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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

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Human Expedition on Mars Timeline 2018 by IUT Astronaut

Team Members

1. MD. Wasif Zaman (Team Leader)

2. Fahad Al Mamun

3. Faisal Hussain

4. M. Saiful Bari

5. MD. Abdullah Al Joha

6. MD. Muhtady Muhaisin

7. Mohammad Abdullah Matin Khan

8. Muhammad Usama Islam

9. Raihan Uddin Ahmed

10. Shabab Bin Karim

11. 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)

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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.

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Contents

Preface 06

Chapter 1 Payload Transfer Maneuver 1.1 Introduction 07 1.2 Description of Payload Transfer Maneuver 07

1.2.1 External Tank 07

1.2.2 Solid Rocket Booster 08

1.2.3 The Orbiter 09

1.3 Earth to LEO Maneuver 10

1.4 Orbital Maneuvering System 11

1.5 Reaction Control System 11

1.6 Electrical Power Distribution System 11

Chapter 2 Launching and Landing Maneuver with In Space Trajectory

2.1 Launching Maneuver 13

2.2 Landing Maneuver 17

2.3 Time Schedule 18

Chapter 3 Mars Surface, Environment & Pressurized Rover

3.1 Landing Site 19

3.2 Mars Atmosphere 20

3.3 Pressurized Rover 21

Chapter 4 Engine and Fuel

4.1 Introduction 23

4.2 VASIMR Engine 23

4.3 Engine Subsystem 24

4.4 Nuclear reactor 26

4.5 Fuel for VASIMR Engine 26

4.6 Human Mission to Mars with 12 MW Input Power 27

4.7 Recent Development of VASIMR Engine 27

4.8 Observations and Recommendations 28

Chapter 5 Electrical Power System

5.1 Introduction 29

5.2 Purpose of the Power System 29

5.3 Design Process 29

5.4 Power Sources 29

5.5 Power Storage and Secondary Power Sources 31

5.6 Power Distribution 32

5.7 Power Regulation and Control 32

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Chapter 6 Control System

6.1 Introduction 33

6.2 Control System Architecture 34

6.3 Attitude and Articulation Control 35

6.4 Attitude and articulation control subsystem 35

6.5 Thruster Operations 37

Chapter 7 Space Communication

7.1 Introduction 38

7.2 Description 38

Chapter 8 Human Health Risks & Solutions

8.1 Introduction 40

8.2 Physical Threats 40

8.3 Radiation challenge 42

8.4 Psychological aspects 43

Conclusion 44

Acknowledgement 45

References 46

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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.

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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 inter-tank area inert, and an electrical system to distribute power and instrumentation signals and provide lightning protection as shown in Figure 1.1.

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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

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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

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(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 T-7 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 -9.00.0 Arm Solid Rocket Boosters -5.00.0 Auto Sequence Start -0.31.0 Main Engine Start -0.06.0 SRB Ignition 0.00.0 Liftoff 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

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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

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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.

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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.

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SLS have launching capability twice a year with proper funding and fuel supply which is shown in Figure 2.1.1.

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

SLS 130-metric-ton

Evolved Configuration

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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 Vcs. 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):

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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.

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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

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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

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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 space-collected 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 water-deposited, 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

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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.

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3.3 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.

Viking atmospheric measurements

Composition

95.32% 2.7% 1.6% 0.13% 0.07% 0.03% trace

carbon dioxide nitrogen argon oxygen carbon monoxide water vapor neon, krypton, xenon, ozone, methane

Surface pressure 1-9 millibars, depending on altitude; average 7 mb

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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.

