DOCSLIB.ORG

Evolution of Spacecraft Orbital Motion Due to Hypervelocity Impacts with Debris and Meteoroids

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Evolution of Spacecraft Orbital Motion Due to Hypervelocity Impacts with Debris and Meteoroids EVOLUTION OF SPACECRAFT ORBITAL MOTION DUE TO HYPERVELOCITY IMPACTS WITH DEBRIS AND METEOROIDS Luc Sagnieres` and Inna Sharf Department of Mechanical Engineering, McGill University, 817 Sherbrooke W., Montreal, QC, H3A 0C3, Canada, Email: [email protected] ABSTRACT From a somewhat different perspective, the impacts be- tween the small debris flux and a large object, be it a functional or defunct spacecraft, will lead to a transfer A framework is developed to include the effect of bom- of momentum from the impactor to the target, thus mod- bardment from space debris and meteoroids on the orbital ifying its orbital parameters. Moreover, at such high ve- motion of active and inoperative satellites by considering locities, the particles ejected during crater formation pro- the transfer of momentum from impactors as a stochastic vide an additional momentum transfer, an effect known process known as a compound Poisson process. The dif- as momentum enhancement [12]. While many of the en- ferential equation for orbital motion is set up as a stochas- vironmental factors influencing an object’s orbit in space tic differential equation and solved in a Monte Carlo sim- have been extensively studied, the effect of bombardment ulation. The orbital propagation is coupled to a determin- of debris and meteoroids on spacecraft orbit propagation, istic attitude propagation model in order to implement the on the other hand, is a research area that is still in its in- additional effect of ejecta momentum. The framework is fancy. The work proposed here follows the methodology applied to four spacecraft and two analytical test cases described in [14] for propagating the attitude motion of are presented in order to obtain a confirmation of the va- tumbling spacecraft undergoing these hypervelocity im- lidity of the stochastic model. The resulting probability pacts, however, applying that approach to orbital motion. distribution functions for the change in semi-major axis after a ten-year propagation period show that for the pre- This method enables the inclusion of hypervelocity im- sented cases, a systematic decrease is present, although pacts into spacecraft orbit propagation models by consid- small compared to observations and estimated decay due ering the transfer of linear momentum from collisions as to other effects. a stochastic jump process known as a compound Pois- son process [14]. The differential equation for orbital motion then becomes a stochastic differential equation Key words: hypervelocity impacts, orbit propagation, (SDE) and is solved in a Monte Carlo simulation by em- stochastic differential equation. ploying independent sets of randomly generated colli- sions [14]. These collisions can be obtained using impact fluxes converted into probability density functions (PDF), 1. INTRODUCTION the former obtained from the European Space Agency’s (ESA) Meteoroid and Space Debris Terrestrial Environ- ment Reference (MASTER) model describing the debris Collisions between small orbital debris and micromete- and meteoroid population around Earth [5]. Furthermore, oroids and active and inoperative satellites are typically the stochastic orbit propagation model can be coupled to a considered in the context of the increasing threat that deterministic attitude propagation model in order to ade- they present to space operations and missions [8]. The quately implement the additional effect of ejecta momen- current estimates of the space debris population in or- tum which is dependent on the orientation of the space- bit are 29,000 above 10 cm, 670,000 above 1 cm, and craft. The momentum contribution from ejecta can be more than 170 million above 1 mm [4]. Furthermore, ev- calculated for every collision by considering the velocity, ery day approximately 100 tons of dust and sand-sized mass and direction distribution of the ejecting particles meteoroids enter Earth’s atmosphere [9]. When collid- from a model developed for ESA defining the character- ing with orbiting spacecraft, these hypervelocity impacts, istics of such ejecta [12]. occurring at speeds up to 60 km per second, can have disastrous consequences because of the structural dam- First, a review of the differential equations governing age they may cause and electromagnetic interference they spacecraft dynamics will be presented in Section 2. Sec- may have with spacecraft systems [3]. When impacting ond, the SDEs for orbital motion will be discussed in large debris, this bombardment is responsible for gener- Section 3 along with the numerical integration method ating more small debris in orbit, further aggravating the used to solve them. In order to assess the importance space debris problem. of these collisions on orbit propagation, the