Opportunities From Using The Variable Specific Impulse Magnetoplasma Rocket For Manned Missions To Mars

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MASTER'S THESIS

Opportunities From Using The Variable Specific Impulse Magnetoplasma Rocket

For Manned Missions To Mars

Vladimir Gurevich

Master of Science (120 credits) Space Engineering - Space Master

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CRANFIELD UNIVERSITY

Vladimir Gurevich

OPPORTUNITIES FROM USING THE VARIABLE SPECIFIC IMPULSE MAGNETOPLASMA ROCKET FOR MANNED MISSIONS

TO MARS

School of Engineering

Astronautics & Space Engineering (SpaceMaster)

MSc

Academic Year: 2011 - 2012

Supervisors Dr. Jennifer Kingston Professor Bob Parkinson

May 2012

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CRANFIELD UNIVERSITY

School of Engineering

Astronautics & Space Engineering (SpaceMaster)

MSc

Academic Year 2011 - 2012

Vladimir Gurevich

OPPORTUNITIES FROM USING THE VARIABLE SPECIFIC IMPULSE MAGNETOPLASMA ROCKET FOR MANNED MISSIONS

TO MARS

Supervisors Dr. Jennifer Kingston Professor Bob Parkinson

May 2012

This thesis is submitted in partial fulfilment of the requirements for the degree of MSc in Astronautics and Space Engineering.

© Cranfield University 2012. All rights reserved. No part of this

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ABSTRACT

Human interplanetary flights have been widely considered as an unconditional priority towards the development of the astronautics for several decades now.

Nevertheless, at the moment there are no certainly available technologies that can safely and effectively bring humans back to the Moon and conquer more distant objects in the solar system. This thesis is dedicated to the case of future expeditions to Mars. An extensive list of problems and critical considerations for this kind of missions includes:

long time life condition maintaining; space radiation protection; microgravity effects minimization; capability of abort options and failure risk minimization. There are a number of concepts and approaches for Human Mars mission design. Each of them under favorable conditions might allow overcoming the existing barriers. The nuclear electric propulsion concept and more specifically a VASIMR engine are examined in details in order to estimate their real potential. The attractive from many points of view VASIMR engine proposition was evaluated as currently unrealistic due to several serious technical challenges including a requirement in developing a new space nuclear generator with high power-to-weight ratio.

Keywords:

Interplanetary flights, Human mission to Mars, VASIMR engine.

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ACKNOWLEDGEMENTS

I would like to express my deepest gratitude to everyone without whom this project would not be possible. First of all I want to thank my kind and responsive supervisors – Doctor Jennifer Kingston and Professor Bob Parkinson – for their goodwill, guidance and support. Also I need to mention specialists outside Cranfield who have found time to answer my questions via e-mail: ir. Marc Naeije (Delft University, Netherlands), Mr. Andrew Ilin (AARC, United States), Mr. Pedro Molina Cabrera (Bionik Labs, Canada), Dr. Donald Rapp (United States).

My sincere appreciation is extended to the SpaceMaster Consortium for giving me this unique opportunity to study on this prestigious program together with such talented and wonderful young professionals from all over the world. Personal thanks to my bravest colleague Mr. Victor Huarcaya Azañon for million ideas and an inspiring example of how a thirst for space discoveries can overcome all life challenges including cancer.

I would like to address the last but not the least words of gratitude to my caring parents. Thank you for your infinite love, support and sometimes inexplicable faith during the project and during my whole life.

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LIST OF FIGURES

Figure 1-1. Basic questions of the project. ... 2

Figure 2-1. Cumulative human spaceflight time allocation for different nations in hours. ... 6

Figure 2-2. The Earth-Mars Hohmann transfer [9]. ... 10

Figure 2-3. Earth-Mars distance [11]. ... 11

Figure 2-4. General Thrust-Isp correlation for various propulsion concepts. ... 12

Figure 2-5. Selective classification of various propulsion concepts. ... 13

Figure 2-6. Operational principle of the VASIMR engine concept [14]. ... 15

Figure 2-7. The spacecraft configuration of the PLT concept [16]. ... 16

Figure 3-1. Mars mission architecture in its phases for which alternative options can be implied. ... 19

Figure 3-2. Architecture for the DRM-3 (chemically propelled) mission (adapted from [17, 18]). ... 20

Figure 3-3. Delta-v calculations for the simplified Earth-Mars orbit [17]. ... 25

Figure 3-4. Trajectory for the short stay mission [24]. ... 27

Figure 3-5. Trajectory for the long stay mission [24]... 28

Figure 3-6. The SLS concept art interpretation [28]. ... 30

Figure 3-7. Schematic of Mars Aerocapture [17]. ... 34

Figure 4-1. Daily regenerative life support requirements [6]. ... 37

Figure 5-1 Relationship between available Solar Energy Flux ... 49

Figure 5-2. Comparison of Mars mission trip time for nuclear and solar powered designs [39]. ... 50

Figure 5-3. Schematic of nuclear Brayton CCMHD space power system [45]. ... 54

Figure 6-1. A simplified VASIMR working cycle. ... 56

Figure 6-2. Fuel consuption estimations for different flight mods (based on [50]). ... 58

Figure 6-3. The VASIMR propelled spacecraft art interpretation [52]. ... 60

Figure 7-1. Allocation of five Lagrangian points in a two-body system (Earth-Moon system). ... 71

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Figure 7-4. The dimensions of the habitat module (based on [59]). ... 74

Figure 7-5. Trajectory optimisation variable parameters [62]. ... 80

Figure 7-6. Results of Multi-objective evolutionary algorithm simulations of an Earth- Mars mission [62]. ... 81

Figure 7-7. Helio-transfer trajectory optimisation (adapted from [39]). ... 83

Figure 7-8. Specific impulse profile optimisation (adapted from [39]). ... 83

Figure A-1. The Martian surface [64]. ... 98

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LIST OF TABLES

Table 2-1. Human spaceflights chronology. ... 5

Table 3-1. Chronology of some of the notable human Mars mission propositions (based on [20, 21]). ... 21

Table 3-2. Initial values for Earth and Mars ... 24

Table 3-3. Overview of the short stay mission options [23]. ... 26

Table 3-4. Overview of the long stay mission options [23]. ... 28

Table 3-5. Different energy types of trips characteristic [17]. ... 31

Table 4-1. ECLSS estimations for a conservative human Mars mission [29]. ... 38

Table 4-2. ECLSS estimations amendmented for using water-air recycling [29]. ... 38

