DOCSLIB.ORG

Interplanetary Transfers, Which Are Optimal with Respect to Fuel Consumption

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Interplanetary Transfers, Which Are Optimal with Respect to Fuel Consumption Astrodynamics (AERO0024) 10. Interplanetary Trajectories Gaëtan Kerschen Space Structures & Systems Lab (S3L) Motivation 2 6. Interplanetary Trajectories 6.1 Patched conic method 6.2 Lambert’s problem 6.3 Gravity assist 3 6. Interplanetary Trajectories 6.1 Patched conic method 6.1.1 Problem statement 6.1.2 Sphere of influence 6.1.3 Planetary departure 6.1.4 Hohmann transfer 6.1.5 Planetary arrival 6.1.6 Sensitivity analysis and launch windows 6.1.7 Examples 4 ? Hint #1: design the Earth-Mars transfer using known concepts Hint #2: division into simpler problems Hint #3: patched conic method 5 What Transfer Orbit ? Constraints ? Dep. “V1” Arr. Dep. Arr. Dep. “V2” Arr. Motion in the heliocentric reference frame6 Planetary Departure ? Constraints ? v v Earth/ Sun V ? 1 Motion in the planetary reference frame7 Planetary Arrival ? Similar Reasoning Transfer ellipse SOI SOI Arrival Departure hyperbola hyperbola 8 Patched Conic Method Three conics to patch: 1. Outbound hyperbola (departure) 2. The Hohmann transfer ellipse (interplanetary travel) 3. Inbound hyperbola (arrival) 6.1.1 Problem statement 9 Patched Conic Method Approximate method that analyzes a mission as a sequence of 2-body problems, with one body always being the spacecraft. If the spacecraft is close enough to one celestial body, the gravitational forces due to other planets can be neglected. The region inside of which the approximation is valid is called the sphere of influence (SOI) of the celestial body. If the spacecraft is not inside the SOI of a planet, it is considered to be in orbit around the sun. 6.1.1 Problem statement 10 Patched Conic Method Very useful for preliminary mission design (delta-v requirements and flight times). But actual mission design and execution employ the most accurate possible numerical integration techniques. 6.1.1 Problem statement 11 Sphere of Influence (SOI) ? Let’s assume that a spacecraft is within the Earth’s SOI if the gravitational force due to Earth is larger than the gravitational force due to the sun. GmE m sat Gm S m sat 22 rrE,, sat S sat 5 rE, sat 2.5 10 km 6.1.2 Sphere of influence 12 Sphere of Influence (SOI) Third body: sun or planet m Spacecraft d j m2 r ρ d ρ rr3 Gm j 3 3 rd m1 Central body: Disturbing O sun or planet function (L04) 6.1.2 Sphere of influence 13 If the Spacecraft Orbits the Planet G m m p:planet pv rsv rsp rr Gm v: vehicle pv3 pv s 3 3 rpv r sv r sp s:sun rpv A p P s Perturbation acceleration due Primary to the sun gravitational acceleration due to the planet 6.1.2 Sphere of influence 14 If the Spacecraft Orbits the Sun G msv m rrpv sp p:planet rr Gm v: vehicle sv3 sv p 3 3 rsv r pv r sp s:sun rsv A s P p Perturbation acceleration due Primary to the planet gravitational acceleration due to the sun 6.1.2 Sphere of influence 15 SOI: Correct Definition due to Laplace It is the surface along which the magnitudes of the acceleration satisfy: P P p s AAsp Measure of the planet’s Measure of the deviation of influence on the orbit of the the vehicle’s orbit from the vehicle relative to the sun Keplerian orbit arising from the planet acting by itself 2 5 mp rrSOI sp ms 6.1.2 Sphere of influence 16 SOI: Correct Definition due to Laplace P P If p s the spacecraft is inside the SOI of the planet. AAsp Ap The previous (incorrect) definition was 1 As The moon lied outside the SOI and was in orbit about the sun like an asteroid ! 