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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

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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

25


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

26


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 39.948 131.3 1.0079 20.179

Atomic Volume (cm³/mole) 22.4 37.3 14.4 16.7

Density @293k (g/cm³) 0.001784 0.00588 0.0000899 0.0009

State Gas Gas Gas Gas

Melting Point 83.85 161.3 14.01 24.53

Boiling Point 87.3 165 20.28 27.1

Specific heat capacity (J/gK) 0.52 0.158 14.304 0.904

Heat of vaporization (KJ/mol) 6.447 12.636 0.904 1.7326

Heat of Fusion (KJ/mol) 1.888 2.297 0.117 0.3317

1st Ionization energy (KJ/mol) 1520.5 1170.4 1312 2080.6

2nd Ionization Energy (KJ/mol) 2665.8 2046.4 _ 3952.2

3rd Ionization Energy (KJ/mol) 3930.8 3097.2 _ 6121.9

Thermal Conductivity (W/mK) 0.0177 0.00565 0.1805 0.05

Cost in $ (/100 g) 0.5 120 12 33

Ionization Energy/Cost (eV) 100 80 200 150

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

27


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:

VX-200

Power(KW) Thrust(N) Exhaust Speed(Km/s) ƞ

200 5.7 50 72%

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Table 4.7: Recent development chart

VASIMR thrust and power requirements Values

Burn time required 39 days

Input power for each plasma rocket 200KW

Variable Specific impulse for each plasma rocket 4000-30000 sec

Initial mass of s/c before burn at insertion 34.7 ton

Thrust of each VASIMR 5 Newton

Total no of VASIMR motor required 45

Orbital velocity of earth around sun 29.78 km/s

Total power required 9 MW

Total thrust generated 225 Newton

Ideal exhaust velocity achieved by VASIMR 51.18 km/s

Total Propellant mass 15462.24 kg

Delta V 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.

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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

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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 Pu-238 Ce-144 Sr-90 Cm-242

Half-life, years

0.378 86.8 0.781 28.0 0.445

Watts/gram 141 0.55 0.93 0.93 120

S/Watts 570 3000 250 250 495

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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

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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)

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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.

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.

Orbiter Satellite Atmospheric Probe Landing Module

Surface Rover Surface Penetrator

34


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.

Figure 6.3: Order of the phases in which they occur.

Mission Control Sytem

Data Sent from the Spacecraft

Assembling and Transfer of data to station or stations

Operational Instruction Based on data analysis

Data collection after Operation

Completion

Transfer to ISS(TISS)

Trajectory Acquisition(TA)

Mission Orbit(MO)

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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 is made sufficiently large to provide good disturbance rejection and command tracking.

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

36


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

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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

Sun

Horizon

Earth

MARS

Orbiter

Gyros

Attitude

Control

Algorithm

Sensor

Inputs

Reaction

Wheel

Assembly

Pulsed

Plasma

Thruster

Electric

Hydrazine

Thruster

Spin

Stabilization

Assembly

Control

Output

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.

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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 Ka-band 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].

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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.

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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 de-acceleration. 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

41


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.

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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.

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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.

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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.

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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.

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“VASIMR Human mission to mars”(2011). [3] George Musser and Mark Alpert, “HOW TO GO TO MARS”(2000). [4] “Architecture Study Report Summary”, Inspiration MARS. [5] Damon F. Landau and James M. Longuski, “Trajectories for Human Missions to

Mars, Part 1: Impulsive Transfers ”(2006). [6] “Mars Gravity Model 2011 (MGM2011)”, Curtin University and wikipidia.org. [7] Politechnico de Milano, “ Electric Power System(EPS)”(2010). [8] David W. Miller and John Keesee, “Spacecraft power system”(2011) [9] Virginia Tech,” Spacecraft power system”(2012) [10] 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)

[11] 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.

[12] 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)

[13] “Basic of Space Flight II”, NASA JPL. [14] S_band, Available: http://en.wikipedia.org/wiki/S_band [15] Xiaoyou Yu, Fang Yu, Weibing Hou, Xiaochun Wang, “State-of-the-Art of

Transmission Protocols for Deep Space Communication Networks”(2010) [16] Xiao Song, Li Yunsong, Bai Baoming, ZhouYouxi, “The Key Technologies of Deep

Space Communications”.(2011) [17] M. Collen Gino, “ Human Spaceflight”(2012). [18] Francis A. Cucinotta,” Uncertainty Analysis in Space Radiation” (2011). [19] 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)

http://en.wikipedia.org/wiki/S_band

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