Table 4-3 Recommended Dose Limits for Astronauts [1]. ... 40

Table 4-4. Physiological effects of microgravity. ... 44

Table 5-1 Assumed mission parameters [39]. ... 50

Table 5-2 Space radioisotope thermoelectric generators designs [40]. ... 51

Table 5-3 Comparison of characteristics of space nuclear fission reactors [41]. ... 52

Table 5-4. Details of CCMHD nuclear space power system [45]. ... 53

Table 6-1. Comparision of gases that could be used as a propellant ... 59

Table 7-1. Mars mission summary. ... 67

Table 7-2. Propulsion system assumptions [39]. ... 69

Table 7-3. Habitat module mass breakdown ... 75

Table 7-4. Power consumption of the spacecraft subsystems. ... 77

Table 7-5. Estimated Earth-Mars mass budget [based on 39]. ... 78

Table 7-6. Estimated return trip mass budget [based on 39]. ... 78

Table 7-7. Parametric results for roundtrip missions for VASIMR engine models with different levels of power supply [39] ... 82

Table 8-1. Approaches for a Human mission to Mars. ... 87

Table A-1. Martian orbital characteristics [65]. ... 99

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LIST OF ABBREVIATIONS

AARC Ad Astra Rocket Company ACS Attitude Control System

AU Astronomic Units

CCMHD Closed Cycle Magnetohydrodynamic CTV Crew Transfer Vehicle

DRM Design Reference Mission

ECLSS Environmental Control and Life Support System EDL Entry, Descent and Landing

EM Electromagnetic

EMI Electromagnetic Interference ERV Earth Return Vehicle

FRC Field Reversed Configuration GCR Galactic Cosmic Ray

GPHS General Purpose Heat Source

HR High-Reflection

HTGR High Temperature Gas Reactor HTLI High-Thrust-Low-Impulse

ICRH Ion Cyclotron Resonance Heating IMLEO Initial Mass in Low Earth Orbit ISRU In-Situ Resource Utilization ISS International Space Station LEO Low Earth Orbit

LMO Low Mars Orbit

LTHI Low-Thrust-High-Isp MAV Mars Ascent Vehicle MHD Magnetohydrodynamic

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MMRTG Multi-Mission Radioisotope Thermoelectric Generator NASA National Aeronautics and Space Administration PLT Photonic Laser Thruster

PMD Power Management and Distribution PPU Power Processing Unit

RF Radio Frequency

RTG Radioisotope Thermoelectric Generator SEP Solar Energy Particle

SLS Space Launch System

SNAP Systems Nuclear Auxiliary Power TCS Thermal Control System

TEI Trans Earth Injection TMI Trans Mars Injection TPI Trans Planet Injection

US United States

VASIMR Variable Specific Impulse Magnetoplasma Rocket

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1 INTRODUCTION

1.1 Topicality

40 years ago, when people on Earth had got used to the idea that surface of Moon is marked by human being traces, there was no question whether the next height – people walking on Mars – would be conquered. The question was “when?” Years have passed. Human society has done incredible progress. Technological evolution has overcome bounds of imagination. But Mars is still out of reach and there is still no certain answer to the question “when?” In fact the most optimistic predictions expect a manned flight to Mars to become possible not earlier than in 20 years – this is exactly the same expectations people have made 20 years ago [1], which leads us to unfavourable conclusions.

If we want we can search for an explanation in diversified scope of political and economic reasons, conspiracy theories, but nevertheless the truth is that the interplanetary manned missions are (i) extremely dangerous, (ii) technically complicated, (iii) associated with gigantic costs. And although someone can argue that there are no insuperable technical obstacles [2], manned interplanetary travel in fact is the most challenging direction of astronautics in the conceivable future. That is why in order to risk a life of at least a well-trained space crew, in order to bring together the brightest brains of generation and in order to eventually spend colossal amount of tax money, we need a really good reason for necessity of this journey beside putting a colourful flag and taking notable pictures.

Why our goal is to send people to Mars at the first place? What we cannot achieve with relatively cheap, unpretentious and immortal robots? Most of the reasons that can be found in both American and non-American recent sources still can be considered as inconclusive echoes of the cold war. During the unpleasant realities of the global absurd competition it was important to be the first in literally everything by any cost: first in the ice hockey, first in Hanoi, first in the orbit. Fortunately the morbid craving for self-affirmation and political prejudices do not rule the world anymore and the reasons like “we must go to Mars because we are the greatest nation ever lived”

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The real motivation in sending astronauts to other planets is the same as the motivation the great explorers had have. Why Ferdinand Magellan, Leif Ericson and Christopher Columbus started their dangerous and tangled expeditions? Doubtfully they did it in order to proof superiority of their races or political systems. They did it to create the world we know, love, and basically live in today. Astronautics is essentially the same with sailing explorations, but involves additional dimension. Our goal can be formulated as a search for life, colonisation, and scientific spinoffs – but in fact the real reason is in changing world by extending human life area and borders of our imagination.

1.2 Identification of key challenges

The main nerve of the project is concentrated in a wide investigation of possibilities to overcome the interplanetary barrier for manned missions. A very simplified motivational description of the project can be shown in figure below.

Figure 1-1. Basic questions of the project.

The central objectives can be formulated as following:

 Comprehensive study of aspects of the human mission

 Identification of key challenges

 Evaluation and discussion of possible solution

 Trade-off of the feasible propulsion approaches for the mission

 Assessment of the potential space nuclear reactors

 Depth analysis of a selected prospective concept (VASIMR)

 Based on assumptions system design of a conditional mission with a VASIMR engine and a nuclear space generator

 Getting an answer to the question about the possibilities of an expedition to Mars in the near future

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The project objective statement:

Within 5 months’ time perform a technical investigation of manned interplanetary mission challenges; comprehend whether the variable specific impulse magnetoplasma rocket can match and improve the mission requirements; give a reasonable prediction for future of interplanetary space expeditions.

1.3 Thesis layout

This report covers a wide range of topics and discussions. The interdisciplinary specific of the project has affected the thesis structure and coherent link between sections is rather complicated. This is very important to spend some time and clarify logical portrait of the entire report.

