25
Human Expedition on Mars Timeline 2018 by IUT Astronaut
completed its conceptual design, as shown in Figure 4.4, but funding was cut before the
reactor could be completed and tested. With the proper funding, the RMBLR concept could
be completed with necessary reconfigurations, tested, and implemented within a fairly
short timeline.
Figure 4.4: Rotating Multi-Megawatt Boiling Liquid-Metal Reactor (RMBLR)
4.5. Fuel for VASIMR Engine [1]
After having so much research and calculations it is observed that these four propellants
mention bellow can be considered as better options to be used as VASIMR engine
propellant.
Propellant Properties
Argon(Ar)
Xenon(Xe)
Hydrogen(H)
Neon(Ne)
Atomic Weight
Atomic Volume (cm³/mole)
Density @293k (g/cm³)
State
Melting Point
Boiling Point
Specific heat capacity (J/gK)
39.948
22.4
0.001784
Gas
83.85
87.3
0.52
131.3
37.3
0.00588
Gas
161.3
165
0.158
1.0079
14.4
0.0000899
Gas
14.01
20.28
14.304
20.179
16.7
0.0009
Gas
24.53
27.1
0.904
Heat of vaporization (KJ/mol)
Heat of Fusion (KJ/mol)
1st Ionization energy (KJ/mol)
2nd Ionization Energy (KJ/mol)
6.447
1.888
1520.5
2665.8
12.636
2.297
1170.4
2046.4
0.904
0.117
1312
_
1.7326
0.3317
2080.6
3952.2
3rd Ionization Energy (KJ/mol)
Thermal Conductivity (W/mK)
Cost in $ (/100 g)
Ionization Energy/Cost (eV)
3930.8
0.0177
0.5
100
3097.2
0.00565
120
80
_
0.1805
12
200
6121.9
0.05
33
150
Figure 4.5: Fuel element comparison
Argon, Neon, Hydrogen, Xenon are the gases that can be used. Different properties and
parameters are shown in the table above. After seeing and considering all the possible
26
Human Expedition on Mars Timeline 2018 by IUT Astronaut
properties of all gases it seems that any of the above gas can be used. The major concern
for this kind of design process is cost and Weight. As it is shown in the table the cheapest
available gas is argon. Argon is about $40/kg vs. Xenon about $2000/Kg. So, Argon can
be considered as the best fuel option for the engine. But it should also be noted that
heavier gases like ammonia, nitrogen or water can be considered as possible candidate as
propellant too as they are denser and can produce denser plasma.
4.6. Human Mission to Mars with 12 MW Input Power [2]
The first studies of human missions to Mars, based on VASIMR® propulsion technology,
were conducted using HOT (Hybrid Optimization Technique) software. Those studies
demonstrated the capability of a 12 MW mission to transit to Mars within 3 months, which
is about twice as fast as the DRM (NASA Design Reference Mission) to Mars, assuming
chemical propulsion technology [Drake, 1998The Earth-Mars heliocentric transfer, which
takes 91 days and utilizes 36 mT of propellant, was calculated by Copernicus software
with an optimized, variable Isp schedule in the range of 4,000 to 30,000 s, which delivers
the specified payload with minimum propellant in the required time. At Mars arrival, the
relative velocity is 6.8 km/sec, the Mars Travel Vehicle (45 mT) and empty propellant tanks
(4 mT) are separated from the orbital transfer vehicle (OTV).After they reach the mars
surface safe and secured then the DAV(65 mT) is separated from the spacecraft along
with 2 crew members.
The mass budget for described mission according to our assumption is as follows:
 Mdepart = 250 mT
 E-M Propellant(ME-M =45mT)
 M-E Propellant(MM-E =42mT)
 MTV(MMTV =45mT)
 DAV (MDAV = 65 mT)
 Structure(Mstruc = 9mT)
 Tanks(Mtanks=5mT)
 Power propulsion corresponding 4Kg/KW specific mass
 Power Propulsion System(48mT) includes reactors(21mT),
thrusters (21mT),thruster radiators(2.4mT) and reactor radiators(3.6mT)
 Others: 1mT.
4.7. Recent Development of VASIMR Engine [1]
Recent development of VASIMR engine VX-200 in Ad Astra laboratory shows promising
aspects. In the data sheet shown below the latest performance of the engine using Argon
as fuel is showed. Along with that another table is given bellow showing the maximum
possible performance capacity of VX-200 using latest development:
Below a performance data of VX-200 engine is given using Argon propellant:
Power(KW)
200
Thrust(N)
5.7
VX-200
Exhaust Speed(Km/s)
50
ƞ
72%
27
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Table 4.7: Recent development chart
VASIMR thrust and power requirements
Burn time required
Input power for each plasma rocket
Variable Specific impulse for each plasma rocket
Initial mass of s/c before burn at insertion
Thrust of each VASIMR
Total no of VASIMR motor required
Orbital velocity of earth around sun
Total power required
Total thrust generated
Ideal exhaust velocity achieved by VASIMR
Total Propellant mass
Delta V
Values
39 days
200KW
4000-30000 sec
34.7 ton
5 Newton
45
29.78 km/s
9 MW
225 Newton
51.18 km/s
15462.24 kg
21.403 km/s
4.8. Observations and Recommendations
4.8.1 Observations
The advantages of the VASIMR are high power, high specific impulse, constant power
variable specific impulse, and potentially long lifetime due to magnetic confinement of the
plasma steam. The rocket relies on efficient plasma production in the first stage using a
helicon plasma source. Though the engine is still in the development stage it is possible to
make it ready within 2018 for full scale Mars mission. From the above discussion we can
come to the following observations:
1. The designed rocket geometry shows that proposed rocket is smaller than
conventional rockets.
2. The VASIMR rocket with the specific impulse of 5000 sec, force of 5 Newton per
engine, coupled power of 200kW and 72% thruster efficiency is capable of taking
34.7 ton weight to mars in 39 days.
3. Although it is very complicated and challenging to operate in high temperatures, the
VASIMR engine would be less expensive and could bring a revolution in space
missions. Some performance parameters of engine performance are calculated
based on available data and equations.
4. Flux density, Ion density, Ion velocity in different magnetic field and temperature
are the topics in early stage of research.
4.8.2 Recommendations:
1. Using heavier species like Nitrogen, Ammonia and water are possible candidates of
propellants as they are denser and produce dense plasma.
2. Efficiency of ICRF booster stage increases with the increase in plasma density.
Plasma density can be increased by additional gas input and increased helicon
stage power input.
3. Higher coupled voltage more than 200kW should be developed with more
improved thruster efficiency.
4. Mass flow rate should be increased in order to get more thrust force.
28
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Chapter 5
Electrical Power System
5.1. Introduction
The Electrical Power System of the spacecraft will be consists of the following divisions:
 Power Source
 Energy Storage
 Power Distribution
 Power Regulation and Control
5.2. Purpose of the Power System:





