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 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 TestBuild satellite mock-up Prepare and execute a series of ground
testsBuild working satelliteLaunch into LEO to test SATS technology
10
Frans Ebersohn – Group LeaderAustin BondBrannen ClarkJoshua HempelJulianne LarsonAgustin MaquiAndrew Schaeperkoetter
Albedo Change Demo
11
SPADE – Static Preliminary Albedo Demonstration Experiment
12
Experiment Mission Profile
1. Orient spacecraftV
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 sizeV
15
Surface Charging
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
Capacitor Design
Plate Area .217 m2
Assumed Resistivity 2.82 × 10-8 Ω·m
Dielectric Constant 9.15
Distance between Plates .01 m
Potential Difference Between Plates
.1 V
Capacitance 1.76 x 10-9 F
Charge on top plate 1.76 x 10-10 C
• Conducting Plates made of Aluminum• Required Electric Field = 10 – 20 Volts/Meter• Three strips of different roughness from grinding
17
Mass BudgetMass of powder (MACP) .694 kg
Powder canister .313 kg
Pressurant gas .139 kg
Pressurant tank 0.3 kg
Tribogun 0.6 kg
Control electronics 0.4 kg
Tubing and valves 0.5 kg
Total 2.946 kg
18
System Sizing
Pressurized Inert Gas
Albedo Change Particles
Liquefaction flow
ACP Chamber
Mixing Chambe
r
Tribo ionization tubeAverage radius = RT
Length = LT
Fraction of Mixture Volume Occupied by ACPs
Pressurant Gas Tank Radius
Powder Canister Radius
0.5 7.46 cm 4.99 cm
Tribo Ionization Tube Average Radius (R T )
Tribo Ionization Tube Length ( LT )
0.1 cm 33.2 cm
19
Power Budget
Tribogun Power Capacitor Power Total
80 Watts 14.3 Watts 94.3 Watts
Charge capacitor with a 12 Volt potential for ~1 second
Power Required to charge = 14.3 Watts
20
Spray Profile
Average Particle Velocity
1 m/s
Mass Flow Rate 0.0025 kg/s
Dispensing Time 276 s
Force on Craft 0.0025 N
Max Travel Time of Particles
0.691 s
21
Minimum Imaging Requirements
Image Rate (fps)
Total Frames
During Spray
1 276
During Curing (30 mins)
0.0167 30
Mission Total 306
22
Travis Jacobs – Group LeaderDarkhan AlimzhanovJonathan EllensPatrick HarringtonCathy SpohnheimerBarrett Wright
Telecommunications and Instrumentation
23
Communication
a. Earth angular radius, sin 71.4E ER R H
Parameter Formula Value
3 24
max max min max
deg
b. Period, 1.658669 10 6378.14 91.5min
c. max nadir angle, sin sin cos 70.8d
P P H P
0max max min max max
max max max max max
eg
d. max Earth cen. angle, 90 14.2deg
e. max distance, sin sin 1656.8ED D R D k
1min min minf. min Earth cen. angle, sin (sin sin 3.2deg
cos cos cos )
g. min nadir
pole gs
pole gs
m
lat lat
lat lat long
1min min min min
min
m
angle, tan ( sin sin 44.9deg
1 sin cos )
h. max elevation angle,
0ax max min min max
min min min min min
ma
90 41.9deg
i. min distance, sin sin 508.4
j. max angular rate,
ED D R D km
x max min max
min max
2 52.1deg min
k. azimuth range, cos 2 tan tan 154.3deg
l. time in view,
ER H PD
T
0 1max min 180 cos cos cos 7.04min T P T
Launch site: Baikonur Cosmodrome at 45.9° north latitude into 350 km circular orbitGround station: College Station, TX (30.6° N, 96.3° W)Lnode = -138° εmin = 5°
24
Communication Sizing
Data rate sized to transmit 300 images of 460 KB each to earth56 kbps
Transmission times7 minute ground pass
2 minute command/acquisition time 3 minute communication time 2 minute buffer in case of trouble
25
Communication
Cubesat Transceiver (CTR-450)
Power 20 mW
Mass 200 g
Operating Temp -20C to 60C
Operating Frequencies 370 – 470 MHz
Data Rates 1 – 154 kbps
UHF Monopole Antenna (ANT-100)
Mass 100 g
Operating Temp -40C to 80C
UHF Beacon
Mass 85 g
Power 1.9 W (during transmission)
Operating Temp -20C to 60C
Cost $11,250 (7500 Euro)
26
Camera
Camera is mission critical component, need something that has been flight tested
Used on Cosmos-1TPS Camera
Mass 120 g (including radiation shielding)
Power 0.5 W
Resolution 640 x 480 pixels (0.3 Megapixels)
27
Computer
GUMSTIX Not space qualified, needs testing on
operating temperature and radiation Runs on Linux Overo Earth
256 MB RAM256 MB FlashMicro SD card slotPower ~ 3WMass <100 g
28
GPS Antenna and Receiver
GPS Reciever (GPS-12-V1)
Position Accuracy 10 m
Velocity Accuracy 0.03 m/s
Mass 200 g
Power 1 W
Operating Temp -40C to 85 C
GPS Antenna (ANT-GPS)
Mass 82 g
Operating Temp -55C to 85C
29
Total Requirements
Mass 887 g
Power 6.4 W
Operating Temp -20C to 60C1
1does not include computer and camera
30
Nathan Jones – Group LeaderKenneth BarnesChristina DaughtreyBrandt KingBrian KuehnerTimothy LoweryJared Wissel
Propulsion and Attitude Control
31
Attitude SensingSensor Parameters
4 Sun Sensors Mass = 1.2 kgPower < 1.2 W
Magnetometer Mass = 0.14 kgPower < 300 mW
Sun Sensor
Magnetometer
32
Torque RodsMTR-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.