At the beginning (Section 2) the brief overview of manned space flight objectives and mechanisms for interplanetary missions is given. The most attractive propulsion concepts are introduced. This is followed by the detailed analysis of objectives and challenges we face by designing space missions to Mars (Section 3) and human deep space journeys in general (Section 4). Guided by conclusions that for such missions the nuclear energy is the only effective power source the detailed overview of development of nuclear space generators is given (Section 5). The emphasis is made to the concept that most likely would be able to provide sufficient amount of energy in the nearest future (CCMHD). The Variable Specific Impulse Magnetoplasma Rocket concept (claimed as a central topic of the research) is analysed in details in order to identify its advantages and disadvantages, as well as to examine its suitability for human Mars flights (Section 6). The climax of the project is the mission design (Section 7) preceded by assumptions discussion. The mission design consists of mission scenarios trade-off, subsystems integration and mission optimisation. At the conclusion (Section 8) obtained results are summarized and further work is suggested.

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2 FUNDAMENTALS OF INTERPLANETARY MANNED MISSIONS

We are starting immersion in one of the most complex subjects in astronautcs.

An essential part of the research is in understanding the basics of (i) specificity of human spaceflights and of (ii) astrodynamics for interplanetary missions. The section consists of formulation of objectives (2.1); a brief historical background (2.2);

complexity overview associated with such systems (2.3); mechanisms of interplanetary spaceflights (2.4); rocket concepts overview (2.5). In more details these fundamentals will be discussed in section 3 (astrodynamics and propulsion points of view) and in section 4 (human spaceflight aspects).

2.1 Main objectives of performing interplanetary space mission

In order to design a successful mission there is need to understand all (even the most obvious) objectives and key performance steps of the mission. The objectives below can be described as the second level. By going deeper in details these objectives can be splitted in a more specific and broad set of objectives of the levels 3 and 4.

 Choosing a destination

 Formulating scientific value of the mission

 Formulating mission requirements

 Designing the mission

 Financial cover of the mission

 Crew selection

 The mission preparation

 Launch and Earth departure

 Fast and safe journey to the target planet orbit

 The planet entry, descent and landing

 Productive surface stay

 Returning the crew safely on Earth 2.2 History of human spaceflights

The space exploration was the area of technology development perhaps the most affected by the cold war. The Soviet Union and the US have used the peaceful by definition space arena for their villainous political interests: ideological propaganda and military deterrence. Despite all negative and unethical factors this ruthless competition (known as the Space Race) has dramatically accelerated space exploration in the Sixties

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and Seventies. This has proved that astronautics is one of the rare areas of human life that could be developed much slower in presence of free trading than by direct government initiatives.

In particular the motivation for the manned spaceflights were especially affected by the political conjuncture. This was more important to send a man in space before a competitor than to send him at all. The most important milestones of the human spaceflights are presented in Table 2-1.

Table 2-1. Human spaceflights chronology.

Accomplishment Date Country

First man in space 12.04.1961 First woman in space 16.06.1963

First spacewalk 08.03.1965

First Moonwalk 21.07.1969

First manned orbital station 19.04.1971 First reusable orbital spacecraft 12.04.1981

First Marswalk ?

For the moment only 3 countries have experience with manned space missions:

Russia (as the only successor of the Soviet space program), the United States and the People's Republic of China. However, more than 40 countries have in total send their citizens to space aboard Soviet, American, Russian, and Chinese spacecraft. The total number of manned space launches up to the December 2011 is 284; the number of individual astronauts is 525 [3]. There were 4 accidents resulting in crew members death, claimed the lives of 18 astronauts and cosmonauts. Human spaceflight statistics can be summarized in Figure 2-1, showing the cumulative time allocation of entire 951849 hours (22 November 2011) for different nations [3].

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Figure 2-1. Cumulative human spaceflight time allocation for different nations in hours.

2.3 Difficulties of manned interplanetary missions

In the recent scientific publications arguments against using human spaceflights can be found more often than before [4]. In the age of effective artificial intelligence the reasons for sending poorly adaptable to difficult conditions humans seem especially unpractical. Since our objectives attentively involve the presence of a crew, instead of trading-off between real and virtual astronauts it is more useful to discuss problems caused by the presence of a man on-board.

2.3.1 Life support

The necessity of supporting life conditions and providing satisfactory consumables is crucial not only for human spaceflights but for any space mission that involves living organisms starting from the very first space traveller – the dog Laika on board of Sputnik-2 in 1957. The first space life-support system designed to keep Laika alive for a week consisted of a small oxygen generator, devices for avoiding oxygen poisoning and for absorbing carbon dioxide, a simple air-conditioning system, a week food supply, a waste collector, and chains that in fact had prevented all dog’s movement [5]. Because of the system imperfection Laika has died during first 5-7 hours after launch due to overheating and emotional stress. This example of cruelty demonstrates among other things the importance of the proper life support.

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When we talk about human flights to distances beyond the LEO we talk about supporting life conditions for the crew during long time: up to several years.

Throughout this time there are requirements in: maintaining a breathable atmosphere through air revitalization; keeping a comfortable temperature-humidity environment;

providing water and food supply; supporting a complex contaminant control; organizing waste management. The environmental control and life support system (ECLSS) should be designed and configurated in a way which would allow an optimal collaboration of all these processes.

2.3.2 Radiation

The harmful for humans and equipment radiation environment of outer space is one of key constraints for extensive crewed missions. The three main types of radiation hazards in space are Van Allen radiation, Solar Energetic Particle events (SEPs) and Galactic Cosmic Rays (GCRs). Beside the external space radiation there might be sources of hazard radiation right on the spacecraft. For all missions that involve usage of nuclear elements for generating either power or thrust, there is an additional need in protecting astronauts from radiation exposure caused by on-board nuclear plant. Space radiation effects and the possible solutions of shielding against it are discussed in the section 4.2.

2.3.3 Low gravity deconditioning

Human space flights imply placement of astronauts in conditions free from gravitational force coverage. In fact all crewed missions are seriously complicated by weightless effects. This is epitomized in impossibility for normal and convenient work and vital functionality of the crew.

Even more serious problem associated with long-duration micro-gravity experience is physiological detriment. Zero G environment induces a range of physiological impairments, bone demineralization, cardiovascular deconditioning along with changes in fluid distribution within the body, skeletal muscle atrophy [6]. The ways of preventing these effects are in minimizing transfer time; intensive physical

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2.3.4 Reliability

Even a minor fault during the flight potentially leads to a very serious risk of fatal failure for the whole mission and therefore for the crew. Complex missions (such as interplanetary manned flights) require equally complicated and explicably expensive back-up plans: variable mission abort options, multiple vehicles (“lifeboat”) [7].