6.1.2 Sphere of influence 17 SOI Radii SOI radius Planet SOI Radius (km) (body radii) Mercury 1.13x105 45 Venus 6.17x105 100 OK ! Earth 9.24x105 145 Mars 5.74x105 170 Jupiter 4.83x107 677 Neptune 8.67x107 3886 6.1.2 Sphere of influence 18 Validity of the Patched Conic Method The Earth’s SOI is 145 Earth radii. This is extremely large compared to the size of the Earth: The velocity relative to the planet on an escape hyperbola is considered to be the hyperbolic excess velocity vector. vvSOI This is extremely small with respect to 1AU: During the elliptic transfer, the spacecraft is considered to be under the influence of the Sun’s gravity only. In other words, it follows an unperturbed Keplerian orbit around the Sun. 6.1.2 Sphere of influence 19 Outbound Hyperbola The spacecraft necessarily escapes using a hyperbolic trajectory relative to the planet. When this velocity vector is added to the planet’s heliocentric velocity, the result is the spacecraft’s heliocentric velocity on the Hyperbolic excess interplanetary elliptic transfer orbit at the speed SOI in the solar system. 2 2 2 2 2 Lecture 02: v v vesc v SOI v esc Is vSOI the velocity on the transfer orbit ? 6.1.3 Planetary departure 20 Magnitude of VSOI The velocity vD of the spacecraft relative to the sun is imposed by the Hohmann transfer (i.e., velocity on the transfer orbit). H. Curtis, Orbital Mechanics for Engineering Students, Elsevier. 6.1.3 Planetary departure 21 Magnitude of VSOI By subtracting the known value of the velocity v1 of the planet relative to the sun, one obtains the hyperbolic excess speed on the Earth escape hyperbola. 2R v v v sun 2 1 v SOI D 1 RRR 1 1 2 Imposed Lecture05 Known 6.1.3 Planetary departure 22 Direction of VSOI What should be the direction of vSOI ? For a Hohmann transfer, it should be parallel to v1. 6.1.3 Planetary departure 23 Parking Orbit A spacecraft is ordinary launched into an interplanetary trajectory from a circular parking orbit. Its radius equals the periapse radius rp of the departure hyperbola. H. Curtis, Orbital Mechanics for Engineering Students, Elsevier. 6.1.3 Planetary departure 24 ΔV Magnitude and Location 2 2 h rvp Lecture 02: r e 1 p (1 e ) known known v 2 a e 1 2 2 h h rp v 2 r h 1 v p a e2 1 h 1 1 v vpc v cos rrpp e 6.1.3 Planetary departure 25 Planetary Departure: Graphically Departure to outer or inner planet ? H. Curtis, Orbital Mechanics for Engineering Students, Elsevier. 6.1.3 Planetary departure 26 Circular, Coplanar Orbits for Most Planets Inclination of the orbit Planet Eccentricity to the ecliptic plane Mercury 7.00º 0.206 Venus 3.39º 0.007 Earth 0.00º 0.017 Mars 1.85º 0.094 Jupiter 1.30º 0.049 Saturn 2.48º 0.056 Uranus 0.77º 0.046 Neptune 1.77º 0.011 Pluto 17.16º 0.244 6.1.4 Hohmann transfer 27 Governing Equations vD sun 2R2 vv 1 v1 D 1 RRR 1 1 2 sun 2R1 vA vv2 A 1 RRR v2 2 1 2 Signs ? v2 - vA, vD - v1 >0 for transfer to an outer planet v2 - vA, vD - v1 <0 for transfer to an inner planet 6.1.4 Hohmann transfer 28 Schematically Transfer to outer planet Transfer to inner planet 6.1.4 Hohmann transfer 29 Arrival at an Outer Planet For an outer planet, the spacecraft’s heliocentric approach velocity vA is smaller in magnitude than that of the planet v2. v2 v vA vv2 A v 2 and v have opposite signs. 6.1.5 Planetary arrival 30 Arrival at an Outer Planet The spacecraft crosses the forward portion of the SOI 6.1.5 Planetary arrival 31 Enter into an Elliptic Orbit If the intent is to go into orbit around the planet, then must be chosen so that the v burn at periapse will occur at the correct altitude above the planet. 2 rp 1 2 rvp h(1 e )2 2 (1 e ) v vp,, hyp v p capture v rp r p r p r p 6.1.5 Planetary arrival 32 Planetary Flyby 2 rvp Otherwise, the specacraft will simplye continue1 past periapse on a flyby trajectory exiting the SOI with the same relative speed v it entered but with the velocity vector rotated through the turn angle . 1 2sin1 e 6.1.5 Planetary arrival 33 Sensitivity Analysis: Departure The maneuver occurs well within the SOI, which is just a point on the scale of the solar system. One may therefore ask what effects small errors in position and velocity (rp and vp) at the maneuver point have on the trajectory (target radius R2 of the heliocentric Hohmann transfer ellipse). 21 v R 2 r r v 21p p p 2 R2 Rv1 D vD v r p r p v D v p 1 2sun 6.1.6 Sensitivity analysis and launch windows 34 Sensitivity Analysis: Earth-Mars, 300km Orbit 11 3 2 3 2 sun 1.327 10 km /ss ,1 398600 km / 66 R12149.6 10 km, R 227.9 10 km, rp 6678 km vD 32.73 km / s , v 2.943 km / s R rv 2 3.127pp 6.708 R2 rpp v A 0.01% variation in the burnout speed vp changes the target radius by 0.067% or 153000 km. A 0.01% variation in burnout radius rp (670 m !) produces an error over 70000 km. 6.1.6 Sensitivity analysis and launch windows 35 Sensitivity Analysis: Launch Errors Ariane V Trajectory correction maneuvers are clearly mandatory. 6.1.6 Sensitivity analysis and launch windows 36 Sensitivity Analysis: Arrival The heliocentric velocity of Mars in its orbit is roughly 24km/s.
