HD 174884: a Strongly Eccentric, Short-Period Early-Type Binary System Discovered by Corot
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Astrodynamics
Politecnico di Torino SEEDS SpacE Exploration and Development Systems Astrodynamics II Edition 2006 - 07 - Ver. 2.0.1 Author: Guido Colasurdo Dipartimento di Energetica Teacher: Giulio Avanzini Dipartimento di Ingegneria Aeronautica e Spaziale e-mail: [email protected] Contents 1 Two–Body Orbital Mechanics 1 1.1 BirthofAstrodynamics: Kepler’sLaws. ......... 1 1.2 Newton’sLawsofMotion ............................ ... 2 1.3 Newton’s Law of Universal Gravitation . ......... 3 1.4 The n–BodyProblem ................................. 4 1.5 Equation of Motion in the Two-Body Problem . ....... 5 1.6 PotentialEnergy ................................. ... 6 1.7 ConstantsoftheMotion . .. .. .. .. .. .. .. .. .... 7 1.8 TrajectoryEquation .............................. .... 8 1.9 ConicSections ................................... 8 1.10 Relating Energy and Semi-major Axis . ........ 9 2 Two-Dimensional Analysis of Motion 11 2.1 ReferenceFrames................................. 11 2.2 Velocity and acceleration components . ......... 12 2.3 First-Order Scalar Equations of Motion . ......... 12 2.4 PerifocalReferenceFrame . ...... 13 2.5 FlightPathAngle ................................. 14 2.6 EllipticalOrbits................................ ..... 15 2.6.1 Geometry of an Elliptical Orbit . ..... 15 2.6.2 Period of an Elliptical Orbit . ..... 16 2.7 Time–of–Flight on the Elliptical Orbit . .......... 16 2.8 Extensiontohyperbolaandparabola. ........ 18 2.9 Circular and Escape Velocity, Hyperbolic Excess Speed . .............. 18 2.10 CosmicVelocities -
Orbital Mechanics of Gravitational Slingshots 1 Introduction 2 Approach
Orbital Mechanics of Gravitational Slingshots Final Paper 15-424: Foundations of Cyber-Physical Systems Adam Moran, [email protected] John Mann, [email protected] May 1, 2016 Abstract A gravitational slingshot is a maneuver to save fuel by using the gravity of a planet to accelerate or decelerate a spacecraft. Due to the large distances and high speeds involved, slingshots require precise accuracy to accomplish | the slightest mistake could cause the whole mission to fail. Therefore, we have developed a cyber-physical system to model the physics and prove the safety and efficiency of powered and unpowered gravitational slingshots. We present our findings and proof in this paper. 1 Introduction A gravitational slingshot is a maneuver performed to increase or decrease the speed of a spacecraft by simply approaching planetary bodies. A spacecraft's usefulness and maneuverability is basically tied to the amount of fuel it can carry, and the more fuel a spacecraft holds, the more fuel it needs to carry that fuel into orbit. Therefore, gravitational slingshots are a very appealing way to save mass, and therefore money, on deep-space missions since these maneuvers do not require any fuel. As missions conducted by national and private space programs become more frequent and ambitious, the need for these precise maneuvers will increase. Therefore, we have created a cyber-physical system that models the physics of a gravitational slingshot for a spacecraft approaching a planet. In the "Approach" section of this paper, we give a brief overview of the physics involved with orbits and gravitational slingshots. In the "Models and Properties" section of this paper, we describe what assumptions and simplifications we made to model these astrophysics in a way for us to prove our desired properties with KeYmaeraX. -
P.Roceedings