2.3.5 Human factors

Since manned interplanetary missions are in paramount of difficulty, the role of a human (as the main and only advantage to unmanned missions) is to be implicated in the complex mission control circuit in the optimal way. This requires years of training as well as designing of advanced control loops.

2.3.6 Communication lag

Tolerable for unmanned long distance missions delays in communication link between spacecraft and ground stations bring complexity in manned flights. However, presence of ‘human intelligence’ on board has to redistribute duties between the ground and the crew. This is one of the keys to compensate the absence of immediate communication link.

2.3.7 Physiological and psychological aspects of the mission

The manned mission to Mars is conjugated with unique and until now never unattainable task of sending human beings to another planet. The total mission time is estimated from several months up to few years. The distance of round trip is over hundreds of million kilometres. There was no opportunity to directly study the hazards space environment effects on humans before. All these emotional factors lead to the stupendous psychological pressure on the astronauts. Even with the most realistic mission simulations on Earth, the crew of the first Mars ship will be the first to experience the whole range of physiological effects.

Besides already discussed issues caused by ionizing radiation and low-gravity environment the HUMEX study formulates the following effects of a long-duration interplanetary mission [8]:

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 Physical effects of a prolonged low-light environment

 Psychological effects of isolation from earth

 Psychological effects of lack of community due to lack of real-time connections with earth

 Social effects of several humans living under crowded conditions for over one earth year

 Inaccessibility of terrestrial medical facilities 2.4 General principles of interplanetary missions

The first goal of any interplanetary mission is to escape the gravity well of the planet of origin. After that depending on if the destination object is closer or further away from the Sun the spacecraft has to increase or decrease its speed, due to the difference in speeds of rotation of the two objects (planets) around the Sun, and enter a heliocentric orbit that intersects the destination orbit. Eventually the spacecraft brakes into an orbit around the destination planet. For chemical propelled missions these manoeuvres are generally achieved by two powered phases – an Earth departure burn and an arrival burn at the destination planet [7].

The main requirement in sense of necessary fuel mass (and consequently available mass for carried payload) for reaching another planet is delta-v (change in speed), required to match velocity with another planet. The Tsiolkovsky rocket equation relates delta-v with specific impulse I_sp (effective exhaust velocity) and stage mass fraction (initial mass Mo divided by final mass Mfin which is basically initial mass minus used propellant Mf) is presented in the equation (2-1).

(

) (2-1)

The minimum delta-v requirements for a transfer between two coplanar circular orbits characterize the simple Hohmann transfer, where changes in velocity are made only at aphelion and perihelion [7]. The principle is based on the fact that a tangent orbit between original and target is the lowest energy route for transfer from one orbit to another. The first application of thrust moves spacecraft from the original circular orbit to an elliptical transfer orbit. During the whole heliocentric transfer the main engine is

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Figure 2-2. The Earth-Mars Hohmann transfer [9].

The Hohmann transfer covers an angle of π radians and consequently it takes a half of the orbital period of the outer orbit to accomplish the transfer. That is why in terms of planetary relative positions there is a specific requirement that at the point when the spacecraft reaches the destination orbit there is 180 degree between the launch point and the target (figure 2.2). Time for the Earth-Mars transit using the Hohmann is 8.5 months [10]. For a back-and-forth mission the amount time will exceed to 32 months due to rather a long surface stay caused by necessity to wait for a proper interposition of the two planets.

But even for not Hohmann trajectories we need to take into account cyclicity of the relative position of the planets. Mars orbit is 1.524 times greater than Earth’s orbit around the Sun. Because of this Mars rotates around the Sun slower than Earth and the planets’ positional arrangement relative to each other changes with time considerably.

Because of this the Earth-Mars distance is relevant for mission operations and communications. The period of the Earth-Mars distance variation cycle is 26 months in average and it is demonstrated in Figure 2-3.

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Figure 2-3. Earth-Mars distance [11].

The two-way trip time can be obtained from the equation below [10]:

( )

(2-2)

W is an integer number which represents the type of the transfer. For a fast round trip with a dual-Hohmann transfer W=1 [10]. are angular velocities of Earth and Mars.

2.5 Rocket concepts for interplanetary missions

2.5.1 Propulsion systems overview

Propulsion is the heart of spacecraft, a driving force for rapid explorations of depths of space. One of the reasons for a renunciation of the manned Moon flights and for persistent postponing of Mars missions is in inexpediency of the classical chemical rockets. The ultimate goal of propulsion engineers is development of a new powerful thruster that could combine generally mutually exclusive characteristics of high specific

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Figure 2-4. General Thrust-Isp correlation for various propulsion concepts.

Classical high thrust mission approach for interplanetary flights (impulsive trajectories) uses short time burns in desired points in order to provide energy for starting the mission (gravitational escape, heliocentric transfer) and to capture the destination planet’s orbit. The most commonly used high-thrust propulsion systems (in fact the most commonly used systems among all types of propulsion) are chemical. For the majority of such systems the principle is in chemical reaction between a fuel and an oxidant, which releases the energy (exhaust). The limitation for chemical propulsions is in decomposition or combustion of molecular compounds. Thermal rockets generate reactive thrust by heating (nuclear or solar heating) and subsequent expelling of the propellant. The biggest problem of thermal rocket concept is enormous total mass of the system.

Low thrust missions are not capable of providing short and efficient burns like chemical rockets and thereby require a higher specific impulse to offset the less efficient propellant utilization. High specific impulse in turn provides a more efficient propellant consumption, but leads to inflated power requirements. The thrusters operating on electric power have a better predisposition for fast deep space journeys. However, their

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performance is considerably limited by power source capabilities and heavy plant- propulsion systems.

The trade-off between the space propulsion concepts (Figure 2-5) is based on (i) feasibility, (ii) propulsion characteristics, (iii) simplicity, (iv) ecological factors. It is worth saying that beside these propulsion systems there are other solutions like aerobraking that basically also provides delta-v and can be considered as a propulsion technique (see 3.3).

Figure 2-5. Selective classification of various propulsion concepts.