developed Proc. 7th European Conference on Space Debris, Darmstadt, Germany, 18–21 April 2017, published by the ESA Space Debris Office Ed. T. Flohrer & F. Schmitz, (http://spacedebris2017.sdo.esoc.esa.int, June 2017) model is then applied to several pieces of orbital ob- 3. INCLUDING HYPERVELOCITY IMPACTS IN jects in Section 4: the defunct European environmen- ORBITAL MOTION tal satellite Envisat, representing a potential active de- bris removal (ADR) target in low-Earth orbit (LEO); a high area-to-mass ratio (HAMR) object modeled after As hypervelocity impacts are random, they will have to multi-layer insulation debris in a highly elliptic orbit with be considered as a stochastic process and the differen- apogee at geostationary altitude, chosen to represent an tial equation for orbital motion will have to be set up as upper bound when the effect of the collisions is strongest; an SDE. A stochastic process is one that describes the one of the operational LAGEOS spherical satellites in evolution of a system’s random variables over time, and, a medium Earth orbit (MEO), satellites having been the unlike deterministic processes, can expand in multiple, cause of some debate due to unexplained semi-major axis or infinitely many, paths [11]. Including stochastic pro- decay; and lastly to the Japenese satellite Ajisai, also an cesses in differential equations is useful when wanting operational spherical satellite undergoing an orbital de- to describe uncertainties or random variations in any dy- cay due to aerodynamic drag. Finally, in Section 5, two namical system. Solving SDEs requires the use of numer- analytical test cases are presented to provide additional ical integration techniques that have been a subject of in- confidence in the validity of the propagation model. terest in the last few decades as most integration methods used for solving ordinary differential equations perform poorly when applied to SDEs [11]. 2. SPACECRAFT DYNAMICS 3.1. Collisions as a compound Poisson process Understanding the dynamics of satellites in orbit requires analysis of the six degrees of freedom governing the or- As explained in [14], the momentum transfer due to hy- bit and attitude of the spacecraft: three for translation pervelocity impacts can be represented as a compound and three for orientation. Three vector differential equa- Poisson process if a PDF describing the impacts can be tions describe the evolution of the corresponding vari- determined, and if the time between collisions can be ables. First is the dynamics equation for orbital motion shown to follow an exponential distribution. PDFs for in an Earth-centered inertial (ECI) coordinate frame [2]: the mass and velocity of the impactors can be obtained µ X by transforming the velocity and mass impact fluxes of ¨r(t) = − r(t) + a (t; r(t)) (1) r(t)3 j the MASTER-2009 model for a certain input orbit. A 2D j PDF for direction can also be obtained in an analogous where r is the position as a function of time t, r = krk, manner from the directional impact flux as a function of µ is the Earth’s gravitational parameter, and ai represents azimuth and elevation angles. These PDFs can then be the non-gravitational accelerations, and assuming a per- combined to define the linear momentum of an impactor. fectly uniform spherical Earth, ignoring Earth’s oblate- ness and third-body interactions. The time between collisions can be assumed to follow an exponential distribution due to the analogy between ki- The second differential equation, the dynamics equation netic gas theory and objects sweeping the space environ- for attitude motion, relates the evolution of angular veloc- ment filled with other debris and meteoroids [5]. There- ity, !, to the sum of the external torques about the center fore, the PDF for time between collisions can be obtained of mass of the body, τ j [6]: from the total expected number of collisions as outputted by MASTER-2009. The four PDFs obtained then define × X I!_ (t) + !(t) I!(t) = τ j(t; q(t)) (2) the impactors that a spacecraft in the specified orbit will j collide with. where I is the matrix representation of the inertia ten- From Eq. (1), the differential equation of motion can be sor of the rigid body in the centroidal body-fixed frame set up in differential form: T T and q = q0 qv is the attitude parametrization, cho- 0 1 sen here to be quaternions. µ X d_r(t) = − r(t) + a (t; r(t)) dt (5) @ r(t)3 j A Finally, the third differential equation is the kinematic j equation for orientation: Considering the momentum of the satellite, defined as 1 _q(t) = ΩΩ(!)q(t) (3) p = M _r with M the mass of the spacecraft, Eq. (5) can 2 be rewritten as: where, expressed in terms of the body-referenced compo- 0 1 µ X nents of !: dp(t) = M − r(t) + a (t; r(t)) dt (6) @ r(t)3 j A 2 0 −!x −!y −!z3 j ! 0 ! −! Ω = 6 x z y7 (4) 4!y −!z 0 !x 5 Including the term for the momentum from hypervelocity !z !y −!x 0 impacts as a function of the compound Poisson process, Yt, Eq.