Supply a continuous source of electric power to spacecraft loads during mission life
Control and distribute electrical power to spacecraft.
Support power requirements for average and peak electrical power.
Provide converters for ac and regulated dc power buses.
Protect the spacecraft payload against failures within EPS.
5.3. Design Process [7]
Table 5.3: Electrical design block diagram.
5.4. Power Sources [9]
The Primary power sources are
 Primary Batteries
 Fuel Cells
 Radio Isotope Generators (RTG)
 Nuclear Reactor
 Solar Dynamics
 Solar Cells
29
Human Expedition on Mars Timeline 2018 by IUT Astronaut
5.4.1. Primary Batteries


Lithium sulphur dioxide and lithium thionyl chloride
Silver- Zinc
5.4.2 Fuel Cells
A fuel cell is a device that directly converts the chemical energy of reactants into low
voltage electricity, via electrochemical reaction. It is similar to the conventional battery.
Figure 5.4.2: Fuel Cell
Because of its high energy density when stored as a cryogenic liquid, hydrogen has
become the fuel of choice for aerospace application. The corresponding oxidant is liquid
oxygen. The reaction releases two free electrons and the waste product is water which
may be an advantage for manned missions in space.
5.4.3 RTG[7]
The design of RTG is simple by the standards of nuclear technology. The main component
is a sturdy container of a radioactive material. Radioactive decay of the fuel produces heat
which flows through the thermocouples to the heat sink, generating electricity in the
process. For the spaceflight use the fuel must produce a large amount of energy per mass
and volume. Isotopes MUST NOT produce significant amounts of gamma, neutron
radiation, or penetrating radiation in general though other decay modes or decay chain
products.
 In this paper it is proposed to use Pu238 in the RTG as it has the lowest shielding
requirements and longest half-life.
 It has a half-life of 87.7 years, losing 0.78% of their capacity per year.
 Reasonable energy density.
 Exceptionally low gamma and neutron radiation levels.
RTG material properties
Property
Po-210
Half-life,
0.378
years
Watts/gram 141
S/Watts
570
Pu-238
86.8
Ce-144
0.781
Sr-90
28.0
Cm-242
0.445
0.55
3000
0.93
250
0.93
250
120
495
30
Human Expedition on Mars Timeline 2018 by IUT Astronaut