The feasible candidates for long and heavy missions like manned interplanetary flights are Hall-effect ion propulsion, Photonic Laser Propulsion, Field Reversed Configuration Thruster and VASIMR [12, 13]. Introduction for these concepts is presented below. Choice in favour of VASIMR for our investigation is primary not because of its impressive technical advantages but due to the big attention around the VASIMR development among the scientific and general public and our willingness to answer a question about its real feasibility.

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2.5.2 Hall-effect thruster

Hall-effect ion thruster is one of the oldest electrical propulsion systems that have been in development since the 1960s. The principle of work is based on the drifting of moving ions in a crossed electric and magnetic fields (the Hall Effect) along with accelerating of these ions (Faraday Motor Effect) [7]. The propellant (most commonly it is Xenon) is ionized by electrons trapped in magnetic field. Thereupon the resultant ions are accelerated to high speed which results in producing of thrust. Later the ions are neutralized in the plume.

Propulsion characteristics of this thruster are rather moderate: Thrust around 3N and Specific Impulse 80 km/s for Hall-effect thruster in kW power range. Since it is not probable that in near future there will be a perceptible increase in these values the ion thrusters are unlikely to be considered as the main propulsion concept for human space travel. However its temperate size and demands make ion thruster a good candidate to use it as an auxiliary system, for example for the attitude control.

2.5.3 VASIMR engine

The Variable Specific Impulse Magnetoplasma Rocket is currently a widely advertised electromagnetic propulsion concept. Relatively inexpensive gas (like Hydrogen) is used for creating plasma by helicon antenna ionization, which then accelerated by an ion cyclotron resonance frequency booster with variable intensity. The working propellant rectilinear movement is supported by superconducting magnets. The expanding magnetic field at the magnetic nozzle converts the thermal motion of hot plasma into directed flow.

The key advantage of the VASIMR is the implication of principle of changing specific impulse in different phases of the flight accordingly to thrust/propellant/delta-v optimization requirements. The VASIMR concept allows using advantages of high thrust for overcoming gravitational fields in turn with high specific impulse used for the most of the journey routine. This combination might lead to more efficient fuel consuming and most importantly to significant helio-transfer time reduction in comparison to constant specific impulse engines.

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Figure 2-6. Operational principle of the VASIMR engine concept [14].

In order to use appreciable benefits of this concept for human interplanetary flights (short journey time, effective fuel expense) the VASIMR engine requires hundreds of MW of power and currently exorbitant specific mass of less than 1 kg/kW.

Taking into account that in the nearest future the problem of satisfying these demands most probably will not be solved there is a number of sceptical opinions about expediency of VASIMR [15].

2.5.4 Photonic Laser Thruster

Photonic propulsion is an interesting and promising concept for futuristic space missions. The PLT requires at least two spacecraft furnished with large high-reflection (HR) mirrors. One of the spacecraft is equipped with a gigantic photonic laser that is aimed towards the mirror of the second spacecraft which contains the mission payload.

This spacecraft can be designed lightweight since the mirror is the only main propulsion component of the layout. After a series of reflections the laser beam creates an outward force on both spacecraft. The thrust of the photons is amplified by an active resonant optical cavity. The possible specific impulse is estimated up to incredibly high values (300000 km/s) [16]. However, the medium power lasers not likely to give high levels of thrust to the spacecraft and there will be need in additional thrust amplification by using the reflectant mirrors as cavity traps for the beamed photons.

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Figure 2-7. The spacecraft configuration of the PLT concept [16].

The main limitation of this concept is the range in which the laser can thrust. A maximum range for feasible mirror size and laser characteristics is 400000 km [12]

which is extremely short in the scale of interplanetary distances. Another difficulty is a necessity in extremely precise and stable attitude control system. Thereby the PLT is not a momentary solution for planning a deep space mission but rather a propulsion system of future.

2.5.5 Field Reversed Configuration Thruster

A FRC thruster is an electromagnetic propulsion system with its origin in nuclear fusion technology and benefits from effective plasma magnetic confinement.

The term Field Reversed Configuration is basically a special state of plasma. The name Reversed is adopted because the FRC magnetic field is produced in opposite direction of the external field.

The formation of FRC plasma is happening in three important steps: (i) the propellant ionization by radio frequency waves (at the same fashion with VASIMR), (ii) using rotating magnetic field to create the plasmoid, (iii) the confinement of the created plasmoid in equilibrium and free of contact with the walls [12]. Afterwards the FRC plasmoid is accelerated through the field reversing magnetic coils.

The potential of this concept is hundreds km/s of specific impulse with low power supply [17] and thousands with high power supply. The main advantages are simplicity and opportunity to couple propulsion system with nuclear fusion power generator. The problems are heavy weight and need in extremely high power. At the same time this concept does not seem to have such essential drawbacks as VASIMR

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and PLT have. In case of serious development progress in fusion technique the FRC can become a key propulsion for future space missions.

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3 MISSION ARCHITECTURE

The understanding of the mission architecture recounted in this chapter is based on the initial strategy and fundamental considerations aimed to answer a question “How do we reach Mars?” It should be not confused with a more profound mission design given in section 7, which uses results of this mission architecture analysis as well as the mission crucial elements overview (Chapter 4) as starting points for portraying the mission elements and scenario.

The section starts with an introduction to the mission architecture in case of human Mars flight. Main principles are illustrated in the example of the architecture for the Design Reference Mission (3.1). The basis for the mission architecture design is the expense estimations that are used as an indication factor for the trade-off between mission options (3.2). The key options of reducing the large initial mass on the Earth orbit and therefore the total mission cost are introduced (3.3). Trajectory analysis and transfer duration for classical approach are examined in order to get an appreciation for the complex problems faced by the mission (3.4). The section also contains introduction and description of the key mission architecture components: launch to Earth orbit (3.5), Trans Mars and Trans Earth injections (3.6) and orbit insertion (3.7).

3.1 Mars mission architecture overview

Any logical study of a space mission starts with initial planning and architecture options trade-off. The idea is to find a compromise solution that meets scientific demands with technological and financial means. The main problem of analysing futuristic missions like human interplanetary flights is necessity to take into account a great number of assumptions and suppositions. The majority of proposed architectures for this moment fully rely on the technologies that are not flight proved. Some of these concepts exist only on paper.

Alternative options for different Mars mission architectures can be specified as in Figure 3-1. Here is important to mention that many of practical architectures (majority of high-thrust concepts) do not consider any alternatives for the interplanetary transfer assuming a trip on empty tanks without producing thrust. With this architecture

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the spacecraft reaches the destination planet using its own momentum. This approach, however, has some advantages pointed in 2.5.