Load more

Recommended publications
  • PDF (03Ragozzine Exo-Interiors.Pdf)
    20 Chapter 2 Probing the Interiors of Very Hot Jupiters Using Transit Light Curves This chapter will be published in its entirety under the same title by authors D. Ragozzine and A. S. Wolf in the Astrophysical Journal, 2009. Reproduced by permission of the American Astro- nomical Society. 21 Abstract Accurately understanding the interior structure of extra-solar planets is critical for inferring their formation and evolution. The internal density distribution of a planet has a direct effect on the star-planet orbit through the gravitational quadrupole field created by the rotational and tidal bulges. These quadrupoles induce apsidal precession that is proportional to the planetary Love number (k2p, twice the apsidal motion constant), a bulk physical characteristic of the planet that depends on the internal density distribution, including the presence or absence of a massive solid core. We find that the quadrupole of the planetary tidal bulge is the dominant source of apsidal precession for very hot Jupiters (a . 0:025 AU), exceeding the effects of general relativity and the stellar quadrupole by more than an order of magnitude. For the shortest-period planets, the planetary interior induces precession of a few degrees per year. By investigating the full photometric signal of apsidal precession, we find that changes in transit shapes are much more important than transit timing variations. With its long baseline of ultra-precise photometry, the space-based Kepler mission can realistically detect apsidal precession with the accuracy necessary to infer the presence or absence of a massive core in very hot Jupiters with orbital eccentricities as low as e ' 0:003.
    [Show full text]
  • Orbital Lifetime Predictions
    Orbital LIFETIME PREDICTIONS An ASSESSMENT OF model-based BALLISTIC COEFfiCIENT ESTIMATIONS AND ADJUSTMENT FOR TEMPORAL DRAG co- EFfiCIENT VARIATIONS M.R. HaneVEER MSc Thesis Aerospace Engineering Orbital lifetime predictions An assessment of model-based ballistic coecient estimations and adjustment for temporal drag coecient variations by M.R. Haneveer to obtain the degree of Master of Science at the Delft University of Technology, to be defended publicly on Thursday June 1, 2017 at 14:00 PM. Student number: 4077334 Project duration: September 1, 2016 – June 1, 2017 Thesis committee: Dr. ir. E. N. Doornbos, TU Delft, supervisor Dr. ir. E. J. O. Schrama, TU Delft ir. K. J. Cowan MBA TU Delft An electronic version of this thesis is available at http://repository.tudelft.nl/. Summary Objects in Low Earth Orbit (LEO) experience low levels of drag due to the interaction with the outer layers of Earth’s atmosphere. The atmospheric drag reduces the velocity of the object, resulting in a gradual decrease in altitude. With each decayed kilometer the object enters denser portions of the atmosphere accelerating the orbit decay until eventually the object cannot sustain a stable orbit anymore and either crashes onto Earth’s surface or burns up in its atmosphere. The capability of predicting the time an object stays in orbit, whether that object is space junk or a satellite, allows for an estimate of its orbital lifetime - an estimate satellite op- erators work with to schedule science missions and commercial services, as well as use to prove compliance with international agreements stating no passively controlled object is to orbit in LEO longer than 25 years.