It is suggested to use Stirling Radioisotope Generator (SRG) that uses free
piston stirling engines couples to linear alternators to convert heat to electricity. It
can be considered as alternative to RTG as it is still in research level under NASA
supervision.
Figure 5.4.3: SRG
5.4.4. Nuclear Reactor
The same reactor used to supply power to the VASIMR engine can be used to
generate electricity too.
5.4.5. Solar Dynamics
It’s a system that Provide higher efficiencies for solar power production. The advantages of
solar dynamic systems over solar photovoltaic is that dynamic systems
 Have higher thermal efficiency of about 30% whereas photovoltaic have that
of 3-4%.
 Can be used for higher power levels.
5.4.6. Solar Cell
Solar cells are semiconductors in which light of even relatively low energy, such as visible
photons, can kick electrons out of valance band and into the higher energy conduction
band creating electric current at a voltage related to the band gap energy.
5.4.7. Multiple Junction cells


Increase efficiency exploiting more spectrum.
Semiconductors connected in series with diode.
5.4.8. Solar Cells Technologies
Mono crystalline silicon cells are Well proven base technology in steady state performance
(around 30-50 W/Kg 13% EOL)
Multi-junction cells are Well proven based technology in constant evolution of performance
(23-30% targeted EOL). They have radiation resistance gain but with critical technologies.
5.5. Power Storage and Secondary Power Sources:


Secondary Batteries(accumulators)
Regenerative Fuel Cells
31
Human Expedition on Mars Timeline 2018 by IUT Astronaut
5.6. Power Distribution [8]
It consists of Cabling Fault protection and switch gear to turn on and off spacecraft loads,
Command decoders to command specific loads as shown in Figure 5.6. Power distribution
architecture depends on the dimension and complexity of spacecraft and demands in
terms of power. Two architectures can be considered
 Distributed: Each load has its own dedicated feeding and control system
 Centralized: Everything is controlled from central bus.
 Bus voltage required for the spacecraft would be 100-150V bus with a total power of
more than 2 KW.
Figure 5.6: Power distribution block diagram
5.7. Power Regulation and Control [8]
The regulation and control system can be formed by the following elements:
 Shunt Dump Module (SDM)
 Mode Control Unit (MCU)
 Battery Charge Regulator (BCR)
 Battery Discharge Regulator (BDR)
 Battery Management Unit (BMU)
 Power Conversion and Distribution Unit (PCDU)
32
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Chapter 6
Control System
6.1. Introduction
Control system of a Mars spacecraft that is designed for human exploration on the surface
of MARS is best described in some divisions. It is far more convenient to describe the
various aspects of the control rather than combining all of the concerned matters. The key
elements of the control system associate the interaction between various routine
operations and the analysis of information gathered by concerned operations. Owing to the
maintenance of cost efficiency and design parameters along with the issue of human
safety the various inspectional and communicative satellites and probes that are to be
used as follows.
Orbiter Satellite
Atmospheric Probe
Surface Rover
Landing Module
Surface Penetrator
Figure 6.1: Control System Components
All the above mentioned probes and satellites will maintain frequent and effective sharing
of data along with the analysis and sometimes some information about the consequent
routine or mitigating operation to be performed.
The control operations based on which phases they are to be executed are divided among
the Ground station (GS). Attending the issue of human of human safety, the GS would
collect and analyze all the data that are provided and give instructions for maneuvering.
The commonly referred ground station that is known as ‘Mission Control System’ [10]
consists of a computer system that connected to one or more ground stations which
provides the communication with the mother spacecraft. It has the privilege to dictate the
operation to be executed by the spacecraft control maneuvers and at some degree it
would provide the schedule along with necessary information for some routine work
making the maneuvering and propulsion of the spacecraft as smooth as possible.
33
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Mission Control
Sytem
Data collection
after Operation
Completion
Operational
Instruction Based
on data analysis
Data Sent from the
Spacecraft
Assembling and
Transfer of data to
station or stations
Figure 6.2: Overview of Mission control System actions.
The communication with the earth ground station or stations must be maintained on a
constant level without any interference or any kind of signal disruption symbolizing the
spacecraft has deviated from it’s trajectory. And should deviation occur the maneuvering
system would be designed as such to mitigate or counter such an event with an adequate
and subtle measure. Robust attitude control system is designed to counter error
occurrence. Along with control operation performed from the ground it is possible to design
more automation in control system.
6.2. Control System Architecture
There are three distinct phases through which the attitude control system must operate:



Transfer to ISS(TISS)
Trajectory acquisition(TA)
Mission Orbit(MO)
Transfer to ISS begins with the payload transfer procedure and during which the various
mandatory parts of the mission are rendezvoused together. Trajectory acquisition begins
with the starting from ISS and continues up to the transfer to the MARS orbital and its
order of occurrence is shown in figure 6.3.
Transfer to ISS(TISS)
Trajectory Acquisition(TA)
Mission Orbit(MO)
Figure 6.3: Order of the phases in which they occur.
34
Human Expedition on Mars Timeline 2018 by IUT Astronaut
6.3. Attitude and Articulation Control
The spacecraft's attitude, its orientation in space, must be stabilized and controlled so that
on board experiments may accomplish precise pointing for accurate collection and
subsequent interpretation of data, so that the heating and cooling effects of sunlight and
shadow may be used intelligently for thermal control, and so that propulsive maneuvers
may be executed in the right direction.

Spin Stabilization: Stabilization is accomplished by rotating the spacecraft mass,
thus using gyroscopic action as the stabilizing mechanism Spin-stabilized
spacecraft provide a continuous sweeping desirable for fields and particle
instruments, but they require complicated systems to de-spin antennas or optical
instruments which must be pointed at targets.[10]

Three-axis stabilization: Stabilization is accomplished by nudging a spacecraft
back and forth within a dead band of allowed attitude error, using small thrusters
and reaction wheels. Three-axis controlled spacecraft will point optical instruments
without having to de-spin them.

Reaction wheels: Electrically-powered wheels mounted in three orthogonal axes
aboard a spacecraft. To rotate the vehicle in one direction, you spin up the proper
wheel in the opposite direction. To rotate the vehicle back, you slow down the
wheel. Excess momentum that builds up in the system due to external torques must
be occasionally removed from the system via propulsive maneuvers.
6.4. Attitude and articulation control subsystem
The onboard computer that manages the tasks involved in spacecraft stabilization via its
interface equipment. For attitude reference, star trackers, star scanners, solar trackers,
sun sensors, and planetary limb trackers are used. Gyroscopes are carried for attitude
reference for those periods when celestial references are not being used. The AACS also
controls the articulation of the spacecraft's moveable appendages such as solar panels
and optical instrument scan platforms.[10]
 Pitch Control: The basis of pitch control operation using RWA is a tachometer loop
that maintains the speed of the reaction wheel. It is desirable to keep the speed at a
nominal level usually ±10%, to keep the spacecraft gyroscopically stiff.
The simple control scheme used is to multiply the difference between the desired speed
and measured speed by a gain and to filter the measured speed by a first order filter.
The resulting closed loop system is [11]
(
)(
)
(
)
(
)
Where
is called the filter cutoff.
was chosen to provide adequate damping and
made sufficiently large to provide good disturbance rejection and command tracking.