Load more

Recommended publications
  • Electric Propulsion System Scaling for Asteroid Capture-And-Return Missions
    Electric propulsion system scaling for asteroid capture-and-return missions Justin M. Little⇤ and Edgar Y. Choueiri† Electric Propulsion and Plasma Dynamics Laboratory, Princeton University, Princeton, NJ, 08544 The requirements for an electric propulsion system needed to maximize the return mass of asteroid capture-and-return (ACR) missions are investigated in detail. An analytical model is presented for the mission time and mass balance of an ACR mission based on the propellant requirements of each mission phase. Edelbaum’s approximation is used for the Earth-escape phase. The asteroid rendezvous and return phases of the mission are modeled as a low-thrust optimal control problem with a lunar assist. The numerical solution to this problem is used to derive scaling laws for the propellant requirements based on the maneuver time, asteroid orbit, and propulsion system parameters. Constraining the rendezvous and return phases by the synodic period of the target asteroid, a semi- empirical equation is obtained for the optimum specific impulse and power supply. It was found analytically that the optimum power supply is one such that the mass of the propulsion system and power supply are approximately equal to the total mass of propellant used during the entire mission. Finally, it is shown that ACR missions, in general, are optimized using propulsion systems capable of processing 100 kW – 1 MW of power with specific impulses in the range 5,000 – 10,000 s, and have the potential to return asteroids on the order of 103 104 tons. − Nomenclature
    [Show full text]
  • Optimisation of Propellant Consumption for Power Limited Rockets
    Delft University of Technology Faculty Electrical Engineering, Mathematics and Computer Science Delft Institute of Applied Mathematics Optimisation of Propellant Consumption for Power Limited Rockets. What Role do Power Limited Rockets have in Future Spaceflight Missions? (Dutch title: Optimaliseren van brandstofverbruik voor vermogen gelimiteerde raketten. De rol van deze raketten in toekomstige ruimtevlucht missies. ) A thesis submitted to the Delft Institute of Applied Mathematics as part to obtain the degree of BACHELOR OF SCIENCE in APPLIED MATHEMATICS by NATHALIE OUDHOF Delft, the Netherlands December 2017 Copyright c 2017 by Nathalie Oudhof. All rights reserved. BSc thesis APPLIED MATHEMATICS \ Optimisation of Propellant Consumption for Power Limited Rockets What Role do Power Limite Rockets have in Future Spaceflight Missions?" (Dutch title: \Optimaliseren van brandstofverbruik voor vermogen gelimiteerde raketten De rol van deze raketten in toekomstige ruimtevlucht missies.)" NATHALIE OUDHOF Delft University of Technology Supervisor Dr. P.M. Visser Other members of the committee Dr.ir. W.G.M. Groenevelt Drs. E.M. van Elderen 21 December, 2017 Delft Abstract In this thesis we look at the most cost-effective trajectory for power limited rockets, i.e. the trajectory which costs the least amount of propellant. First some background information as well as the differences between thrust limited and power limited rockets will be discussed. Then the optimal trajectory for thrust limited rockets, the Hohmann Transfer Orbit, will be explained. Using Optimal Control Theory, the optimal trajectory for power limited rockets can be found. Three trajectories will be discussed: Low Earth Orbit to Geostationary Earth Orbit, Earth to Mars and Earth to Saturn. After this we compare the propellant use of the thrust limited rockets for these trajectories with the power limited rockets.