Figure 3-1. Mars mission architecture in its phases for which alternative options can be implied.

One of architecture concept examples – the Design Reference Mission 3.0 (DRM-3) proposed by NASA Johnson Space Center Mars in 1997 [18] – is presented in Figure 3-2. Here the crew, the cargo and the Earth Return Vehicle (ERV) are considered separately in order to optimize masses on destination. Crew pursues Mars orbit insertion by high energy path and after Entry, Descent and Landing process (EDL) 30 t module reaches the Mars surface. Cargo, estimated to be around 43 t on destination, includes Mars Ascent Vehicle (MAV) for transporting astronauts from the Mars surface to the departure orbit. According to this concept the MAV uses for fuel the In Situ Resource Utilization (ISRU) products from the Mars surface. The ERV travels to Mars orbit by low energy trajectory where it rendezvous with MAV and continues on high energy path to the Earth direction with crew on board. The mass of ERV at Mars is 29 t.

The DRM-3 was based on a series of improbable assumptions about the ISRU potential, the life-support system capabilities and the Habitat module mass [17]. The total destination mass of 73 t is rather understated and it will be not sufficient to perform a full surface mission and an ascent manoeuvre. Nonetheless this example illustrates one of the common approaches to the architecture of this specific mission.

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Figure 3-2. Architecture for the DRM-3 (chemically propelled) mission (adapted from [17, 18]).

3.2 Expense estimations

In order to estimate possible expenses it is common to use the Initial Mass in Low Earth Orbit (IMLEO) as a relative index of mission costs. The IMLEO is a direct indicator of requirements for a launch vehicle and for a number of launches with following orbit rendezvouses and assembles. Each of architectures involves a serial of sets leading from departure to return. The mission can be divided by conditional milestones:

LAUNCH PAD → LOW EARTH ORBIT (LEO) ↔ MARS ORBIT ↔ MARS SURFACE ↓

EARTH

Estimation of IMLEO therefore is done by the following steps [19]:

1. Estimated the payload masses of vehicles on every state (backwards):

M₁ - mass transported to Mars surface and returned to Earth M2 - mass transported to Mars orbit and returned to Earth

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M4 - mass to Mars orbit

2. Estimate the mass ratios for every step using models for delta-v

G1 - mass ratio for journey from Earth to the Mars surface and back to Earth G2 - mass ratio for journey from Earth to the Mars orbit and back to Earth G3 - mass ratio for journey to Mars surface

G₄ - mass ratio for journey to Mars orbit 3. Calculate IMLEO:

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The estimated IMLEO value dictates the scale of the planned mission and is a significant proportional cost index at the initial stage of the budget evaluation.

Comparison of IMLEOs for different Mars mission proposals is presents in Table 3-1.

These values demonstrate a gradual evolution of proposed missions from the early unrealistic hypotheses about a prompt Mars colonization (Wernher von Braun) to more down-to-earth Nuclear thermal proposals.

Table 3-1. Chronology of some of the notable human Mars mission propositions (adapted from [20, 21]).

Year Proposal Name Propulsion IMLEO (t)

1952 Von Braun Mars Expedition Nitric acid / Hydrazine 37200 1956 Von Braun / Willy Ley Mars

Expedition Nitric acid / Hydrazine 3400

1956 Soviet Martian Piloted

Complex Liquid oxygen / Kerosene 1360

1957 Ernst Stuhlinger Mars Study Nuclear electric 660 1960 Lewis Research Centre Mars

Expedition Nuclear thermal 614

1971 NASA Mars expedition Liquid Oxygen / Liquid Hydrogen 1900 1982 Mars via Solar Sail (UK) Solar Sail + Aerocapture 900

1986 Energia Mars-1986 Nuclear electric 365

1991 Mars Direct Nuclear thermal //

Liquid Oxygen / Liquid Hydrogen 220 // 220 1996 Design Reference Mission 3 Nuclear thermal 410 1998 DLR + Alenia Mars study Nuclear thermal //

Liquid Oxygen / Liquid Hydrogen 420 // 545 One of the crucial concerns about a human Mars mission is enormous fuel mass requirement which affects the entire scenario starting from the launch. That is why the

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Since the fuel mass inter alia depends on the spacecraft mass and can dramatically vary during the mission design procedure, the IMLEO therefore has a variable value and is determined by a series of iterations. The essential mass ratios for different chemical propellants and types of transfer are given in the Appendix B.

3.3 Ways of decreasing IMLEO

It has been mentioned that the Mars mission is practically unrealizable because of the combination of adverse factors including very large IMLEO. There are several propositions of reducing the IMLEO and consequently increasing the mission feasibility [7]:

 Nuclear thermal propulsion

 Nuclear electric propulsion

 Solar electric propulsion

 Aerobraking at Mars

 In Situ Resource Utilization on Mars

However, none of these options is currently available to integrate into the mission architecture in the way that it could undoubtedly save an appreciable portion of mass. From the other hand there is need in only one of them in order to bring the IMLEO down to the realistic ground around 200 t [7]. This is difficult to name an obvious favourite over these 5 options since all of them have advantages and disadvantages. Dependent from technical development progress any of these technologies can become an enabler for spaceflight beyond LEO.

3.3.1 Nuclear thermal propulsion

Nuclear thermal rockets might have potentially a very impressive specific impulse and thrust-to-weight ratio (10km/s and ~3-10 consequently [7]). But a huge nuclear plant traveling around the solar system and theoretical danger of its sudden explosion brings a lot of public opposition to such proposals.

3.3.2 Nuclear electric propulsion

Theenvironment considerations are also true for the variety of electric propulsion systems with high power demands that require a nuclear source on board.

The VASIMR as one of such concepts has several significant advantages in terms of

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fuel consumption and reusability, but technical drawbacks (weight of total power- propulsion system, high temperature levels, environmental issues, uncertainty with pace of energy generators development) do not allow us to confidently give preference to the Nuclear electric propulsions either. The requirement of the high power-to-weight is extremely unrealistic at the moment (see chapter 5).

