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Satellite Splat II: an Inelastic Collision with a Surface-Launched Projectile and the Maximum Orbital Radius for Planetary Impact
European Journal of Physics Eur. J. Phys. 37 (2016) 045004 (10pp) doi:10.1088/0143-0807/37/4/045004 Satellite splat II: an inelastic collision with a surface-launched projectile and the maximum orbital radius for planetary impact Philip R Blanco1,2,4 and Carl E Mungan3 1 Department of Physics and Astronomy, Grossmont College, El Cajon, CA 92020- 1765, USA 2 Department of Astronomy, San Diego State University, San Diego, CA 92182- 1221, USA 3 Physics Department, US Naval Academy, Annapolis, MD 21402-1363, USA E-mail: [email protected] Received 26 January 2016, revised 3 April 2016 Accepted for publication 14 April 2016 Published 16 May 2016 Abstract Starting with conservation of energy and angular momentum, we derive a convenient method for determining the periapsis distance of an orbiting object, by expressing its velocity components in terms of the local circular speed. This relation is used to extend the results of our previous paper, examining the effects of an adhesive inelastic collision between a projectile launched from the surface of a planet (of radius R) and an equal-mass satellite in a circular orbit of radius rs. We show that there is a maximum orbital radius rs ≈ 18.9R beyond which such a collision cannot cause the satellite to impact the planet. The difficulty of bringing down a satellite in a high orbit with a surface- launched projectile provides a useful topic for a discussion of orbital angular momentum and energy. The material is suitable for an undergraduate inter- mediate mechanics course. Keywords: orbital motion, momentum conservation, energy conservation, angular momentum, ballistics (Some figures may appear in colour only in the online journal) 4 Author to whom any correspondence should be addressed. -
Orbital Mechanics Course Notes
Orbital Mechanics Course Notes David J. Westpfahl Professor of Astrophysics, New Mexico Institute of Mining and Technology March 31, 2011 2 These are notes for a course in orbital mechanics catalogued as Aerospace Engineering 313 at New Mexico Tech and Aerospace Engineering 362 at New Mexico State University. This course uses the text “Fundamentals of Astrodynamics” by R.R. Bate, D. D. Muller, and J. E. White, published by Dover Publications, New York, copyright 1971. The notes do not follow the book exclusively. Additional material is included when I believe that it is needed for clarity, understanding, historical perspective, or personal whim. We will cover the material recommended by the authors for a one-semester course: all of Chapter 1, sections 2.1 to 2.7 and 2.13 to 2.15 of Chapter 2, all of Chapter 3, sections 4.1 to 4.5 of Chapter 4, and as much of Chapters 6, 7, and 8 as time allows. Purpose The purpose of this course is to provide an introduction to orbital me- chanics. Students who complete the course successfully will be prepared to participate in basic space mission planning. By basic mission planning I mean the planning done with closed-form calculations and a calculator. Stu- dents will have to master additional material on numerical orbit calculation before they will be able to participate in detailed mission planning. There is a lot of unfamiliar material to be mastered in this course. This is one field of human endeavor where engineering meets astronomy and ce- lestial mechanics, two fields not usually included in an engineering curricu- lum. -
Control of Long-Term Low-Thrust Small Satellites Orbiting Mars
