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Apophis Mitigation Technology Flight Test: Design Review and Mission Overview

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luisa
0 diapositives · 125 Views · Jul 03, 2014

The Apophis Mitigation Technology Flight Test, detailing the design review and mission overview. It encompasses the background, mission statement, and specific requirements for the test. Key areas of focus include the albedo change demonstration, telecommunications, instrumentation, propulsion, attitude control, structures, thermal management, and budget scheduling. The test aims to assess innovative technologies for asteroid deflection to safeguard Earth from potential impacts.

#apophis mitigation #technology flight test #asteroid deflection #mission overview
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Apophis Mitigation Technology Flight Test: Design Review and Mission Overview
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  1. Apophis Mitigation Technology Flight Test

    Design Review
  2. Overview Introduction Background Mission Statement Requirements Mission Overview Albedo Change Demo Telecommunications and Instrumentation Propulsion and Attitude Control Structures, Thermal and Power Budget and Scheduling Conclusions
  3. Introduction Students were challenged to develop an actual proposal for a large scale (~ $30M) space flight experiment and submit it to NASA Headquarters and the International Institute for Lunar and NEO Research. This is a collaboration among Texas A&M, NASA Ames Research Center and King Abdul Aziz City for Science and Technology, Saudi Arabia. Objective: To design a low Earth orbit demonstration of a technology capable of deflecting the dangerous Near-Earth asteroid Apophis from possible Earth impact.
  4. Background Discovered on June 19th, 2004 by R. A. Tucker, D. J. Tholen and F. Bernardi at Kitt Peak Orbital models have identified several close Earth approaches with the 2029 approach potentially passing through a gravitational keyhole which could swing Apophis into a collision path in 2036 As an Aten class NEA, Apophis rated as high as 4 on the Torino scale but has been downgraded to a 0 as its probability of impact has been commonly accepted at 1/45,000
  5. Background Cont. Learning Through Research (LTR) educational program at Texas A&M gave rise to the project Fall ’07 inaugurated a sequence of three undergraduate classes addressing the design of the Apophis Preliminary Exploration Platform (APEP) mission Fall ’08 design class was challenged to develop a system that can significantly modify the orbit of Apophis by its close approach in ’36
  6. Mission Statement The Apophis Mitigation Technology Flight Test will demonstrate the feasibility of an albedo change prototype on a target surface in Low Earth Orbit to practice mitigation of dangerous Near Earth Objects in a controlled environment.
  7. Requirements Mission Purpose and Conditions The purpose of the AMT Flight Test is to demonstrate successful operation of the Surface Albedo Treatment Subsystem (SATS) in the Low Earth Orbit (LEO) environment, subject to the following conditions: The test article shall be a scaled version of the SATS design intended to fly on the 2022 Apophis Exploration Mission (AEMP), as described in the NRC proposal. Operational characteristics to be verified are restricted to the successful deposition of albedo change material onto a test object together with the intended albedo change in the test object. The mission shall furnish instrumentation to verify (1) dispensing cone angle, (2) flow and deposition rates, (3) coverage efficiency, (4) albedo change on the test object.
  8. Requirements Cont. Mission Parameters Total mass of the AMT spacecraft(s) shall not exceed 50 kg Launch shall be no later than 2011 Margins on mass, power and cost shall be 30% Mission cost, excluding launch and operations, shall not exceed $30 million (2009)
  9. Mission Overview Apophis Mitigation Technology Flight Test Build satellite mock-up Prepare and execute a series of ground tests Build working satellite Launch into LEO to test SATS technology
  10. Frans Ebersohn – Group Leader Austin Bond Brannen Clark Joshua Hempel Julianne Larson Agustin Maqui Andrew Schaeperkoetter Albedo Change Demo
  11. SPADE – Static Preliminary Albedo Demonstration Experiment Pressurant Gas Canister Powder Canister Torque Rod (3) Sun Sensor (4) Tribogun Tube Electronics Bay Batteries Antenna Camera Test Surface
  12. Experiment Mission Profile V∞ 1. Orient spacecraft 2. Charge test surface then remove power supply 3. Initiate tribogun and spray test surface