    [Show full text]
  • Asteroid Retrieval Mission
    Where you can put your asteroid Nathan Strange, Damon Landau, and ARRM team NASA/JPL-CalTech © 2014 California Institute of Technology. Government sponsorship acknowledged. Distant Retrograde Orbits Works for Earth, Moon, Mars, Phobos, Deimos etc… very stable orbits Other Lunar Storage Orbit Options • Lagrange Points – Earth-Moon L1/L2 • Unstable; this instability enables many interesting low-energy transfers but vehicles require active station keeping to stay in vicinity of L1/L2 – Earth-Moon L4/L5 • Some orbits in this region is may be stable, but are difficult for MPCV to reach • Lunar Weakly Captured Orbits – These are the transition from high lunar orbits to Lagrange point orbits – They are a new and less well understood class of orbits that could be long term stable and could be easier for the MPCV to reach than DROs – More study is needed to determine if these are good options • Intermittent Capture – Weakly captured Earth orbit, escapes and is then recaptured a year later • Earth Orbit with Lunar Gravity Assists – Many options with Earth-Moon gravity assist tours Backflip Orbits • A backflip orbit is two flybys half a rev apart • Could be done with the Moon, Earth or Mars. Backflip orbit • Lunar backflips are nice plane because they could be used to “catch and release” asteroids • Earth backflips are nice orbits in which to construct things out of asteroids before sending them on to places like Earth- Earth or Moon orbit plane Mars cyclers 4 Example Mars Cyclers Two-Synodic-Period Cycler Three-Synodic-Period Cycler Possibly Ballistic Chen, et al., “Powered Earth-Mars Cycler with Three Synodic-Period Repeat Time,” Journal of Spacecraft and Rockets, Sept.-Oct.
    [Show full text]
  • Up, Up, and Away by James J
    www.astrosociety.org/uitc No. 34 - Spring 1996 © 1996, Astronomical Society of the Pacific, 390 Ashton Avenue, San Francisco, CA 94112. Up, Up, and Away by James J. Secosky, Bloomfield Central School and George Musser, Astronomical Society of the Pacific Want to take a tour of space? Then just flip around the channels on cable TV. Weather Channel forecasts, CNN newscasts, ESPN sportscasts: They all depend on satellites in Earth orbit. Or call your friends on Mauritius, Madagascar, or Maui: A satellite will relay your voice. Worried about the ozone hole over Antarctica or mass graves in Bosnia? Orbital outposts are keeping watch. The challenge these days is finding something that doesn't involve satellites in one way or other. And satellites are just one perk of the Space Age. Farther afield, robotic space probes have examined all the planets except Pluto, leading to a revolution in the Earth sciences -- from studies of plate tectonics to models of global warming -- now that scientists can compare our world to its planetary siblings. Over 300 people from 26 countries have gone into space, including the 24 astronauts who went on or near the Moon. Who knows how many will go in the next hundred years? In short, space travel has become a part of our lives. But what goes on behind the scenes? It turns out that satellites and spaceships depend on some of the most basic concepts of physics. So space travel isn't just fun to think about; it is a firm grounding in many of the principles that govern our world and our universe.