SSC18-PII-26 Control of Long-Term Low-Thrust Small Satellites Orbiting Mars Christopher Swanson University of Florida 3131 NW 58th Blvd. Gainesville FL [email protected] Faculty Advisor: Riccardo Bevilacqua Graduate Advisor: Patrick Kelly University of Florida ABSTRACT As expansion of deep space missions continue, Mars is quickly becoming the planet of primary focus for science and exploratory satellite systems. Due to the cost of sending large satellites equipped with enough fuel to last the entire lifespan of a mission, inexpensive small satellites equipped with low thrust propulsion are of continuing interest. In this paper, a model is constructed for use in simulating the control of a small satellite system, equipped with low- thrust propulsion in orbit around Mars. The model takes into consideration the fuel consumption and can be used to predict the lifespan of the satellite based on fuel usage due to the natural perturbation of the orbit. This model of the natural perturbations around Mars implements a Lyapunov Based control law using a set of gains for the orbital elements calculated by obtaining the ratio of the instantaneous rate of change of the element over the maximum rate of change over the current orbit. The thrust model simulates the control of the semi-major axis, eccentricity, and inclination based upon a thrust vector fixed to a local vertical local horizontal reference frame of the satellite. This allows for a more efficient fuel usage over the long continuous thrust burns by prioritizing orbital elements in the control when they have a higher relative rate of change rather than the Lyapunov control prioritizing the element with the greatest magnitude rate of change. -
Stable Orbits in the Small-Body Problem an Application to the Psyche Mission
Stable Orbits in the Small-Body Problem An Application to the Psyche Mission Stijn Mast Technische Universiteit Delft Stable Orbits in the Small-Body Problem Front cover: Artist illustration of the Psyche mission spacecraft orbiting asteroid (16) Psyche. Image courtesy NASA/JPL-Caltech. ii Stable Orbits in the Small-Body Problem An Application to the Psyche Mission by Stijn Mast to obtain the degree of Master of Science at Delft University of Technology, to be defended publicly on Friday August 31, 2018 at 10:00 AM. Student number: 4279425 Project duration: October 2, 2017 – August 31, 2018 Thesis committee: Ir. R. Noomen, TU Delft Dr. D.M. Stam, TU Delft Ir. J.A. Melkert, TU Delft Advisers: Ir. R. Noomen, TU Delft Dr. J.A. Sims, JPL Dr. S. Eggl, JPL Dr. G. Lantoine, JPL An electronic version of this thesis is available at http://repository.tudelft.nl/. Stable Orbits in the Small-Body Problem This page is intentionally left blank. iv Stable Orbits in the Small-Body Problem Preface Before you lies the thesis Stable Orbits in the Small-Body Problem: An Application to the Psyche Mission. This thesis has been written to obtain the degree of Master of Science in Aerospace Engineering at Delft University of Technology. Its contents are intended for anyone with an interest in the field of spacecraft trajectory dynamics and stability in the vicinity of small celestial bodies. The methods presented in this thesis have been applied extensively to the Psyche mission. Conse- quently, the work can be of interest to anyone involved in the Psyche mission. -
Relative Orbital Motion Dynamical Models for Orbits About Nonspherical Bodies
Relative Orbital Motion Dynamical Models for Orbits about Nonspherical Bodies Item Type text; Electronic Thesis Authors Burnett, Ethan Ryan Publisher The University of Arizona. Rights Copyright © is held by the author. Digital access to this material is made possible by the University Libraries, University of Arizona. Further transmission, reproduction or presentation (such as public display or performance) of protected items is prohibited except with permission of the author. Download date 01/10/2021 07:16:25 Link to Item http://hdl.handle.net/10150/628098 Relative Orbital Motion Dynamical Models for Orbits about Nonspherical Bodies by Ethan R. Burnett Copyright c Ethan R. Burnett 2018 A Thesis Submitted to the Faculty of the Department of Aerospace and Mechanical Engineering In Partial Fulfillment of the Requirements For the Degree of Master of Science With a Major in Aerospace Engineering In the Graduate College The University of Arizona 2 0 1 8 2 3 Acknowledgments I would like to thank my advisor, Dr. Eric Butcher, for his encouragement and excellent teaching. I greatly enjoyed working with my research colleagues: Mohammad Maadani, Bharani Malladi, David Yaylali, Jingwei Wang, Morad Nazari, and Arman Dabiri. I thank my parents and my brothers for their steadfast support. 4 Vita Education Master of Science, Aerospace Engineering Emphasis: Dynamics and Control University of Arizona, Tucson, AZ May 2018 Bachelor of Science, Aerospace Engineering Minor, Physics University of Arizona, Tucson, AZ May 2016 Publications Ethan Burnett and Eric Butcher, \Linearized Relative Orbital Motion Dynamics in a Ro- tating Second Degree and Order Gravity Field," AAS 18-232, AAS/AIAA Astrodynamics Specialist Conference, Snowbird, UT, August 19 { 23, 2018 (submitted) Ethan Burnett, Eric Butcher, and T. -