is
Yaw Control: The roll yaw control system must accomplish two requirements. First,
it must attenuate the external disturbances on the spacecraft. The second
requirement is nutation damping. Since there is no passive source of nutation
damping the control system must damp the nutation. The first part of the controller
takes the low frequency approximation to the closed loop system and selects a pair
of gains to meet the pointing requirements. This approach, a purely proportional
35
Human Expedition on Mars Timeline 2018 by IUT Astronaut
control, is the simplest. The nominal plant of an earth pointing momentum bias
spacecraft has poles at orbit rate and at the nutation frequency.
The low-frequency approximations of the roll yaw equations are [11]
And the torque command is
[ ]
[
]
Since only roll is measured.
A good approximation of spin-stabilized roll pitch and RWA is given in the below
figure.
Figure 6.4: Diagram showing positioning of pitch axis and reaction wheel axis
including Yaw
pitch
Sensors: The sensors that are to be used in the attitude control algorithm input are –
 Horizon Sensor Assembly(HSA)
 Sun Sensor Assembly(SSA)
 Earth Sensor Assembly(ESA)
 Mars orbiter Sensor Assembly(MOSA)
 Magnetic Sensors Assembly(MGSA)
 MARS Surface landers and
 Atmospheric probe
 Gyros
 Reaction Wheel Assembly(RWA) Tachometer
36
Human Expedition on Mars Timeline 2018 by IUT Astronaut
The attitude control logarithm employed is closed loop control method concerning the
human safety issue. A block diagram showing the actuators and sensor output is shown in
the below figure. The error computation are performed using these processed output of the
sensors and then these are outputs are delivered to the control system phases. Then the
outputs of the control system are distributed to the actuators such as rocket engine
thrusters (REA), electric hydrazine thrusters (EHT) shown on the right hand side.[11]
The control system architecture is shown below with a simple diagram
Control
Output
Sun
Horizon
Earth
Sensor
Inputs
Reaction
Wheel
Assembly
Pulsed
Plasma
Thruster
Attitude
Control
Algorithm
MARS
Orbiter
Gyros
Electric
Hydrazine
Thruster
Spin
Stabilization
Assembly
Figure 6.6: Simple Expression of Control Architecture
6.5. Thruster Operations [11]
During thruster operations the 3-axis torque commands generated by the controllers are
fed into the simplex algorithm, along with the positions and thrust vectors of the available
thrusters. The simplex linear programming algorithm is used to determine the optimal set
of pulse width commands. Optimal is defined as using the minimum amount of fuel
necessary to produce the requested torque. The use of simplex allows the operator to
account for thruster misalignments, plume disturbances and center-of-mass motion by
changing ground loadable Parameters without any reprogramming. The simplex
implementation is customized for this application making it efficient enough for use with a
relatively slow computer. The torque distribution law automatically limits the number of
iterations in simplex and tests each pulse width command for validity. All of the thruster
control loops use thruster pulse width modulation [13]. The minimum pulse width is relatively
large which can lead to limit cycling if the disturbances are small (which is particularly true
during non-station-keeping operation). The control system allows the operator to choose a
pulsing period that is longer than the control period. For example, the station-keeping
loops run at 2 Hz but a typical pulsing period will be 8 seconds, meaning that thrusters will
only fire once every 16 control cycles. This reduces limit cycling significantly.
37
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Chapter 7
Space Communication
7.1. Introduction
Space communication for a round trip from Earth to Mars become so important and
complicated as it needs to ensure Quality of service (Qos) over 400,000,000 km ( distance
between earth & mars ) where round trip for light costs around 44 minute.
7.2. Description
The total communication consists 3 steps.
1. Near Earth Communication
2. Deep Space communication
3. Deploying dedicated network assets.
Figure 7.1: Basic Communication system
The Near earth communication is maintained by the S-band and Ka-band links, the service
called by Tracking and Data Relay Satellite System (TDRSS) as shown in Figure 7.1. The
S-band is a part of microwave band and electromagnetic spectrum. It is defined by an
IEEE standard for radio waves with frequencies that range from 2 to 4 GHz [14]. The Kaband links are able to interact with ground station with 1-3 Gbps speed from the low-Earth
orbits. From the launching through Trans Mars Injection (TMI) the communication system
follows TDRSS based system [15].
38
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Figure 7.2: Architecture of deep Space communication Network [16]
Figure 7.2 describe the backbone and planetary network where backbone network deals
with Earth, other planets, the moon, satellites and relay satellites and planetary network
deals with planetary satellite network and planetary surface network.
With the X-band and Ka-band links, DSN gets much higher speed, still round –trip time
(RTT) costs around 8.5 min to 40 min depends on orbital location of the planets and bit
error rates are on order 10-1. On the mars, a high speed connection between user and
spacecraft is needed to cope up with the landing and exploration of mars. For a safe
landing, choosing better position for landing, dedicated telecommunication or other survey
a relay satellite would be very helpful. For an energy efficient means for communication
between a Mars user and earth this relay satellite could also play a key role in supporting
communications between spatially separated users at Mars. In short a dedicated relay
spacecraft orbited Based on assessments to date must be planted for inter planetary
communication.
The whole mission at the environment of mars is very precious that it would start transfer
the mission dealings live to the earth by DSCN. A big part of the exploration is based on
the picture taken from the environment of the mars. In deep space communication system,
the storage and transmission of image data holds a large part of the band-width. So in
order to satisfy the requirement for bandwidth and storage capacity, high efficient image
compressing coding method is needed.
39
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Chapter 8
Human Health Risks & Solutions
8.1. Introduction
In spaceflight there are various risks on human health. This can be divided into two
aspects.
 Physical
 Psychological
8.2. Physical Threats
8.2.1. Effect of G. Force [17]
One of the basic problems that an astronaut faces is effect of G force. A lack of training
may cause blackout with blood circulation being impaired which has relation with tunnel
vision problem. Due to G force the problems that are caused is compacted bladders, RBC
burst, subdural hematomas cessation of circulation. And it is very obvious in this case that
an astronaut has to face this effect.