    [Show full text]
  • NASA Process for Limiting Orbital Debris
    NASA-HANDBOOK NASA HANDBOOK 8719.14 National Aeronautics and Space Administration Approved: 2008-07-30 Washington, DC 20546 Expiration Date: 2013-07-30 HANDBOOK FOR LIMITING ORBITAL DEBRIS Measurement System Identification: Metric APPROVED FOR PUBLIC RELEASE – DISTRIBUTION IS UNLIMITED NASA-Handbook 8719.14 This page intentionally left blank. Page 2 of 174 NASA-Handbook 8719.14 DOCUMENT HISTORY LOG Status Document Approval Date Description Revision Baseline 2008-07-30 Initial Release Page 3 of 174 NASA-Handbook 8719.14 This page intentionally left blank. Page 4 of 174 NASA-Handbook 8719.14 This page intentionally left blank. Page 6 of 174 NASA-Handbook 8719.14 TABLE OF CONTENTS 1 SCOPE...........................................................................................................................13 1.1 Purpose................................................................................................................................ 13 1.2 Applicability ....................................................................................................................... 13 2 APPLICABLE AND REFERENCE DOCUMENTS................................................14 3 ACRONYMS AND DEFINITIONS ...........................................................................15 3.1 Acronyms............................................................................................................................ 15 3.2 Definitions .........................................................................................................................
    [Show full text]
  • NEAR of the 21 Lunar Landings, 19—All of the U.S
    Copyrights Prof Marko Popovic 2021 NEAR Of the 21 lunar landings, 19—all of the U.S. and Russian landings—occurred between 1966 and 1976. Then humanity took a 37-year break from landing on the moon before China achieved its first lunar touchdown in 2013. Take a look at the first 21 successful lunar landings on this interactive map https://www.smithsonianmag.com/science-nature/interactive-map-shows-all-21-successful-moon-landings-180972687/ Moon 1 The near side of the Moon with the major maria (singular mare, vocalized mar-ray) and lunar craters identified. Maria means "seas" in Latin. The maria are basaltic lava plains: i.e., the frozen seas of lava from lava flows. The maria cover ∼ 16% (30%) of the lunar surface (near side). Light areas are Lunar Highlands exhibiting more impact craters than Maria. The far side is pocked by ancient craters, mountains and rugged terrain, largely devoid of the smooth maria we see on the near side. The Lunar Reconnaissance Orbiter Moon 2 is a NASA robotic spacecraft currently orbiting the Moon in an eccentric polar mapping orbit. LRO data is essential for planning NASA's future human and robotic missions to the Moon. Launch date: June 18, 2009 Orbital period: 2 hours Orbit height: 31 mi Speed on orbit: 0.9942 miles/s Cost: 504 million USD (2009) The Moon is covered with a gently rolling layer of powdery soil with scattered rocks called the regolith; it is made from debris blasted out of the Lunar craters by the meteor impacts that created them.
    [Show full text]
  • Study of Extraterrestrial Disposal of Radioactive Wastes
    NASA TECHNICAL NASA TM X-71557 MEMORANDUM N--- XI NASA-TM-X-71557) STUDY OF N74-22776 EXTRATERRESTRIAL DISPOSAL OF READIOACTIVE WASTES. PART 1: SPACE TRANSPORTATION. AN.. DESTINATION CONSIDERATIONS. FOR, (NASA) Unclas 64 p HC $6.25 CSCL 18G G3/05 38494 STUDY OF EXTRATERRESTRIAL DISPOSAL OF RADIOACTIVE WASTES by R. L. Thompson, J. R. Ramler, and S. M. Stevenson Lewis Research Center Cleveland, Ohio 44135 May 1974 This information is being published in prelimi- nary form in order to expedite its early release. I STUDY OF EXTRATERRESTRIAL DISPOSAL OF RADIOACTIVE WASTES Part I Space Transportation and Destination Considerations for Extraterrestrial Disposal of Radioactive W4astes by R. L. Thompson, J. R. Ramler, and S. M. Stevenson Lewis Research Center CT% CSUMMARY I NASA has been requested by the AEC to conduct a feasibility study of extraterrestrial (space) disposal of radioactive waste. This report covers the initial work done on only one part of the NASA study, the evaluation and comparison of possible space destinations and space transportation systems. Only current or planned space transportation systems have been considered thus far. The currently planned Space Shuttle was found to be more cost-effective than current expendable launch vehicles by about a factor of 2. The Space Shuttle requires a third stage to perform the waste disposal missions. Depending on the particular mission, this third stage could be either a reusable space tug or an expendable stage such as a Centaur. Of the destinations considered, high Earth orbits (between geostationary and lunar orbit altitudes), solar orbits (such as a 0.90 AU circular solar orbit) or a direct injection to solar system escape appear to be the best candidates.