    [Show full text]
  • Future Space Launch Vehicles
    Future Space Launch Vehicles S. Krishnan Professor of Aerospace Engineering Indian Institute of Technology Madras Chennnai - 600 036, India (Written in 2001) Introduction Space technology plays a very substantial role in the economical growth and the national security needs of any country. Communications, remote sensing, weather forecasting, navigation, tracking and data relay, surveillance, reconnaissance, and early warning and missile defence are its dependent user-technologies. Undoubtedly, space technology has become the backbone of global information highway. In this technology, the two most important sub-technologies correspond to spacecraft and space launch vehicles. Spacecraft The term spacecraft is a general one. While the spacecraft that undertakes a deep space mission bears this general terminology, the one that orbits around a planet is also a spacecraft but called specifically a satellite more strictly an artificial satellite, moons around their planets being natural satellites. Cassini is an example for a spacecraft. This was developed under a cooperative project of NASA, the European Space Agency, and the Italian Space Agency. Cassini spacecraft, launched in 1997, is continuing its journey to Saturn (about 1274 million km away from Earth), where it is scheduled to begin in July 2004, a four- year exploration of Saturn, its rings, atmosphere, and moons (18 in number). Cassini executed two gravity-assist flybys (or swingbys) of Venus one in April 1998 and one in June 1999 then a flyby of Earth in August 1999, and a flyby of Jupiter (about 629 million km away) in December 2000, see Fig. 1. We may note here with interest that ISRO (Indian Space Research Organisation) is thinking of a flyby mission of a spacecraft around Moon (about 0.38 million km away) by using its launch vehicle PSLV.
    [Show full text]
  • Gravity-Assist Trajectories to Jupiter Using Nuclear Electric Propulsion
    AAS 05-398 Gravity-Assist Trajectories to Jupiter Using Nuclear Electric Propulsion ∗ ϒ Daniel W. Parcher ∗∗ and Jon A. Sims ϒϒ This paper examines optimal low-thrust gravity-assist trajectories to Jupiter using nuclear electric propulsion. Three different Venus-Earth Gravity Assist (VEGA) types are presented and compared to other gravity-assist trajectories using combinations of Earth, Venus, and Mars. Families of solutions for a given gravity-assist combination are differentiated by the approximate transfer resonance or number of heliocentric revolutions between flybys and by the flyby types. Trajectories that minimize initial injection energy by using low resonance transfers or additional heliocentric revolutions on the first leg of the trajectory offer the most delivered mass given sufficient flight time. Trajectory families that use only Earth gravity assists offer the most delivered mass at most flight times examined, and are available frequently with little variation in performance. However, at least one of the VEGA trajectory types is among the top performers at all of the flight times considered. INTRODUCTION The use of planetary gravity assists is a proven technique to improve the performance of interplanetary trajectories as exemplified by the Voyager, Galileo, and Cassini missions. Another proven technique for enhancing the performance of space missions is the use of highly efficient electric propulsion systems. Electric propulsion can be used to increase the mass delivered to the destination and/or reduce the trip time over typical chemical propulsion systems.1,2 This technology has been demonstrated on the Deep Space 1 mission 3 − part of NASA’s New Millennium Program to validate technologies which can lower the cost and risk and enhance the performance of future missions.
    [Show full text]
  • Determination of Optimal Earth-Mars Trajectories to Target the Moons Of
    The Pennsylvania State University The Graduate School Department of Aerospace Engineering DETERMINATION OF OPTIMAL EARTH-MARS TRAJECTORIES TO TARGET THE MOONS OF MARS A Thesis in Aerospace Engineering by Davide Conte 2014 Davide Conte Submitted in Partial Fulfillment of the Requirements for the Degree of Master of Science May 2014 ii The thesis of Davide Conte was reviewed and approved* by the following: David B. Spencer Professor of Aerospace Engineering Thesis Advisor Robert G. Melton Professor of Aerospace Engineering Director of Undergraduate Studies George A. Lesieutre Professor of Aerospace Engineering Head of the Department of Aerospace Engineering *Signatures are on file in the Graduate School iii ABSTRACT The focus of this thesis is to analyze interplanetary transfer maneuvers from Earth to Mars in order to target the Martian moons, Phobos and Deimos. Such analysis is done by solving Lambert’s Problem and investigating the necessary targeting upon Mars arrival. Additionally, the orbital parameters of the arrival trajectory as well as the relative required ΔVs and times of flights were determined in order to define the optimal departure and arrival windows for a given range of date. The first step in solving Lambert’s Problem consists in finding the positions and velocities of the departure (Earth) and arrival (Mars) planets for a given range of dates. Then, by solving Lambert’s problem for various combinations of departure and arrival dates, porkchop plots can be created and examined. Some of the key parameters that are plotted on porkchop plots and used to investigate possible transfer orbits are the departure characteristic energy, C3, and the arrival v∞.