Circular Orbit
Circular orbit A circular orbit is the orbit with a fixed distance around the barycenter, that is, in the shape of a circle. Below we consider a circular orbit in astrodynamics or celestial mechanics under standard assumptions. Here the centripetal force is the gravitational force, and the axis mentioned above is the line through the center of the central mass perpendicular to the plane of motion. In this case, not only the distance, but also the speed, angular speed, potential and kinetic energy are constant. There is no periapsis or apoapsis. This orbit has no radial version. A circular orbit is depicted in the top-left Contents quadrant of this diagram, where the Circular acceleration gravitational potential well of the central mass shows potential energy, and the Velocity kinetic energy of the orbital speed is Equation of motion shown in red. The height of the kinetic Angular speed and orbital period energy remains constant throughout the constant speed circular orbit. Energy Delta-v to reach a circular orbit Orbital velocity in general relativity Derivation See also Circular acceleration Transverse acceleration (perpendicular to velocity) causes change in direction. If it is constant in magnitude and changing in direction with the velocity, we get a circular motion. For this centripetal acceleration we have where: is orbital velocity of orbiting body, is radius of the circle is angular speed, measured in radians per unit time. The formula is dimensionless, describing a ratio true for all units of measure applied uniformly across the formula. If the numerical value of is measured in meters per second per second, then the numerical values for will be in meters per second, in meters, and in radians per second. -
Positional Astronomy : Earth Orbit Around Sun
www.myreaders.info www.myreaders.info Return to Website POSITIONAL ASTRONOMY : EARTH ORBIT AROUND SUN RC Chakraborty (Retd), Former Director, DRDO, Delhi & Visiting Professor, JUET, Guna, www.myreaders.info, [email protected], www.myreaders.info/html/orbital_mechanics.html, Revised Dec. 16, 2015 (This is Sec. 2, pp 33 - 56, of Orbital Mechanics - Model & Simulation Software (OM-MSS), Sec 1 to 10, pp 1 - 402.) OM-MSS Page 33 OM-MSS Section - 2 -------------------------------------------------------------------------------------------------------13 www.myreaders.info POSITIONAL ASTRONOMY : EARTH ORBIT AROUND SUN, ASTRONOMICAL EVENTS ANOMALIES, EQUINOXES, SOLSTICES, YEARS & SEASONS. Look at the Preliminaries about 'Positional Astronomy', before moving to the predictions of astronomical events. Definition : Positional Astronomy is measurement of Position and Motion of objects on celestial sphere seen at a particular time and location on Earth. Positional Astronomy, also called Spherical Astronomy, is a System of Coordinates. The Earth is our base from which we look into space. Earth orbits around Sun, counterclockwise, in an elliptical orbit once in every 365.26 days. Earth also spins in a counterclockwise direction on its axis once every day. This accounts for Sun, rise in East and set in West. Term 'Earth Rotation' refers to the spinning of planet earth on its axis. Term 'Earth Revolution' refers to orbital motion of the Earth around the Sun. Earth axis is tilted about 23.45 deg, with respect to the plane of its orbit, gives four seasons as Spring, Summer, Autumn and Winter. Moon and artificial Satellites also orbits around Earth, counterclockwise, in the same way as earth orbits around sun. Earth's Coordinate System : One way to describe a location on earth is Latitude & Longitude, which is fixed on the earth's surface. -