Source of Acquisiti on NASA Contractor/Grantee .P .ROCEEDINGS , . -. -. fi.t.... ~ I I I -- -_ .. - T & D [] 0 FIFTH 0 0 S T A T IISPACEI E S CONGRESS COCOA BEACH, FLORIDA I - , , ',- MARCH 11,12,13,14,1968 ~ / 5i: .3 1/, ! ..?t~b , .~ I - , I FOREWORD Each' spring the Canaveral Council of Technical Societies sponsors a symposium devoted to the' accomplishments of oU,r space programs and plans for future activities. These pro ceedings provide a permanent record of the papers presented at our Fifth Space Congress held in Cocoa Beach, Florida, March 11 - 14, 1968. The Fifth Space Congress theme, "Our Goals in Space Operations ~ ', was, chosen to provide a forum for engineers and scientists to express individual and corporate views on where our nation should be heading in space operation. The papers presented herein depict the broad and varied views of the industrial organizations and government agencies involved in space activities. We believe that these proceedings will provide technical stimulation and serve as a valuable reference for the scientists and engineers working in our space program. On behalf of the Canaveral Council of Teclmical SOCieties, I wish to express our appreciation to the authors who prepared and presented papers at the Fifth Space Congress. M~/?~~~- EDWARDP. W~ General Chairman Fifth Space Congress • I I TABLE OF CONTENTS (Continued) Session Pages Elliptic Capture Orbits for Missions to the Near Planets by F . G. Casal/B. L. Swenson/ A. C. Mascy of Moffett Field, Calif . .............................................................................. 25.3.1-8 A Venus Lander Probe for Manned Fly-By Missions by P. -
1 CHAPTER 10 COMPUTATION of an EPHEMERIS 10.1 Introduction
1 CHAPTER 10 COMPUTATION OF AN EPHEMERIS 10.1 Introduction The entire enterprise of determining the orbits of planets, asteroids and comets is quite a large one, involving several stages. New asteroids and comets have to be searched for and discovered. Known bodies have to be found, which may be relatively easy if they have been frequently observed, or rather more difficult if they have not been observed for several years. Once located, images have to be obtained, and these have to be measured and the measurements converted to usable data, namely right ascension and declination. From the available observations, the orbit of the body has to be determined; in particular we have to determine the orbital elements , a set of parameters that describe the orbit. For a new body, one determines preliminary elements from the initial few observations that have been obtained. As more observations are accumulated, so will the calculated preliminary elements. After all observations (at least for a single opposition) have been obtained and no further observations are expected at that opposition, a definitive orbit can be computed. Whether one uses the preliminary orbit or the definitive orbit, one then has to compute an ephemeris (plural: ephemerides ); that is to say a day-to-day prediction of its position (right ascension and declination) in the sky. Calculating an ephemeris from the orbital elements is the subject of this chapter. Determining the orbital elements from the observations is a rather more difficult calculation, and will be the subject of a later chapter. 10.2 Elements of an Elliptic Orbit Six numbers are necessary and sufficient to describe an elliptic orbit in three dimensions. -
Arxiv:2002.01920V4 [Astro-Ph.HE] 4 Apr 2020 the CHIME Collaboration Announced a P = 16 Day Periodicity from FRB 180916.J0158+65 (The CHIME/FRB Collaboration Et Al
FRB-periodicity: mild pulsars in tight O/B-star binaries Maxim Lyutikov1, Maxim V. Barkov1;2, Dimitrios Giannios1 1 Department of Physics, Purdue University, 525 Northwestern Avenue, West Lafayette, IN 47907-2036, USA 2 Astrophysical Big Bang Laboratory, RIKEN, 2-1 Hirosawa, Wako, Saitama 351-0198, Japan Received/Accepted ABSTRACT Periodicities observed in two Fast Radio Burst (FRB) sources (16 days in FRB 180916.J0158+65 and 160 days in FRB 121102) are consistent with that of tight, stellar mass binary systems. In the case of FRB 180916.J0158+65 the primary is an early −8 −7 −1 OB-type star with mass loss rate M_ ∼ 10 − 10 M yr , and the secondary a neutron star. The observed periodicity is not intrinsic to the FRB's source, but is due to the orbital phase-dependent modulation of the absorption conditions in the massive star's wind. The observed relatively narrow FRB activity window implies that the primary's wind dynamically dominates that of the pulsar, η = Lsd=(Mv_ wc) ≤ 1, where Lsd is pulsar spin-down, M_ is the primary's wind mass loss rate and vw is its velocity. 