  13. Experiment Mission Profile (cont) 4. Allow particles time to cure 5. Observe, record, and transmit data Data
  14. Attitude Requirements Orient so that panel is facing sun when on sun side of earth to allow particles to cure in sunlight Orient so that main body of the craft shields the panel while moving through LEO atmosphere Main body in ‘ram’ direction, panel in ‘plasma wake’ Drag on particles should not be an issue due to shielding and particle size V∞ All information above from Reference 16
  15. Test Surface Charging Parallel plate capacitor Voltmeter attached to measure voltage and thus charge Top surface is positively charged and particles negatively charged E
  16. Capacitor Design Parallel plate capacitor Conducting Plates made of Aluminum Glass Dielectric 8,9 Required Electric Field based on expected Apophis charge density 15 Ereq = 10 – 20 Volts/Meter Three strips of different roughness from grinding Extending from main body
  17. Tribogun Mass Budget
  18. Liquefaction flow Albedo Change Particles ACP Chamber Pressurized Inert Gas Tribo ionization tube Average radius = RT Length = LT Mixing Chamber System Sizing
  19. Tribogun System Power Budget Charge capacitor with a 12 Volt potential for ~1 second Power Required to charge = 14.3 Watts
  20. Albedo Change Particle Dispensing Characteristics
  21. Minimum Visual Camera Imaging Requirements
  22. Travis Jacobs – Group Leader Darkhan Alimzhanov Jonathan Ellens Patrick Harrington Cathy Spohnheimer Barrett Wight Telecommunications and Instrumentation
  23. Communication Launch site: BaikonurCosmodrome at 45.9° north latitude into 350 km circular orbit Ground station: College Station, TX (30.6° N, 96.3° W) Longitude of ascending node = -138° εmin = 5°
  24. Link Budget P = -191.3 + 8.69ln(s) + 4.34ln(R) P(1657 km, 56kbps) = 11.3mW Formulas taken from Reference 5
  25. Communication Sizing Data rate sized to transmit 300 images of 460 KB each to earth 56 kbps Transmission times 7 minute ground pass 2 minute command/acquisition time 3 minute communication time 2 minute buffer in case of trouble
  26. Communication UHF beacon from reference 7 Antenna and Transceiver from Reference 6
  27. Camera Camera is mission critical component, need something that has been flight tested Used on Cosmos-1 Camera from Reference 11
  28. Computer GUMSTIX Not space qualified, needs testing on operating temperature and radiation Runs on Linux Overo Earth 256 MB RAM 256 MB Flash Micro SD card slot Power ~ 3W Mass <100 g Computer from Reference 10
  29. GPS Antenna and Receiver GPS from Reference 6
  30. Total Requirements for Telecommunication and Instrumentation 1does not include computer and camera
  31. Nathan Jones – Group Leader Kenneth Barnes Christina Daughtrey Brandt King Brian Kuehner Timothy Lowery Jared Wissel Propulsion and Attitude Control
  32. Attitude Sensing Sun Sensor Magnetometer Data and photos obtained from Reference 12
  33. Attitude Control Torque Rods MTR-5 Magnetorquer Mass = 0.5 kg Power = 1 W Total of 3 Total Mass = 1.5 kg Total Power = 3 W Use torque rods only, no need for course corrections - only need to control the orientation of the satellite to keep capacitor plate in the sunlight. Data and photo obtained from Reference 12
  34. Control Torque Control Torque Disturbing Torques Gravity gradient Solar pressure Magnetic field variations Atmospheric drag Time to Rotate 90°
  35. ESPA Adapter Ring Sizing is based on the Minotaur IV Standard Fairing SL-ESPA 24 Satellite Total Weight < 97 kg Dimensions 24in x 20in x 28in Lightband Separation System Adds 2.5 kg to total weight on ESPA ring Data and photos obtained from Reference 13
  36. Lightband Separation System Two primary functions Rigidly holds two adjoining vehicles together for shipment and launch Affect separation of those vehicles upon command from an adjoining vehicle Advantages Simplifies payload integration from days to minutes Eliminates hazards associated with pyrotechnics and fracture System indicates to the vehicles the state of separation (separated or joined) Average velocity of separation 0.25 m/s Lowers cost Saves weight Reduces shock Data and photos obtained from Reference 14
  37. Orbital Requirements Reference 5 Inclination comes from assuming we launch due east from Baikonur, which has a latitude of 45.9o. Ballistic Coefficient is calculated using a max and min area that the atmosphere interacts with during the orbit and using Cd = 2. The Orbit Lifetime is found from the ballistic coefficient and the altitude.