A Dedicated Earth-Orbiting Spacecraft for Investigating Gravitational Physics and the Space Environment
aerospace Article METRIC: A Dedicated Earth-Orbiting Spacecraft for Investigating Gravitational Physics and the Space Environment Roberto Peron 1,* ID and Enrico C. Lorenzini 2 1 Istituto di Astrofisica e Planetologia Spaziali (IAPS-INAF), 00133 Roma, Italy 2 Department of Industrial Engineering and CISAS “Giuseppe Colombo”, University of Padova, 35131 Padova, Italy; [email protected] * Correspondence: [email protected]; Tel.: +39-06-499-34-367 Academic Editor: Darren L. Hitt Received: 22 May 2017; Accepted: 13 July 2017; Published: 20 July 2017 Abstract: A dedicated mission in low Earth orbit is proposed to test predictions of gravitational interaction theories and to directly measure the atmospheric density in a relevant altitude range, as well as to provide a metrological platform able to tie different space geodesy techniques. The concept foresees a small spacecraft to be placed in a dawn-dusk eccentric orbit between 450 and 1200 km of altitude. The spacecraft will be tracked from the ground with high precision, and a three-axis accelerometer package on-board will measure the non-gravitational accelerations acting on its surface. Estimates of parameters related to fundamental physics and geophysics should be obtained by a precise orbit determination, while the accelerometer data will be instrumental in constraining the atmospheric density. Along with the mission scientific objectives, a conceptual configuration is described together with an analysis of the dynamical environment experienced by the spacecraft and the accelerometer. Keywords: fundamental physics; general relativity; space environment; atmospheric density; International Terrestrial Reference Frame; measurement of acceleration; satellite laser ranging; global navigation satellite systems PACS: 04.80.Cc; 06.20.-f; 06.30.Gv; 07.87.+v; 14.70.Kv; 91.10.Fc; 91.10.Sp; 91.10.Ws; 95.10.Eg; 95.40.+s 1. -
Lecture 7 Launch Trajectories
AA 284a Advanced Rocket Propulsion Lecture 7 Launch Trajectories Prepared by Arif Karabeyoglu Department of Aeronautics and Astronautics Stanford University and Mechanical Engineering KOC University Fall 2019 Stanford University AA284a Advanced Rocket Propulsion Orbital Mechanics - Review • Newton’s law of gravitation: mMG Fg = r2 – M, m: Mass of the bodies – r: Distance between the center of masses of the two bodies – Fg: Gravitational attraction force between the two bodies – G: Universal gravitational constant • Assume that m is the mass of the spacecraft and M is the mass of the celestial body. Arrange the force expression as (Note that m << M) m 2 Fg = = g m = M G g = r r2 • Here the gravitational parameter has been introduced for convenience. It is a constant for a given celestial mass. For Earth 3 2 = 398,600 km /sec • For circular orbit: centrifugal force balancing the gravitational force acting on the satellite – Orbital Velocity: V = co r r3 – Orbital Period: P = 2 Stanford University 2 Karabeyoglu AA284a Advanced Rocket Propulsion Orbital Mechanics - Review • Fundamental Assumptions: – Two body assumption • Motion of the spacecraft is only affected my a single central body – The mass of the spacecraft is negligible compared to the mass of the celestial body – The bodies are spherically symmetric with the masses concentrated at the center of the sphere – No forces other than gravity (and inertial forces) r r = k r3 • Solution: • Sections of a cone Stanford University 3 Karabeyoglu AA284a Advanced Rocket Propulsion Orbital Mechanics - Review • Solution: – Orbits of any conic section, elliptic, parabolic, hyperbolic – Energy is conserved in the conservative force field. -
Solar System Astronomy Notes