Probable solution: To overcome this problem, alongside proper training the
seating arrangement can be made relative to direction of acceleration and deacceleration. This will reduce the effect to some extent.
8.2.2. Space adaptation sickness [17]
In the first two days of flight, a sickness is caused generally defined as “Space adaptation
sickness” which includes disorientation, pallor, malaise loss of motivation, irritability,
drowsiness, stomach awareness, infrequent but sudden vomiting.

Probable solution: This is not a major problem. Astronaut can adapted them with
the condition within two days. As well as a proper training will help them to adapt
with the condition easily.
8.2.3. Fluid Shift [17]
Engaging into new atmosphere also causes fluid shift which means flow of blood towards
head which remains for 5 days causes blood plasma to decrease by 12% and body water
decrease by 23%. Study shows that SAS experiencing is less in smaller spacecraft,
women are less affected by SAS and suffer less.

Probable solution: Astronaut can adapted them with this condition within five days.
8.2.4. Effect on Muscle [17]
In a long flight muscles of body are affected into high negative impact causing atrophy
less resistant fatigue, uncontrolled muscle Twitches, loss of fine motor control, weakened
ligaments, tearing water and loosing strength. Care study of following two instances will
give a prospective outlook of muscle loss.Berezovoy & Lebedev stayed for 211 days at Mir
40
Human Expedition on Mars Timeline 2018 by IUT Astronaut
and at returning to planet couldn’t walk in that time. Shannon Lucid’s example contains
staying in space for 188 days and could walk a very little.

Probable solution: Study shows that at any causes the muscle of man lost by
25%. Proper muscle exercise and expert advice may help to get the muscle man to
normal level through proper guidance, diet and exercise.
8.2.5. Immune system changes [18]
In flight an astronaut goes through Immune System Change causing weakness, rapid loss
of plasma volume, reduction of RBC, activation of viruses, increase of infection.
8.2.6. Microbes
exchange between astronauts: Due to microgravity
sedimentation of microbes does not take place. So the separation of microbes is
problematic. Therefore in most cases the microbes are airborne.

Probable solution: One of the solutions can be to spread bacterial probes
throughout the spacecraft. Spacecraft sterilization is necessary. And the astronauts
should be quarantined for 1 week prior to them which will reduce the amount of
microbes. Nutritional health is very important for an astronaut, as it helps to broader
state to defend the primary problems and secondary problems in outer space. As
now 70 foods and 20 beverages are allowed in space so, it is highly recommended
to take calcium strengthen food and food with higher protein, Vitamin D, K and fat.
Whilst staying in mars the main problem is real –time communication as the
message delivery time between earth and mars is several minutes, so at
emergency situation the astronaut should be trained to deal with the problem and
take suitable decisions.
8.2.7. Calcium loss
This is a major problem for a spaceflight like this. Calcium loss is very site specific. The
effect is severe in those spots where there is load muscle and load bone. Due to
microgravity the pressure on load bone and muscle decreases, resulting calcium loss.
The sky lab IV study of the calcium level of the astronauts in 84 days:
1) Calcium loss through urination is up to 30 days.
2) Calcium loss through stool is up to 84 days.
The total bone mass is lost 1-1.5% per month as calcium loss results in bone mass loss
and fracture. In long terms mission “Osteoporosis” is inevitable with 20% or more of bone
mass to be lost.
From this statistics we get that this is not at all negligible. There is higher probability of
suffering from kidney stone.

Probable solution: The vital problem can be overcome
by hard training of
resistive exercise eating vitamin D and K tablets, calcium tablets. There may be
cycling and other exercise which will give pressure to load bone and muscle. But
this will not help so much for this kind of long trip. So there is a suggestion of
creating artificial gravity.
41
Human Expedition on Mars Timeline 2018 by IUT Astronaut
o An artificial gravity is practically impossible till now. Theoretically centrifugal
force can be used to create artificial gravity.
o There can be a chamber known as gravity chamber which will rotate around its
own axis producing gravity. So crew can experience gravity at certain period of
time and can perform exercise.
8.2.8. Gas exposure
CO2, other leaked gases from reservoir are exposed to the astronauts.

Probable solution: In normal condition, lithium hydroxide filter traps CO2 and
activated carbon filter trap other gases.
8.2.9. Biofilm [19]
Mixture of bacteria and fungi creates biofilm. It has the power to oxidize copper cable
resulting damage to electronic devices.