    [Show full text]
  • Analysis of Orbital Decay Time for the Classical Hydrogen Atom Interacting with Circularly Polarized Electromagnetic Radiation
    Analysis of Orbital Decay Time for the Classical Hydrogen Atom Interacting with Circularly Polarized Electromagnetic Radiation Daniel C. Cole and Yi Zou Dept. of Manufacturing Engineering, 15 St. Mary’s Street, Boston University, Brookline, MA 02446 Abstract Here we show that a wide range of states of phases and amplitudes exist for a circularly polarized (CP) plane wave to act on a classical hydrogen model to achieve infinite times of stability (i.e., no orbital decay due to radiation reaction effects). An analytic solution is first deduced to show this effect for circular orbits in the nonrelativistic approximation. We then use this analytic result to help provide insight into detailed simulation investigations of deviations from these idealistic conditions. By changing the phase of the CP wave, the time td when orbital decay sets in can be made to vary enormously. The patterns of this behavior are examined here and analyzed in physical terms for the underlying but rather unintuitive reasons for these nonlinear effects. We speculate that most of these effects can be generalized to analogous elliptical orbital conditions with a specific infinite set of CP waves present. The article ends by briefly considering multiple CP plane waves acting on the classical hydrogen atom in an initial circular orbital state, resulting in “jump- and diffusion-like” orbital motions for this highly nonlinear system. These simple examples reveal the possibility of very rich and complex patterns that occur when a wide spectrum of radiation acts on this classical hydrogen system. 1 I. INTRODUCTION The hydrogen atom has received renewed attention in the past decade or so, due to studies involved with Rydberg analysis, chaos, and scarring [1], [2], [3], [4].
    [Show full text]
  • Collisions & Encounters I
    Collisions & Encounters I A rAS ¡ r r AB b 0 rBS B Let A encounter B with an initial velocity v and an impact parameter b. 1 A star S (red dot) in A gains energy wrt the center of A due to the fact that the center of A and S feel a different gravitational force due to B. Let ~v be the velocity of S wrt A then dES = ~v ~g[~r (t)] ~v ~ Φ [~r (t)] ~ Φ [~r (t)] dt · BS ≡ · −r B AB − r B BS We define ~r0 as the position vector ~rAB of closest approach, which occurs at time t0. Collisions & Encounters II If we increase v then ~r0 b and the energy increase 1 j j ! t0 ∆ES(t0) ~v ~g[~rBS(t)] dt ≡ 0 · R dimishes, simply because t0 becomes smaller. Thus, for a larger impact velocity v the star S withdraws less energy from the relative orbit between 1 A and B. This implies that we can define a critical velocity vcrit, such that for v > vcrit galaxy A reaches ~r0 with sufficient energy to escape to infinity. 1 If, on the other hand, v < vcrit then systems A and B will merge. 1 If v vcrit then we can use the impulse approximation to analytically 1 calculate the effect of the encounter. < In most cases of astrophysical interest, however, v vcrit and we have 1 to resort to numerical simulations to compute the outcome∼ of the encounter. However, in the special case where MA MB or MA MB we can describe the evolution with dynamical friction , for which analytical estimates are available.
    [Show full text]
  • Orbital Mechanics Joe Spier, K6WAO – AMSAT Director for Education ARRL 100Th Centennial Educational Forum 1 History
    Orbital Mechanics Joe Spier, K6WAO – AMSAT Director for Education ARRL 100th Centennial Educational Forum 1 History Astrology » Pseudoscience based on several systems of divination based on the premise that there is a relationship between astronomical phenomena and events in the human world. » Many cultures have attached importance to astronomical events, and the Indians, Chinese, and Mayans developed elaborate systems for predicting terrestrial events from celestial observations. » In the West, astrology most often consists of a system of horoscopes purporting to explain aspects of a person's personality and predict future events in their life based on the positions of the sun, moon, and other celestial objects at the time of their birth. » The majority of professional astrologers rely on such systems. 2 History Astronomy » Astronomy is a natural science which is the study of celestial objects (such as stars, galaxies, planets, moons, and nebulae), the physics, chemistry, and evolution of such objects, and phenomena that originate outside the atmosphere of Earth, including supernovae explosions, gamma ray bursts, and cosmic microwave background radiation. » Astronomy is one of the oldest sciences. » Prehistoric cultures have left astronomical artifacts such as the Egyptian monuments and Nubian monuments, and early civilizations such as the Babylonians, Greeks, Chinese, Indians, Iranians and Maya performed methodical observations of the night sky. » The invention of the telescope was required before astronomy was able to develop into a modern science. » Historically, astronomy has included disciplines as diverse as astrometry, celestial navigation, observational astronomy and the making of calendars, but professional astronomy is nowadays often considered to be synonymous with astrophysics.