Solar System Astronomy Notes Victor Andersen University of Houston [email protected] May 19, 2006 Copyright c Victor Andersen 2006 Contents 1 The Sky and Time Keeping 2 2 The Historical Development of Astronomy 10 3 Modern Science 21 4 The Laws of Motion and Gravitation 23 5 Overview of the Solar System 31 6 The Terrestrial Planets 35 6.1 The Earth . 36 6.2 The Moon . 46 6.3 Mercury . 52 6.4 Venus . 55 6.5 Mars . 58 7 The Gas Giant Planets and Their Ring Systems 62 8 The Large Moons of the Gas Giants 71 9 Minor Solar System Bodies 76 10 The Outer Solar System 82 11 The Sun and Its Influence in the Solar System 85 12 Life Elsewhere in the Solar System 92 13 The Formation of the Solar System 95 1 Chapter 1 The Sky and Time Keeping Constellations • Constellations are areas on the sky, usually defined by distinctive group- ings of stars (Modern astronomers define 88, if you are interested a complete list is given in appendix 11 of the textbook). Constellations have boundaries in the same way that countries on the earth do. • Prominent groupings of stars that do not comprise an entire constel- lation are called asterisms (for example, the Big Dipper is just part of the constellation Ursa Majoris). Star Names • Most bright stars have names (usually Arabic). • Stars are also named using the Greek alphabet along with the name of the constellation of which the star is a member. The first five lower case letters of the Greek alphabet are: α (alpha), β (beta), γ (gamma), δ (delta), and (epsilon). -
Motions of Celestial Bodies: Computer Simulations
Educational Software MOTIONS OF CELESTIAL BODIES: COMPUTER SIMULATIONS Eugene I. Butikov Department of Physics St. Petersburg State University St. Petersburg, Russia 2014 ii Contents Preface vii 1 Introduction: Getting Started 1 1.1 List of the Simulation Programs . 1 1.2 How to Operate the Simulation Programs . 2 1.3 Keplerian Motions in Celestial Mechanics . 3 1.4 Numerical and Analytical Methods . 6 I Review of the Simulations 9 2 Kepler’s Laws 11 2.1 Kepler’s First Law . 11 2.2 Kepler’s Second Law . 16 2.3 Kepler’s Third Law . 19 2.4 The Approximate Nature of Kepler’s Laws . 23 3 Hodograph of the Velocity Vector 27 3.1 Hodograph of the Velocity for Closed Orbits . 27 3.2 Hodograph of the Velocity for Open Orbits . 29 4 Satellites and Missiles 33 4.1 Families of Keplerian Orbits . 34 4.1.1 Various Directions of the Initial Velocities . 34 4.1.2 Equal Magnitudes of the Initial Velocities . 38 4.1.3 Different Magnitudes of the Initial Velocities . 40 4.2 Evolution of an Orbit in the Atmosphere . 43 4.2.1 Evolution of an Elongated Elliptical Orbit . 43 4.2.2 Late Stage of the Evolution and Aerodynamical Paradox . 45 4.2.3 Air Density over the Earth . 47 5 Active Maneuvers in Space Orbits 51 5.1 How to Operate the Program . 51 5.2 Space Flights and Orbital Maneuvers . 54 5.2.1 Designing a Space Flight . 54 iii iv CONTENTS 5.2.2 Way Back from Space to the Earth . 55 5.3 Relative Orbital Motion . -
Visualizing Solutions of the Circular Restricted Three
VISUALIZING SOLUTIONS OF THE CIRCULAR RESTRICTED THREE-BODY PROBLEM by NKOSI NATHAN TRIM A thesis submitted to the Graduate School-Camden Rutgers, The State University of New Jersey in partial fulfillment of the requirements for the degree of Master of Science Graduate Program in Mathematical Sciences written under the direction of Dr. Joseph L. Gerver and approved by Dr. Haydee Herrera-Guzman Camden, New Jersey May 2009 ABSTRACT OF THE THESIS by NKOSI NATHAN TRIM Thesis Director: Dr. Joseph L. Gerver The stability of a satellite near the Lagrange points is studied in a Circular Restricted Three-Body Problem (CR3BP). The Runge Kutta method is used to trace out the orbital path of the satellite over a period of time. Various initial positions near the Lagrange points and velocities are used to produce various paths the satellite can take. The primary paths focused on are horseshoe paths. Horseshoe orbits are shown to be sometimes stable and sometimes chaotic. ii Table of contents ABSTRACTOFTHETHESIS................................ ii Tableofcontents .................................... iii Listofillustrations................................ ........ iv Listoftables ....................................... .... vi 1 Introduction...................................... ... 1 2 LagrangePoints .................................... .. 4 2.1 Formulating the Restricted Three Body Problem . ........ 4 2.2 EstablishingStability . .. .. .. .. .. .. .. .. .. .. .. .... 7 3 RungeKuttaMethod .................................. 8 4 SolutionfortheLagrangePoints