3.3.3 Solar electric propulsion

Using large multi-megawatt space solar power can help simply to avoid most of environmental issues. Furthermore, it saves a significant mass by not using a large reactor, while the most efficient for such power sources trajectory with a Venus fly-by can make it possible to perform a comparatively short trip. Unfortunately the size estimation reveals big solar arrays profile. For targeted stowage volume of 15 kW/m³ [47] the size of solar arrays for providing a single megawatt of electrical power would require over 80 cubic meters. Such array would be not only large and fragile, but also would be only suitable for very low accelerations. Besides that there are significant thermal and solar radiation concerns.

3.3.4 Aerobraking at Mars

Aerobraking is not a propulsion concept, but an approach that aims to effectively reduce thrust requirements in stage of placing the spacecraft around Mars. It uses properties of martian atmosphere interaction with a shape of the spaceship and its cross section material as a propellant free decelerator. Potentially it can save a decent mass by fuel and propulsion system lightening. Also it can simplify the Mars lander entry. At the same time this prolongs the trip time and increases the risks the crew might face during the atmospheric drag phase.

3.3.5 In Situ Resource Utilization on Mars

In Situ Resource Utilization (ISRU) is not the least option for many of the Mars flight propositions. A number of hypotheses anticipate potential of this technology to manufacture fuel, oxygen and water directly on the surface of Mars for the planet stay and for the return trip. This might deny the need to bring heavy operating supplies from

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the equipment for resource utilisation. The real potential of this technology cannot be truly estimated until a preparatory robotic mission to Mars will prove how reliable are the claims of the ISRU developers and advocates.

3.4 Trajectory analysis

3.4.1 Earth and Mars relative position

One of the challenges of interplanetary space flights is the fact that objects in the Solar system are not fixed in space and moving around centres of gravity by their unique orbits. Earth and Mars both orbit around the Sun in the same direction but with different velocities and on different distances from the Sun. Characteristics for Earth and Mars are compared in Table 3-2.

Table 3-2. Initial values for Earth and Mars

EARTH MARS

Semi-major axis of

orbit around the Sun 1.00000261 au 1.52371034 au Period of rotation

around the Sun 365.256 days 686.97 days Equatorial radii 6378.14 km 3396.2 km Gravitational constant 398600 km³/s² 42828 km³/s²

Low orbit altitude ~200 km ~300 km

As it was discussed in 2.4 due to the significant difference in periods of rotation around the Sun the distance between Earth and Mars is not a constant value but varies between ~0.5 au and ~2.5 au in a 26 month cycle. That is why Mars missions have to comply to launch windows intervals. This lengthens periods of waiting on different phases of the mission architecture and makes the human mission even more problematic.

By planning a mission to Mars we should also be aware that Mars and Earth orbits do not lie in one plane and there is an inclination change about 1.5° which has to be overcome. However, by proper launch design it is possible to make this inclination change before reaching LEO and to eliminate this problem.

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3.4.2 Transfer orbit

Assuming the transfer on impulsive trajectories the Hohmann transfer is the most appealing option because of its lenient energy demands. For the most of the conservative interplanetary missions proposing chemical propulsion the Hohmann transfer does not have alternatives.

Figure 3-3. Delta-v calculations for the simplified Earth-Mars orbit [17].

Results of calculations presented in Figure 3-3 suggest that delta-v requirement for Hohmann transfer orbit injection and for following insertion into Low Mars Orbit are 3.62 km/s and -2.10 km/s respectively.

In fact a spacecraft can be inserted in different LMOs, accordingly to the applied slowing down delta-v. The delta-v of 2.10 km/s applied at an altitude of 300 km corresponds to a 300 km circular orbit. For delta-v below 2.10 km/s the orbit would be not circular but elliptical with eccentricity inversely proportional to the delta-v. But if delta-v will be very small (below 0.68 km/s [17]) the spacecraft would be not captured by Mars gravity field.

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3.4.3 Transfer duration

There are two main concepts for human Mars missions based upon their interplanetary trajectories and duration of surface sojourn: short stay missions and long stay (conjunction-class) missions [22].

3.4.3.1 Short stay missions

This type of missions uses the advantage of the relative rotation speed of the planets. This looks reasonable to manage a two-way flight and “familiarization” surface mission in just 15 months with minimized requirements for surface facilities/capabilities.

Table 3-3. Overview of the short stay mission options [23].

Delta-v (km/s)

Total mission (years)

Surface stay (days)

Trip to Mars (days)

Return trip (days)

35 1.06 7 112 268

38 1.12 38 106 264

40 1.14 55 102 259

45 1.25 117 94 245

50 1.30 147 88 240

55 1.37 193 81 226

60 1.41 223 75 217

This concept has a number of critical problems associated with it. From Table 3-3 we can see that delta-v requirements are unfeasibly high. This increases requirements for IMLEO and moreover declares needs in developing new “exotic”

propulsion system [2]. Another concern is that 85% of the mission time is supposed to be a long duration deep space flight. This means increasing of risk factors and necessity of extra protection. But the bigger challenge is that the proposed trajectory for returning flight aimed to have perihelion near orbit of Venus in extreme closeness to the Sun (Figure 3-4). Temperature issues might transfer the human Mars mission using this concept from the “futuristic missions” category into category of “visionary space missions”.

However, advocates of the Venus swing-by trajectory might rightfully indicate that this option is beneficial from solar energy access point of view and that the thermal

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control issue is less critical than the delta-v requirements. Furthermore, the fact that the crew is spending most of the mission time in the Space environment (hazardous, not fully discovered) rather than at the surface of Mars (almost unknown environment) could be considered as an advantage.

Figure 3-4. Trajectory for the short stay mission [24].

3.4.3.2 Long stay missions

This type of missions is considered as a more realistic. Long stay missions use optimum trajectories in both directions. The propellant requirements are not unreachable (delta-v is about 13 km/s). But because of the launch windows limitations there is inevitable necessity to organize almost 3 year mission with no possibilities for an early return and for abort options.

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Table 3-4. Overview of the long stay mission options [23].

Delta-v (km/s)

Total mission (years)

Surface stay (days)

One-way trip (days)

11 2.70 456 265

12 2.72 500 246

13 2.70 548 219

14 2.68 577 201

15 2.65 599 184

18 2.64 621 171

20 2.62 639 159

25 2.61 672 141

30 2.60 694 128

35 2.58 719 111

40 2.56 741 97

45 2.55 759 86

Figure 3-5. Trajectory for the long stay mission [24].

