1. 1 Design Review

2. 2 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. 3 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. 4 Discovered on June 19 th, 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. 5 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. 6 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. 7 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 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. 8 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. 9 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. 10 Frans Ebersohn – Group Leader Austin Bond Brannen Clark Joshua Hempel Julianne Larson Agustin Maqui Andrew Schaeperkoetter

11. 11 SPADE – Static Preliminary Albedo Demonstration Experiment

12. 12 1. Orient spacecraft V 2. Charge test surface then remove power supply 3. Initiate tribogun and spray test surface

13. 13 4. Allow particles time to cure 5. Observe, record, and transmit data Data

14. 14 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

15. 15 Parallel plate capacitor Voltmeter attached to measure voltage and thus charge Top surface is positively charged and particles negatively charged Glass Dielectric Multiple surface roughness to test

16. 16 Plate Area.217 m 2 Assumed Resistivity2.82 × 10 -8 Ω·m Dielectric Constant9.15 Distance between Plates.01 m Potential Difference Between Plates.1 V Capacitance1.76 x 10 -9 F Charge on top plate1.76 x 10 -10 C Conducting Plates made of Aluminum Required Electric Field = 10 – 20 Volts/Meter Three strips of different roughness from grinding

17. 17 Mass of powder (M ACP ).694 kg Powder canister.313 kg Pressurant gas.139 kg Pressurant tank 0.3 kg Tribogun0.6 kg Control electronics0.4 kg Tubing and valves0.5 kg Total2.946 kg

18. 18 Pressurized Inert Gas Albedo Change Particles Liquefaction flow ACP Chamber Mixing Chamber Tribo ionization tube Average radius = R T Length = L T Fraction of Mixture Volume Occupied by ACPs Pressurant Gas Tank Radius Powder Canister Radius 0.57.46 cm4.99 cm Tribo Ionization Tube Average Radius (R T ) Tribo Ionization Tube Length ( L T ) 0.1 cm33.2 cm

19. 19 Tribogun PowerCapacitor PowerTotal 80 Watts14.3 Watts94.3 Watts Charge capacitor with a 12 Volt potential for ~1 second Power Required to charge = 14.3 Watts

20. 20 Average Particle Velocity1 m/s Mass Flow Rate0.0025 kg/s Dispensing Time276 s Force on Craft0.0025 N Max Travel Time of Particles0.691 s

21. 21 Image Rate (fps) Total Frames During Spray1276 During Curing (30 mins) 0.016730 Mission Total306

22. 22 Travis Jacobs – Group Leader Darkhan Alimzhanov Jonathan Ellens Patrick Harrington Cathy Spohnheimer Barrett Wright

23. 23 Launch site: Baikonur Cosmodrome at 45.9° north latitude into 350 km circular orbit Ground station: College Station, TX (30.6° N, 96.3° W) L node = -138° ε min = 5°

24. 24 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

25. 25 Cubesat Transceiver (CTR-450) Power20 mW Mass200 g Operating Temp-20C to 60C Operating Frequencies370 – 470 MHz Data Rates1 – 154 kbps UHF Monopole Antenna (ANT-100) Mass100 g Operating Temp-40C to 80C UHF Beacon Mass85 g Power1.9 W (during transmission) Operating Temp-20C to 60C Cost$11,250 (7500 Euro)

26. 26 Camera is mission critical component, need something that has been flight tested Used on Cosmos-1 TPS Camera Mass120 g (including radiation shielding) Power0.5 W Resolution640 x 480 pixels (0.3 Megapixels)

27. 27 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

28. 28 GPS Reciever (GPS-12-V1) Position Accuracy10 m Velocity Accuracy0.03 m/s Mass200 g Power1 W Operating Temp-40C to 85 C GPS Antenna (ANT-GPS) Mass82 g Operating Temp-55C to 85C

29. 29 Mass887 g Power6.4 W Operating Temp-20C to 60C 1 1 does not include computer and camera

30. 30 Nathan Jones – Group Leader Kenneth Barnes Christina Daughtrey Brandt King Brian Kuehner Timothy Lowery Jared Wissel

31. 31 SensorParameters 4 Sun SensorsMass = 1.2 kg Power < 1.2 W MagnetometerMass = 0.14 kg Power < 300 mW Sun Sensor Magnetometer

32. 32 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 with respect to the sun.