Probable solution: The electronic devices should be coated with plastic to avoid
oxidation.
8.3. Radiation challenge [19]
Radiation problem is the most hazardous problem which keeps us in the dilemma that
whatever “Mars Expedition” is possible or not due to space radiation cataracts conforms in
eye, malignant tumors are formed, genetic code is altered(infertility & sterility), birth
defects are happened.
As for present statuesque for dealing with space radiation for “short flight”, astronaut suits
are designed, the spaceship is designed with compounds which includes aluminum to
defend radiation. But when it comes to “long distance flight”, it’s seen that maximum of the
radiation can’t be blocked. As “space radiation blocking for long flights” is still not sure, so,
some hypothetical solutions can be proposed. The spaceship could be designed with the
compounds which will block the total radiation. Aluminum is more preferable in this regard.
The space suit can also be “doubly alloyed” for radiation protection. As radiation itself is a
form of energy, it can be transformed to heat energy and passed away to space through
exhaustion tunnel. Also the tunnel radiation can be transformed to “Usable source of
energy” at space craft. A hypothetical model can be proposed, where the radiating
particles can react to the particles kept as a coating outside the spacecraft forming
complex compound and thus can be disposed to space. This will sound little childish but
as we know “venom cuts venom”, so, a coating of artificial lab made radiating particles can
cut out the space radiation. The real earth model can be proposed for the spacecraft. The
radiation towards the earth is protected by a magnetic field. The same procedure may be
applied to the spacecraft to block the radiation by creating a suitable magnetic field. As the
astronaut is exposed to various hazards its must to keep in healthy and sound so proper
entertainment system including movies, family pictures, and food should be available to
dissect the stress and make the flight a memorable one.
42
Human Expedition on Mars Timeline 2018 by IUT Astronaut
8.4. Psychological aspects [19]
As a matter of fact the estimated time of flight towards mars is 3 months approx. So a long
term flight brings change in psychological pattern which makes the astronaut more difficult
to take decisions in any unexpected circumstances as negative reaction is inevitable in
small flight. So, the astronauts of same culture is encouraged and different sex is
acceptable through wedlock.
There are some other facts that affect psychology of an astronaut:
8.4.1. Sleep loss [18]
Lack of sleep takes an astronaut to fatigue. This will turn him/her to misjudge. It will reduce
the efficiency of the astronauts.
8.4.2. Circadian irregularity [18]
Our body follows a certain rhythm. The temperature of the body rises and falls, metabolism
rate increases with time. Alertness decreases at night. Again alertness increases at
morning. But it may changes if we change our sleep time. Irregular circadian rhythm can
lead to various physical and psychological problems.
To avoid this astronauts should shift their works. There should be a fixed sleep time (they
must take the sleep). From statistics on psycholonotor vigilance task known as PVT
(alertness and effect of fatigue on cognitive performance) we get that a 9hrs sleep gives a
good PVT.
8.4.3. Privacy:
A study says that astronauts are always under camera, so they are psychologically
affected. They require some private time. So there should be one room where they can
breathe freely.
8.4.4. Recreational activity [18]
There should be some recreational activity depending on the astronauts. It will give some
mental happiness to them. Again to meet physical and psychological need it will be a
better choice to select crews who are well acquainted to each other.
43
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Conclusion
The whole plan has been based on the technology available now and which have strong
probability to be developed within the year 2018. From the above mentioned mission
proposal it is evident that Manned mission to mars will now be more achievable with
minimal risks of human life. The article clearly distinguished the fact of using VASIMR
engine as a feasible way in terms of cost efficiency and time efficacy. The trajectories are
carefully designed for timely traversal to mars. Sophisticated calculations are done to
ensure safe landing and launching through DAV. The control system is designed in a
practical way to maintain the best maneuver of control architecture. Various space
communication modules are implemented to ensure the command and control of the CTV
and DAV throughout the mission with the GS (Ground Station). Human health in physical