    [Show full text]
  • Orbital Decay Analysis for Debris Deorbiting Cubesats in Leo: a Case Study for the Velox-Ii Deorbit Mission
    ORBITAL DECAY ANALYSIS FOR DEBRIS DEORBITING CUBESATS IN LEO: A CASE STUDY FOR THE VELOX-II DEORBIT MISSION Sarat Chandra Nagavarapu(1), Amal Chandran(1), and Daniel E. Hastings(2) (1)Satellite Research Centre (SaRC), Nanyang Technological University, Singapore, 639798, Email: [email protected], [email protected] (2)Aeronautics & Astronautics Department, MIT, Cambridge, MA, USA, 02139, Email: [email protected] ABSTRACT Keywords: Space debris; LEO; orbital decay; debris re- moval; VELOX-II; CubeSat. Since the launch of Sputnik 1 into space in 1957, hu- mans have sent thousands of rockets and more than ten thousand satellites into Earth’s orbit. With hundreds of 1. INTRODUCTION more launched every year, in-space collisions and the creation of high-speed debris fragments are becoming in- creasingly likely, threatening future space missions. The Space debris is any man-made object in orbit about the term orbital debris comprises defunct satellites, spent up- Earth that no longer serves a useful function. The Earth per stages, fragments resulted from collisions, etc., which orbit is in serious danger caused by millions of these pose a significant threat to many other operational satel- space debris. Since the beginning of the space age, more lites in their vicinity. With the advent of satellite mega- than 10,680 satellites have been placed into orbit [1]. Out constellations in low Earth orbit (LEO), the LEO envi- of which 6,250 are still in the orbit and only 3,700 are op- ronment is becoming extremely crowded. In addition to erational, leaving an astonishing 2550 dead satellites in the government space organizations, several private space the orbit.
    [Show full text]
  • Atmospheric Drag Effects on Modelled LEO Satellites During the July 2000 Bastille Day Event in Contrast to an Interval of Geomagnetically Quiet Conditions Victor U
    https://doi.org/10.5194/angeo-2020-33 Preprint. Discussion started: 29 June 2020 c Author(s) 2020. CC BY 4.0 License. Atmospheric drag effects on modelled LEO satellites during the July 2000 Bastille Day event in contrast to an interval of geomagnetically quiet conditions Victor U. J. Nwankwo1, William Denig2, Sandip K. Chakrabarti3, Muyiwa P. Ajakaiye1, Johnson Fatokun1, Adeniyi W. Akanni1, Jean-Pierre Raulin4, Emilia Correia4, and John E. Enoh5 1Anchor University, Lagos 100278, Nigeria 2St. Joseph College of Maine, Standish, ME 04084, U.S.A 3Indian Centre for Space Physics, Kolkata 700084, India 4Centro de Rádio Astronomia e Astrofísica Mackenzie, Universidade Presbiteriana Mackenzie, São Paulo, Brazil 5Interorbital systems, Mojave, CA 93502-0662, U.S.A. Correspondence: Victor U. J. Nwankwo ([email protected]) Abstract. In this work we simulated the effects of atmospheric drag on two model SmallSats in Low Earth Orbit (LEO) with different ballistic coefficients during 1-month intervals of solar-geomagnetic quiet and perturbed conditions. The goal of this effort was to quantify how solar-geomagnetic activity influences atmospheric drag and perturbs satellite orbits. Atmospheric drag compromises satellite operations due to increased ephemeris errors, attitude positional uncertainties and premature satel- 5 lite re-entry. During a 1-month interval of generally quiescent solar-geomagnetic activity (July 2006) the decay in altitude (h) 3 2 3 was a modest 0.53 km (0.66 km) for the satellite with the smaller (larger) ballistic coefficient of 2.2 10− m /kg (3.03 10− × × m2/kg). The associated Orbital Decay Rates (ODRs) during this quiet interval ranged from 13 m/day to 23 m/day (from 16 m/day to 29 m/day).