33. 33 Disturbing Torques Gravity gradient Solar pressure Magnetic field variations Atmospheric drag Control Torque Time to Rotate 90°

34. 34 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

35. 35 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

36. 36 ParameterValue Altitude350 km Inclination30 o Ballistic Coefficient58 – 82 kg/m 2 Orbit Lifetime30 – 200 days Inclination comes from assuming we launch due east from Cape Canaveral, which has a latitude of 28.5 o. Ballistic Coefficient is calculated using a max and min area that the atmosphere interacts with during the orbit and using C d = 2. The Orbit Lifetime is found from the ballistic coefficient and the altitude.

37. 37 ParameterRequiredWith 30% Contingency Power4.2 W5.5 W Mass2.8 kg3.7 kg Settling Time5.24 min ESPA RequirementsMax Value Mass< 97 kg Size24in x 20in x 28in ParameterValue Altitude350 km Inclination30 o Ballistic Coefficient58 – 82 kg/m 2 Orbit Lifetime30 – 200 days

38. 38 Erica Furnia – Group Leader Brian Atteberry Wesley Fite Jason York Stephen Oehler Jennifer Wells Altay Zhakatayev

39. 39 Spacecraft Structure

40. 40 Universal Assumptions: 30 LOS passes total, 7 mins/pass Gumstick computers used (~2-6 Watts each)

41. 41  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

42. 42  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

43. 43 Assumptions: Satellite  Dimensions: a=20 in, b=24 in, c=28 in  Area exposed to Sun rays: A 0 =a*b+b*c+a*c. A Total =2*A 0 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 Gs (W/m^2)1420 Earth albedo0.36 H (km)350 emissivity of Earth0.8 Earth IR (W/m^2)244 (Max Hot), 218 (Max Cold) Q int (W)200 (Max Hot), 20 (Max Cold)

44. 44 Option 1 ε affects only T min, while α affects both temperatures Backward calculations: ε d =0.32, α d =0.11 From SMAD:  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 colder T Choose low absorbtivity 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

45. 45 Option 3 Move to hot part of T Choose material with low absorbtivity 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 α x =0.3 in Figure: Results 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 Choose either Option 1 or 2 Optimize T max and T min Cont.

46. 46 Total Mass Budget Mass [kg] Structure, Thermal, and Power Mass Budget Mass [kg] Albedo Change Demo5.9Housing (Satellite Shell)19.1 Telecommunications and Instrumentations1.0Mounting Hardware0.8 Propulsion and Attitude Control2.8Thermal Insulation1.0 Structure, Thermal, and Power35.6Batteries14.8 Total45.3Total35.6 Total (with 30% safety factor)58.9Total (with 30% safety factor)46.3

47. 47 Andrei Kolomenski – Group Leader Danielle Fitch Cory Phillips Kris Keiser Scott Southwell

48. 48 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

49. 49 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.

50. 50  Vacuum test chamber:  Tribo gun pressurization  Zero-g aircraft test  Limited testing time (30-40 sec. intervals of 0-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 Problems with Testing:

51. 51 MTR-5 Torque Rod Three needed for 3-axis control Have been successfully used on 24 LEO Launches Most expensive part of ACS Reliable SSTL 2-axis Sun Sensor Multiple Sun Sensors needed Very low cost and power requirements Extremely reliable and well tested Magnetometer Reliable attitude sensing Over 25 successful missions with 50 magnetometers have been flown Very low mass and power requirements Cost: 1.5 million to 2.5 million Possible margin of error of up to 20% Risk: All components are extremely reliable. Only risk is possible Electromagnetic Interference.

52. 52 Specs:  Field of view: 100°  Embedded lossless image compression  Auto-exposure capability  Multiple image storage capacity  10 Mbits/s data rate  Minimal resolution of 1024 x 1024 pixels  Power consumption: 1.8 W  Camera mass with bracket: 221 g  Approximately camera bracket is 3” x 5” Cost:  $5,000-10,000 (depending on customization) Risk:  Radiation damage  Component aging (probably not an issue)  Temperature drift Example Camera: X-CAM Micro-Camera

53. 53 Specs:  Common antenna  A single metal patch over a ground plate  Most cost effective model and best scale for our requirements Cost:  Approximately $500 Risk:  Negligible risk S-Band Patch Antenna

54. 54 RDT&E Cost, 1 st Unit Cost and SubTotal Cost values are expressed in thousands of dollars

55. 55 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

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