    [Show full text]
  • A Study of Trajectories to the Neptune System Using Gravity Assists
    Advances in Space Research 40 (2007) 125–133 www.elsevier.com/locate/asr A study of trajectories to the Neptune system using gravity assists C.R.H. Solo´rzano a, A.A. Sukhanov b, A.F.B.A. Prado a,* a National Institute for Space Research (INPE), Sa˜o Jose´ dos Campos, SP 12227-010, Brazil b Space Research Institute (IKI), Russian Academy of Sciences 84/32, Profsoyuznaya St., 117810 Moscow, Russia Received 6 September 2006; received in revised form 21 February 2007; accepted 21 February 2007 Abstract At the present time, the search for the knowledge of our Solar System continues effective. NASA’s Solar System Exploration theme listed a Neptune mission as one of its top priorities for the mid-term (2008–2013). From the technical point of view, gravity assist is a proven technique in interplanetary exploration, as exemplified by the missions Voyager, Galileo, and Cassini. Here, a mission to Neptune for the mid-term (2008–2020) is proposed, with the goal of studying several schemes for the mission. A direct transfer to Neptune is considered and also Venus, Earth, Jupiter, and Saturn gravity assists are used for the transfer to Neptune, which represent new contributions for a possible real mission. We show several schemes, including or not the braking maneuver near Neptune, in order to find a good compromise between the DV and the time of flight to Neptune. After that, a study is made to take advantage of an asteroid flyby opportunity, when the spacecraft passes by the main asteroid belt. Results for a mission that makes an asteroid flyby with the asteroid 1931 TD3 is shown.
    [Show full text]
  • Links to Education Standards • Introduction to Exoplanets • Glossary of Terms • Classroom Activities Table of Contents Photo © Michael Malyszko Photo © Michael
    an Educator’s Guid E INSIDE • Links to Education Standards • Introduction to Exoplanets • Glossary of Terms • Classroom Activities Table of Contents Photo © Michael Malyszko Photo © Michael For The TeaCher About This Guide .............................................................................................................................................. 4 Connections to Education Standards ................................................................................................... 5 Show Synopsis .................................................................................................................................................. 6 Meet the “Stars” of the Show .................................................................................................................... 8 Introduction to Exoplanets: The Basics .................................................................................................. 9 Delving Deeper: Teaching with the Show ................................................................................................13 Glossary of Terms ..........................................................................................................................................24 Online Learning Tools ..................................................................................................................................26 Exoplanets in the News ..............................................................................................................................27 During
    [Show full text]
  • Binary Asteroids and the Formation of Doublet Craters
    ICARUS 124, 372±391 (1996) ARTICLE NO. 0215 Binary Asteroids and the Formation of Doublet Craters WILLIAM F. BOTTKE,JR. Division of Geological and Planetary Sciences, California Institute of Technology, Mail Code 170-25, Pasadena, California 91125 E-mail: [email protected] AND H. JAY MELOSH Lunar and Planetary Laboratory, University of Arizona, Tucson, Arizona 85721 Received April 29, 1996; revised August 14, 1996 found we could duplicate the observed fraction of doublet cra- At least 10% (3 out of 28) of the largest known impact craters ters found on Earth, Venus, and Mars. Our results suggest on Earth and a similar fraction of all impact structures on that any search for asteroid satellites should place emphasis Venus are doublets (i.e., have a companion crater nearby), on km-sized Earth-crossing asteroids with short-rotation formed by the nearly simultaneous impact of objects of compa- periods. 1996 Academic Press, Inc. rable size. Mars also has doublet craters, though the fraction found there is smaller (2%). These craters are too large and too far separated to have been formed by the tidal disruption 1. INTRODUCTION of an asteroid prior to impact, or from asteroid fragments dispersed by aerodynamic forces during entry. We propose that Two commonly held paradigms about asteroids and some fast rotating rubble-pile asteroids (e.g., 4769 Castalia), comets are that (a) they are composed of non-fragmented after experiencing a close approach with a planet, undergo tidal chunks of rock or rock/ice mixtures, and (b) they are soli- breakup and split into multiple co-orbiting fragments.