37 −1 The condition η ≤ 1 requires mildly powerful pulsar with Lsd . 10 erg s . The observations are consistent with magnetically-powered radio emission originating in the magnetospheres of young, strongly magnetized neutron stars, the classical magnetars. Subject headings: stars: binaries: general { stars: magnetars | stars: winds, outflows | fast radio bursts 1. FRB periodicity due to the orbital motion in O/B-NS binary 1.1. Observations and outline of the model arXiv:2002.01920v4 [astro-ph.HE] 4 Apr 2020 The CHIME collaboration announced a P = 16 day periodicity from FRB 180916.J0158+65 (The CHIME/FRB Collaboration et al. -
Bi-Elliptic Hohmann Transfer and One Tangent Burn Transfer Calculations Using the Monte Carlo Simulation
ISSN 2394-7349 International Journal of Novel Research in Engineering and Science Vol. 4, Issue 1, pp: (27-35), Month: March 2017 - August 2017, Available at: www.noveltyjournals.com Bi-Elliptic Hohmann Transfer and One Tangent Burn Transfer Calculations Using the Monte Carlo Simulation Mohamed Abdel M. Allam1, M. E. Awad2, Ibrahim Amin I.3 1Egyptian Armed Forces, Egypt 2Astronomy and Space Science Dept., Faculty of Science, Cairo University, Egypt 3Military Technical College (MTC), Egypt Abstract: One of the objectives in the orbit transfer problem is to achieve the optimal time of flight and the fuel consumption for the orbital transfer maneuver between two orbits. The transfer of satellites in too high orbits as geosynchronous one (geostationary), usually is achieved firstly by launching the satellite in Low Earth Orbit (LEO) (Parking orbit), then in elliptical transfer orbit and finally to the final orbit (Working orbit). In this paper, the Monte Carlo Simulation will be used to determine the optimum three tangent impulses maneuver (Bi-Elliptic Hohmann transfer) and determine the optimum One Tangent Burn transfer. From respective simulation, determine the optimum altitude of the transfer tangent point for Bi-Elliptic Hohmann transfer and determine the optimum angle true anomaly (υtrans) of the One Tangent Burn transfer to create minimum change of velocity and optimum time of flight for transfer and with minimum fuel consumption of this transfer. Keywords: Bi-Elliptic Hohmann Transfer; Coplanar Impulsive Maneuver; Monte Carlo Simulation; One Tangent Burn Transfer; Satellite Orbit. 1. INTRODUCTION R. H. Goddard (1919) was one of the first researchers on the problem of optimal transfers of a spacecraft between two points who suggested optimal approximate solutions for the problem of sending a rocket to high altitudes with minimum fuel consumption [1].After that, there is the very important work done by Hohmann (1925) who solved the problem of minimum ∆V transfers between two circular coplanar orbits. -
AFS 6 Orbit Design Process: 5
AFS 6 Orbit design process: 5. Mission Orbit Design Trades - A New element: Re-usability of launcher - Space-X, a US company started and owned by Elon Musk has a rocket called Falcon-9. It currently cost 54M$ - It has succeeded with a number of launches but the idea is to recover the first stage Exercise: Find informaon on the net and do: 1. Find the es4mates for the resusability, i.e. How many 4mes can a 1:st stage be launched? How many launches will an engine survive? What are the poten4al savings when regular landings of stage 1 is acheived? 