    [Show full text]
  • Kickstart Physics Space
    KICKSTART PHYSICS SPACE 1. SPEED AND ESCAPE VELOCITY 2. PROJECTILE MOTION 3. ACCELERATION AND G-FORCES 4. ‘C’ AND RELATIVITY 5. EINSTEIN Kickstart would like to acknowledge and pay respect to the traditional owners of the land – the Gadigal people of the Eora Nation. It is upon their ancestral lands that the University of Sydney is built. As we share our knowledge, teaching, learning, and research practices within this University may we also pay respect to the knowledge embedded forever within the Aboriginal Custodianship of Country. For more information head to http://sydney.edu.au/science/outreach/high-school/kickstart/index.shtml The University of Sydney School of Physics Space Get to Know the Inner Space Risk Analysis Conduct a risk analysis by filling out the following table. List 3 risks to do with this investigation in the 2nd year lab. Also, list 3 risks in an industry where physics is used. − What are the consequences of those risks? Use the risk matrix below to make your judgment. − What precautions would you take to stop those risks from coming about? − What steps would you take to mitigate those risks? Risk Consequence Precaution/Mitigation Electricity Electrocution (High) Safety Switch Radiation Radiation Sickness (High) Low Dose/Shielding Kickstart Equipment Tripping and Injury (Medium) Don’t Touch Water Slipping and Injury (Medium) No Food/Drink in Lab LN2 Frostbite (High) Personal Protective Equipment Industry – Finance/Tech, Games, Solar Power, Medical Physics, Radiology Physics Research – University, See above, jobs might require See above, jobs will require Industry Academia more specific considerations more specific considerations Public Service – CSIRO, DSTO, NMI, Questacon, Government, Teachers What sort of careers do you think you could get if you studied this topic at the University of Sydney? 1 The University of Sydney School of Physics Space Speed & Escape Velocity There are a number of different types of orbits in orbital systems.
    [Show full text]
  • Samuel YW Low (Samuel [email protected])
    Assessment of Orbit Maintenance Strategies for Small Satellites Samuel Y. W. Low ([email protected]), Singapore University of Technology and Design | Y. X. Chia ([email protected]), DSO National Laboratories Objectives: To explore and validate how various orbit maintenance strategies for small satellites can be optimized for V usage, by varying the tolerance bands as the independent variable. Δ Needs: A computationally efficient and accurate model for the rate of orbit decay, and a model small satellite in LEO as test parameters. Summary: Three methods were proposed: one of forced-Keplerian orbits using constant thrust, one via Hohmann transfers, and one via cyclical direct burns. Results show potential for minimizing V usage, and also reveal and interesting connection as tolerance bands tend to infinitesimally small values. 0. THE ORBITAL DECAY MODEL Δ II. METHOD OF HOHMANN TRANSFERS III. METHOD OF CYCLICAL DIRECT BURNS VELOX-CI Satellite Information Decay Data 2.444 Payload: Radio Occultation Unit from Model km 615 X 608 X 848 mm3 Decay Data 2.566 L (km) Mass = 123 kg from GPS km R (km) Semi major axis = 6928.14 km Fig 2: Decay data comparisons Altitude = 545 km with proposed model Inclination = 15 degrees Fig 1: Model satellite and orbit parameters used in the simulation RAAN, ARP, True Anomaly = 0 Model Satellite & Parameters: VELOX-CI Fig 4: Hohmann transfer, Orbit Characteristics: Circular orbit, near-equatorial against radial axis tolerance band (km) Fig 6: Single ∆V of single the in-track axis tolerance band (km) Orbit Decay Validation Period: 16/12/2015 to 11/11/2016 .
    [Show full text]