    [Show full text]
  • Humanity and Space
    10/17/2012!! !!!!!! Project Number: MH-1207 Humanity and Space An Interactive Qualifying Project Submitted to WORCESTER POLYTECHNIC INSTITUTE In partial fulfillment for the Degree of Bachelor of Science by: Matthew Beck Jillian Chalke Matthew Chase Julia Rugo Professor Mayer H. Humi, Project Advisor Abstract Our IQP investigates the possible functionality of another celestial body as an alternate home for mankind. This project explores the necessary technological advances for moving forward into the future of space travel and human development on the Moon and Mars. Mars is the optimal candidate for future human colonization and a stepping stone towards humanity’s expansion into outer space. Our group concluded space travel and interplanetary exploration is possible, however international political cooperation and stability is necessary for such accomplishments. 2 Executive Summary This report provides insight into extraterrestrial exploration and colonization with regards to technology and human biology. Multiple locations have been taken into consideration for potential development, with such qualifying specifications as resources, atmospheric conditions, hazards, and the environment. Methods of analysis include essential research through online media and library resources, an interview with NASA about the upcoming Curiosity mission to Mars, and the assessment of data through mathematical equations. Our findings concerning the human aspect of space exploration state that humanity is not yet ready politically and will not be able to biologically withstand the hazards of long-term space travel. Additionally, in the field of robotics, we have the necessary hardware to implement adequate operational systems yet humanity lacks the software to implement rudimentary Artificial Intelligence. Findings regarding the physics behind rocketry and space navigation have revealed that the science of spacecraft is well-established.
    [Show full text]
  • The Terrestrial Planet V Hypothesis As the Mechanism for the Origin of the Late Heavy Bombardment
    A&A 535, A41 (2011) Astronomy DOI: 10.1051/0004-6361/201117336 & c ESO 2011 Astrophysics The terrestrial Planet V hypothesis as the mechanism for the origin of the late heavy bombardment R. Brasser and A. Morbidelli Dep. Cassiopée, University of Nice–Sophia Antipolis, CNRS, Observatoire de la Côte d’Azur, 06304 Nice, France e-mail: [email protected] Received 24 May 2011 / Accepted 13 September 2011 ABSTRACT In this study we attempt to model, with numerical simulations, the scenario for the origin of the late heavy bombardment (LHB) proposed in Chambers (2007, Icarus, 189, 386). Chambers suggested that the orbit of a fictitious sub-Mars sized embryo outside Mars became unstable and crossed a part of the asteroid belt until it was removed through ejection or collision. Chambers demonstrated that this kind of evolution can occur at the LHB time, but he did not check whether enough material could be liberated from the belt to be consistent with the magnitude of the LHB. We do that here. In addition to the asteroid belt, we also consider putative belts of small objects located between the orbits of the Earth and Venus (EV belt) or between the Earth and Mars (EM belt). We find that delivering enough material to the Moon from the asteroid belt requires at least a 5 lunar mass embryo crossing the entire asteroid belt up to 3.3 AU for longer than 300 Myr on a low-inclination orbit. We do not witness this to occur in over a hundred simulations. An alternative approach, which relies only on decimating the inner belt (<2.5 AU), requires that the inner belt originally contained 4 to 13 times more mass than the original outer belt.
    [Show full text]