2. Write a short text where you discuss the poten4al impact on scien4fic missions. Will there be more? Will the cost savings on the launcher impact the cost of the payload? Keplerian orbit transfers Single impulsive manoeuvre: Initial and final orbit intersect at the point of impulse. Single manoeuvres can only transfer s/c between intersecting orbits! Simplest case is for a co-planar transfer from circular orbit to elliptic orbit. If velocity is increased è semi- major axis will be expanded At least two manoeuvres are required to transfer s/c between non-intersecting orbits Hohmann transfer orbit è two coplanar manoevres to go from one small circular orbit to a large one Simple calculation accordingly 1 2 µ µ V − = ε ε = − 2 r 2a € € Planet a(AU)transfer T(years) ΔV (km/s) Mercury 0.847 0.289 5.6 Venus 0.931 0.400 3.5 Mars 1.131 0.709 3.6 Jupiter 2.051 2.731 6.3 Saturn 3.137 6.056 7.3 Uranus 5.534 15.972 8.0 Neptune 8.253 30.529 8.3 Pluto 10.572 45.208 8.4 Help is on the way: Gravity assist Specific -
Three Solutions to the Two- Body Problem
Degree project Three solutions to the two- body problem Author: Frida Gleisner Supervisor: Hans Frisk Examiner: Hans Frisk Date: 2013-06-18 Subject: Mathematics Level: Bachelor Department Of Mathematics Contents 1 Introduction 1 2 Different approaches to the two-body problem 2 2.1 Modern solution . .3 2.2 Newton’s solution . .3 2.3 Feynman’s solution . .3 2.4 Table of used assumptions and of what is proved . .3 3 Properties of the ellipse 4 4 A modern solution 7 4.1 One moving body . .7 4.1.1 Calculating the acceleration of r ....................8 4.1.2 The acceleration along '^ ........................9 4.1.3 The acceleration along r^ .........................9 4.1.4 Expressing r as a function of ' ..................... 10 4.2 Two moving bodies . 12 4.2.1 The motion around the center of mass is planar . 13 4.2.2 The method of reduced mass . 15 4.2.3 Finding equations for the locations of the bodies . 16 5 Newton’s solution 17 5.1 One moving body . 17 5.1.1 Proposition I . 17 5.1.2 Proposition VI . 17 5.1.3 Proposition XI . 18 5.2 Two moving bodies . 23 6 Feynman’s solution 25 6.1 One moving body . 25 7 Discussion 31 Appendices 33 A A modern approach of proving the inverse-square law using the equation of the ellipse 33 B Initial value problem, assuming one body in orbit 34 C Initial value problem, two bodies in orbit 36 D The motion of r1 and r2 is parallel 38 E Proposition LXI 39 i Three solutions to the two-body problem Frida Gleisner June 18, 2013 Abstract The two-body problem consists of determining the motion of two gravitationally interacting bodies with given masses and initial velocities. -
An Arc-Length Approximation for Elliptical Orbits
MNRAS 000,1{5 (2019) Preprint 26 November 2019 Compiled using MNRAS LATEX style file v3.0 AN ARC-LENGTH APPROXIMATION FOR ELLIPTICAL ORBITS Ashim B. Karki1? and Aayush Jha2y 1Wolfram Research, Champaign, IL, USA 2Department of Physics, St. Xavier's College, Maitighar, Kathmandu 44600, Nepal Accepted XXX. Received YYY; in original form ZZZ ABSTRACT In this paper, we overlay a continuum of analytical relations which essentially serve to compute the arc-length described by a celestial body in an elliptic orbit within a stipulated time interval. The formalism is based upon a two-dimensional heliocen- tric coordinate frame, where both the coordinates are parameterized as two infinitely differentiable functions in time by using the Lagrange inversion theorem. The parame- terization is firstly endorsed to generate a dynamically consistent ephemerides for any celestial object in an elliptic orbit, and thereafter manifested into a numerical integra- tion routine to approximate the arc-lengths delineated within an arbitrary interval of time. As elucidated, the presented formalism can also be orchestrated to quantify the perimeters of elliptic orbits of celestial bodies solely based upon their orbital period and other intrinsic characteristics. Key words: Methods: Numerical { Celestial Mechanics { Ephemerides 1 INTRODUCTION tations, and librations of the planets) as the need for more accurate planetary ephemerides became evident with the ad- Over the past few decades, the advancement in astrodynam- vancement in space exploration, and a plethora of other rec- ics has instigated a new era of celestial ephemerides. With tifications have been made in this arena hitherto. the advent of novel observational routines - namely, Mi- Notwithstanding the meticulously stringent observations crowave ranging, lunar laser ranging, VLBI measurements, made till date, the contemporary ephemerides are mostly etc. -