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3.5 Earth to low earth orbit

At the moment all space mission architectures (even the most futuristic) have the same departure point – planet Earth. That is why one of the important milestones is choosing a proper launch vehicle or a combination of vehicles. This choice will determine number of launches and necessity to perform orbit rendezvous and assemblies.

The most common and widely used existent launch vehicles like Soyuz-U have typical capability to deliver not more than 7-8 t to LEO. The heaviest ones deliver 20 t.

NASA’s launch vehicle Saturn V delivered payload of 120 t weight to LEO for the Apollo missions. Soviet rocket proposal Vulcan (8 Zenit booster rockets modification for Energia-M rocket) could deliver up to 200 t [25], but was cancelled along with successfully tested 100 t Energia due to the financial concerns.

The human mission to Mars regardless of its architecture will require unprecedented big IMLEO (in case of conservative architectures the initial mass will probably exceed 1000 t [17]), and hence there is an exigency in consideration opportunities for designing a new generation of super heavy lift launch systems [26].

Developing such system (a high loading capacity rocket, launch facilities, ground processing capabilities) inevitably will be a significant part of the total mission cost.

NASA recent hopes were in Constellation human spaceflight program with creating cargo launch vehicle Ares V with a promising 188 t to LEO capability [27].

But in 2010 this program was cancelled. New alternative proposition is a Space Shuttle- derived heavy launch vehicle SLS (Space Launch System). By optimistic expectations its capacity could reach 130 t. This means that in order to perform a human Mars mission there will be necessity to perform complicated orbit assembly operations.

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Figure 3-6. The SLS concept art interpretation [28].

3.6 Trans Planet Injections

This phase of the mission architecture provokes most of speculations and discussions. Mars exploration sceptics rightly use the required mission time as their argument against funding human deep space flight programs. On the other hand there is no consensus between advocates of such programs either. Using of existent propulsion systems implies huge traveling times by definition. Looking for a new type of propulsion systems for a nonconservative mission architecture will lead to mission start delays, ambiguous redistribution of funds, with no guarantee of reducing mission time.

3.6.1 High thrust approach

In general the objectives of the Trans Planet Injection (TPI) can be determined as formulation of delta-v requirements and mass sent toward planet estimations. When we consider a transfer without producing thrust during the most of the transfer (the impulsive approach) the equation to calculate the required delta-v for Trans Mars Injection (TMI) can be derived as following [17]:

√

(3-2)

where v_tot is total speed, v_orb is Orbit velocity, v_inf is the speed rendered to the spacecraft after departing from the gravitational field, R_LEO is the radius of low orbit measured from the centre of planet (Earth for TMI). The similar equation is true for Trans Earth Injection (TEI):

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√

(3-3)

For TMI the values of appropriate vinf vary within launch windows cycle. At each launch opportunity a corresponding map can be calculated that plots arrival dates vs. departure date. In the fast Type I (high energy trips) the spacecraft goes less than 180° around the Sun. In the slow Type II (low energy trips) the spacecraft goes more than 180° around the Sun. The optimum boundary between them is the Hohmann transfer, which by definition is the minimum delta-v transfer between two coplanar circular orbits [7].

If we assume LEO ~200km, and 3 separate modules mission we get:

- Lowest energy cargo delivery (400 days) 3.80 km/s - High energy cargo delivery (250 days) 4.18 km/s - Highest energy crew delivery (175 days) 4.43 km/s

Logically a higher delta-v reduces the trip time by the cost of payload fraction lost in IMLEO. Comparison of different energy trip types is presented in Table 3-5.

Table 3-5. Different energy types of trips characteristic [17].

Year

Lowest energy

trip Low energy trip Fast energy trip Fastest energy trip

V_inf²

(km/s)²

Trip

(days)

ΔV

(km/s)

V_inf²

(km/s)²

Trip

(days)

ΔV

(km/s)

V_inf²

(km/s)²

Trip

(days)

ΔV

(km/s)

V_inf²

(km/s)²

Trip

(days)

ΔV

(km/s)

2016 9 300 3.63 12 175 3.76 15 150 3.89

2018 ¹⁴ ²⁸⁰ ^3.84 ¹² ¹⁷⁵ ^3.76 ¹⁵ ¹⁵⁰ ^4.89

2020 18 400 4.01 23 340 4.22 16 175 3.93 20 150 4.10 2022 14.5 400 3.86 17 350 3.97 22 175 4.18

2024 ¹³ ³⁵⁰ ^3.80 ¹⁶ ³²⁰ ^3.93 ²² ²⁰⁰ ^4.18 ²⁸ ¹⁷⁵ ^4.43 2026 11 300 3.71 12.5 275 3.78 17.5 200 3.99 23 175 4.22

3.6.2 Low thrust approach

Trans planet injection for low thrust propulsion systems is significantly different

Opportunities From Using The Variable Specific Impulse Magnetoplasma Rocket For Manned Missions To Mars

MASTER'S THESIS

Opportunities From Using The Variable Specific Impulse Magnetoplasma Rocket

For Manned Missions To Mars

Vladimir Gurevich

CRANFIELD UNIVERSITY

Vladimir Gurevich

OPPORTUNITIES FROM USING THE VARIABLE SPECIFIC IMPULSE MAGNETOPLASMA ROCKET FOR MANNED MISSIONS

TO MARS

School of Engineering

Astronautics & Space Engineering (SpaceMaster)

MSc

Academic Year: 2011 - 2012

Supervisors Dr. Jennifer Kingston Professor Bob Parkinson

May 2012

CRANFIELD UNIVERSITY

School of Engineering

Astronautics & Space Engineering (SpaceMaster)

MSc

Academic Year 2011 - 2012

Vladimir Gurevich

OPPORTUNITIES FROM USING THE VARIABLE SPECIFIC IMPULSE MAGNETOPLASMA ROCKET FOR MANNED MISSIONS

TO MARS

Supervisors Dr. Jennifer Kingston Professor Bob Parkinson

May 2012

This thesis is submitted in partial fulfilment of the requirements for the degree of MSc in Astronautics and Space Engineering.

© Cranfield University 2012. All rights reserved. No part of this

ABSTRACT

ACKNOWLEDGEMENTS

TABLE OF CONTENTS

LIST OF FIGURES

LIST OF TABLES

LIST OF ABBREVIATIONS

1 INTRODUCTION

2 FUNDAMENTALS OF INTERPLANETARY MANNED MISSIONS

3 MISSION ARCHITECTURE