  38. Summary
  39. Erica Furnia – Group Leader Brian Atteberry Wesley Fite Jason York Stephen Oehler Jennifer Wells – Assistant Program Manager Altay Zhakatayev Structures, Thermal and Power
  40. Top View Left View Front View
  41. Universal Assumptions: 30 LOS passes total, 7 mins/pass Gumstixcomputers used (~2-6 Watts each) Battery Scenarios
  42. ESPA Contraints EELV Secondary Payload Adapter (ESPA) Small Launch scaled version 38.8” primary interface diameter Sized for: Minotaur IV Falcon 1e Taurus Delta II Fits CubeSats up to 180 kg Flight validation costs are low Use existing test facilities Data and photos obtained from Reference 13
  43. ESPA Contraints Minotaur Standard Fairing 8” diameter secondary payload (SPL) interface Can host 6 SPLs measuring 24x20x18.8” and weighing 100 kg Falcon 1e Fairing Payload volume of approx. 2,785 in.3 Full load weight is about 151.95 kg Data and photos obtained from Reference 13
  44. Thermal Analysis Assumptions: Satellite Dimensions: a=20 in, b=24 in, c=28 in Area exposed to Sun rays: A0=a*b+b*c+a*c. ATotal =2*A0 Flat Plate Capacitor Total area of capacitor is exposed to perpendicular sunrays Goals: Tmin~-30 °C, Tmax~70 °C Functions of absorption and emission coefficients of satellite Minimize absorption and emission ↓α = ↓Tmax, ↓ε = ↑Tmax & ↑Tmin Find cheap and reliable way to meet those criteria Find ε and α that provide those temperature limits
  45. Optimize Tmax and Tmin Option 1 ε affects only Tmin, while α affects both temperatures Backward calculations: εd=0.32, αd=0.11 From SMAD5: ¼ mil Aluminized Mylar (degrades in sunlight) ε=0.34 anodized aluminum ε=0.04..0.88 2-5 mil Silvered or Aluminized Teflon α=0.05 and 0.1 vapor deposited aluminum α=0.08..0.17 bare aluminum α=0.09..0.17 Option 2 Move to a colder Temperature Choose low absorptivity and high emissivity material Examples Z93 white paint with α=0.17..0.2 ε=0.92 ZOT paint α=0.18..0.2, ε=0.91 Choose Z93 white paint with α=0.3 and ε=0.92 Tmax=298.7 K, Tmin=186 K Tmax is good, but Tmin is too low Use Patch Heaters
  46. Option 3 Move to hot part of T Choose material with low absorptivity and low emissivity Example: bare aluminum with α=0.17 and ε=0.1 Choose α=0.27 and ε=0.1 Tmax=512 K, Tmin=324 K Tmin is good, but Tmax needs to be fixed In general, it is desirable to be in colder temperature region Need for radiator Hard to satisfy thermal requirements αx=0.3 in Figure: Results Choose either Option 1 or 2 for Satellite Flat Plate Capacitor Case with chosen material from option1: Tmax=305 K, Tmin=232.9 K Case with Z93 white paint: Tmax=280.9 K, Tmin=178.9 K Case with bare aluminum: Tmax=479 K, Tmin=311.5 K Optimize Tmaxand TminCont.
  47. Structure, Thermal, and Power Mass Budget
  48. Total Mass Budget
  49. Andrei Kolomenski – Group Leader Danielle Fitch – Program Manager Cory Phillips Kris Keiser Scott Southwell Budget and Scheduling
  50. Mission Schedule 2010 January – Begin construction of “Mock up” setup April – “Mock up” setup constructed May– Customize setup for vacuum chamber test June– Begin vacuum chamber test July– Customize setup for plasma environment test August/September– Begin plasma environment test October/November– Construct chamber for 0-g test December– Customize setup for 0-g test 2011 January-March – Window for performing 0-g test March-June– Analyze data and optimize design June– Finalize actual design June-November– Construct actual satellite November/December– Launch
  51. Test Setups & Variables We must isolate the effect of the environment on the spray pattern of the Albedo Change Particles (ACPs). Zero-g aircraft test Plasma environment test Ground vacuum chamber test Variables that will vary among the experiments: Material composition, roughness and electric charge of painting surface Variables that remain constant among experiments: Distance between Tribogun and surface Velocity of ACP discharge Mass of powder ejected -All setups will be identical, aside from the environment they are in.
  52. Problems with Testing Vacuum test chamber: Tribogun pressurization Zero-g aircraft test Limited testing time (30-40 sec. intervals of Zero-g conditions) Securing / Suspending the satellite elements during flight Cost and availability of aircraft Plasma environment test Electrical interference with satellite circuitry may cause erroneous measurements Difficulties in data acquisition Complexity and cost of creating a sustainable plasma environment for testing. - All tests require remote actuation of testing technology
  53. Parametric Cost Budget All Cost is expressed in Millions of Dollars
  54. Conclusions Very tight schedule may be difficult to follow, but this experiment will prove the feasibility of distributing the albedo change particles in an actual space environment This validates a unique technology that acts permanently to alter the trajectory of a hazardous NEO Total Mass – 54 kg Total Power – 180 W Total Cost - $ 34 M
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