Orbital Mechanics, Oxford University Press
Astrodynamics (AERO0024) 5A. Orbital Maneuvers Gaëtan Kerschen Space Structures & Systems Lab (S3L) Course Outline THEMATIC UNIT 1: ORBITAL DYNAMICS Lecture 02: The Two-Body Problem Lecture 03: The Orbit in Space and Time Lecture 04: Non-Keplerian Motion THEMATIC UNIT 2: ORBIT CONTROL Lecture 05: Orbital Maneuvers Lecture 06: Interplanetary Trajectories 2 Definition of Orbital Maneuvering It encompasses all orbital changes after insertion required to place a satellite in the desired orbit. This lecture focuses on satellites in Earth orbit. 3 Motivation Without maneuvers, satellites could not go beyond the close vicinity of Earth. For instance, a GEO spacecraft is usually placed on a transfer orbit (LEO or GTO). 4 5. Orbital Maneuvers 5.1 Introduction 5.2 Coplanar maneuvers 5 5. Orbital Maneuvers 5.1 Introduction 5.1.1 Why ? 5.1.2 How ? 5.1.3 How much ? 5.1.4 When ? 6 Orbit Circularization Ariane V is able to place heavy GEO satellites in GTO: perigee: 200-650 km GTO apogee: ~35786 km. GEO 5.1.1 Why ? 7 Orbit Raising: Reboost ISS reboost due to atmospheric drag (ISS, Shuttle, Progress, ATV). The Space Shuttle is able to place heavy GEO satellites in near-circular LEO with a few hundred kilometers altitude. 5.1.1 Why ? 8 Orbit Raising: Evasive Maneuvers See also www.esa.int/SPECIALS/Operations/SEM64X0SAKF_0.html 5.1.1 Why ? 9 Orbit Raising: Deorbiting GEO Satellites Graveyard orbit: to eliminate collision risk, satellites should be moved out of the GEO ring at the end of their mission. Their orbit should be raised by about 300 km to avoid future interference with active GEO spacecraft. -
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. -
Orbit Calculation and Re-Entry Control of Vorsat Satellite
FACULDADE DE ENGENHARIA DA UNIVERSIDADE DO PORTO Orbit Calculation and Re-Entry Control of VORSat Satellite Cristiana Monteiro Silva Ramos PROVISIONAL VERSION Master in Electrical and Computers Engineering Supervisor: Sérgio Reis Cunha (PhD) June 2011 Resumo Desde o lançamento do primeiro satélite artificial para o espaço, Sputnik, mais de 30 000 satélites foram desenvolvidos e foram lançados posteriormente. No entanto, por mais autonomia que o satélite possa alcançar, nenhum satélite construido pelo Homem teria valor se não fosse possível localiza-lo e comunicar com ele. Os olhos humanos foram os primeiros recursos de que os seres humanos dispuseram para observar o Universo. O telescópio surgiu depois e foi durante séculos o único instrumento de exploração do Espaço. Hoje em dia, é possível localizar e seguir o trajecto que o satélite faz no espaço através de programas computacionais que localizam o satélite num dado instante e fazem a previsão do cálculo da órbita. Esta dissertação descreve o desenvolvimento de um sistema de navegação enquadrado no caso de estudo do satélite VORSat. O VORSat é um programa de satélite em desenvolvido na Faculdade de Engenharia da Universidade do Porto (FEUP), Potugal. Este projecto refere-se à construção de um CubeSat, de nome GAMA-Sat, e à reentrada de uma cápsula na Terra (ERC). Os objectivos consistem em determinar os parâmetros que descrevem uma órbita num dado instante (elementos de Kepler) e em calcular as posições futuras do satélite através de observações iterativas, minimizando o erro associado a cada observação. Também se pretende controlar, através de variações do termo de arrastamento, o local de reentrada.