and mental perspective is emphasized with maximum caution and various preventive
measures are discussed. Radiation being the most alarming factor is discussed thoroughly
and suggestions are proposed to overcome this unwanted hazard. At length the proposal
is envisioned to be the best as it discussed all the conundrum to mars and the mission to
mars will now be more acquirable than before.
Before successful manned mission to Moon, everybody used to criticize the possibility of
conducting a safe round trip to the Moon. But technology of ‘60s has done the impossible.
Since then, technology has taken a huge leap of development.
Mars has always been the topic of discussion among the astrophysicists. Ever since the
possibility of presence of water in Martian surface has been declared, the possibility of
existence of life in mars has also increased. A manned mission can reveal the secrets of
the red planet. Hopefully, this small effort of ours can help to set up the milestone in the
history of space science.
44
Human Expedition on Mars Timeline 2018 by IUT Astronaut
Acknowledgement
For our project we would like to sincerely thank Dr. Khondokar Habibul Kabir (Asst.
Professor, Dept. of EEE, IUT) for his guidance, understanding, patience, and most
importantly, his friendly attitude during our work. We remain grateful to him for all the
reviews he made to our paper and making it a successful one.
We would also like to thank Dr. Zubrin known as an aerospace engineer and author, best
known for his advocacy of the manned exploration of Mars. We followed some of his ideas
in “Design Reference Mission” which is adopted by NASA. This mostly helped us to build
the foundation of our paper.
We are also pretty much grateful to the NASA and Mars Society for providing us with lots
of information about Mars exploration, launching-landing of spaceships, Orbital
maneuvering, trajectories, Mars environment and information of all the missions that has
become succeeded so far, communication systems from Mars etc. Apart from that giving
an overall idea for manned exploration of Mars from earth.
Our heartiest gratitude goes to Mr. Mark Carter of Ad Astra Company for providing us lots
of papers and a lot of information about VASIMR engine.
We are also indebted to Mohammed Atiquzzaman, PhD (Professor, School of Computer
Science, University of Oklahoma) for his suggestions and help in the field of space
communication.
At last but not the least we are very grateful to our beloved institution, Islamic University of
Technology (IUT) for providing us with technical support and all other necessary
resources.
45
Human Expedition on Mars Timeline 2018 by IUT Astronaut
References:
[1]
[2]
[3]
[4]
[5]
[6]
[7]
[8]
[9]
[10]
[11]
[12]
[13]
[14]
[15]
[16]
[17]
[18]
[19]
Mitesh D Patel, “ VASIMR Engine desine for a mission to mars”(2011).
Andrew V. Ilin, Leonard D. Cassady, Tim W. Glover, Franklin R. Chang Diaz,
“VASIMR Human mission to mars”(2011).
George Musser and Mark Alpert, “HOW TO GO TO MARS”(2000).
“Architecture Study Report Summary”, Inspiration MARS.
Damon F. Landau and James M. Longuski, “Trajectories for Human Missions to
Mars, Part 1: Impulsive Transfers ”(2006).
“Mars Gravity Model 2011 (MGM2011)”, Curtin University and wikipidia.org.
Politechnico de Milano, “ Electric Power System(EPS)”(2010).
David W. Miller and John Keesee, “Spacecraft power system”(2011)
Virginia Tech,” Spacecraft power system”(2012)
A. Baldi, M. Jones, J.F. Kaufeler and
P.Maigne, “ The Evolution of ESA's
Spacecraft Control Systems”, Flight Control Systems Department, European Space
Operations Centre (ESOC), Darmstadt, Germany.(2012)
Wendy I. Sullivan, Michael A Paluszek Princeton Satellite System, Inc. and Walter
K. Daniel, “ A New Satellite Attitude Control System”, Princeton Satellite System,
Inc. and CTA Space System, Inc.
Yew-Wen Liang, Tzu-Chiang Chu and Chiz-Chung Cheng, “ Robust Attitude Control
for Spacecraft”, Department of Electrical and Control Engineering, National Chiao
Tung University, Hsinchu 30010, Taiwan, Republic of China.(2012)
“Basic of Space Flight II”, NASA JPL.
S_band, Available: http://en.wikipedia.org/wiki/S_band
Xiaoyou Yu, Fang Yu, Weibing Hou, Xiaochun Wang, “State-of-the-Art of
Transmission Protocols for Deep Space Communication Networks”(2010)
Xiao Song, Li Yunsong, Bai Baoming, ZhouYouxi, “The Key Technologies of Deep
Space Communications”.(2011)
M. Collen Gino, “ Human Spaceflight”(2012).
Francis A. Cucinotta,” Uncertainty Analysis in Space Radiation” (2011).
N.A. Schwadron, H.E Spance, L. Townsend, C.Zeittlin, D.Fry, W. Farrel, “
Understanding and Predicting the Space radiation Hazard- a Critical element of
human exploration beyond low earth orbit”.(2011)
46
Human Expedition on Mars Timeline 2018 by IUT Astronaut

Download Report