Attitude Control
33
Control Torque
Disturbing Torques Gravity gradient Solar pressure Magnetic field
variations Atmospheric drag
Control Torque
τ μ B
212.4 A-mB 30 μTOrbit inclination 30°
μ
, 1.217e-4 N-md totτ
sin(i)1.860e-4 N-m
rodτ μ B
Time to Rotate 90°
2 2
2
1( )
122.025 kg-m
I m w h
,
25.24 min
rod d tot
It
34
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
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
Lightband Separation System
36
Orbital Requirements
Parameter Value
Altitude 350 km
Inclination 30o
Ballistic Coefficient 58 – 82 kg/m2
Orbit Lifetime 30 – 200 days
Inclination comes from assuming we launch due east from Cape Canaveral, which has a latitude of 28.5o.
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.
37
Summary
Parameter Required With 30% Contingency
Power 4.2 W 5.5 W
Mass 2.8 kg 3.7 kg
Settling Time 5.24 min
ESPA Requirements Max Value
Mass < 97 kg
Size 24in x 20in x 28in
Parameter Value
Altitude 350 km
Inclination 30o
Ballistic Coefficient 58 – 82 kg/m2
Orbit Lifetime 30 – 200 days
38
Erica Furnia – Group LeaderBrian AtteberryWesley FiteJason YorkStephen OehlerJennifer WellsAltay Zhakatayev
Structures, Thermal and Power
39
Spacecraft Structure
40
Current Power BreakdownUniversal Assumptions:• 30 LOS passes total, 7 mins/pass• Gumstick computers used (~2-6 Watts each)
Subsystem Power (W) Time (h) W-h Battery
Magnetometers 0.3 3.5 1.05 Lithium Monoflouride
Sun Sensors 0.4 3.5 1.4 Lithium Monoflouride
Torque Rods 3 3.5 10.5 Lithium Monoflouride
Computer 6 Varies (see Scenarios) Varies (see Scenarios) Lithium Monoflouride
Transmission 0.02 3.5 0.07 Lithium Monoflouride
Camera 1.8 3.5 6.3 Lithium Thionyl Chloride
Tribo Gun 100 1 1 Lithium Thionyl Chloride
Computer Runtime Lithium Thionyl Chloride (kg)
Lithium Monoflouride (kg) % of Total Mass Budget % Total Mass Budget (30% error margin)
1 week 0.5 5.5 11.9 15.5
2 weeks (chosen) 0.5 11.0 22.8 29.6
3 weeks 0.5 16.4 33.7 43.8
4 weeks 0.5 21.8 44.6 58.0
Battery Scenarios
41
ESPA Constraints
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
ESPA Constraints
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
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
Gs (W/m^2) 1420
Earth albedo 0.36
H (km) 350
emissivity of Earth 0.8
Earth IR (W/m^2) 244 (Max Hot), 218 (Max Cold)
Qint (W) 200 (Max Hot), 20 (Max Cold)
44
Optimize Tmax and Tmin
Option 1 ε affects only Tmin, while α affects both temperatures Backward calculations: εd=0.32, αd=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
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 • αx=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 Tmax and Tmin
Cont.
46
Mass Budget Percentage
Albedo Change Demo
Telecommunications andInstrumentations
Propulsion and Attitude Control
Structure, Thermal, and Power
Total Mass BudgetMass [kg]
Structure, Thermal, and Power Mass Budget
Mass [kg]
Albedo Change Demo 5.9 Housing (Satellite Shell) 19.1Telecommunications and Instrumentations 1.0 Mounting Hardware 0.8
Propulsion and Attitude Control 2.8 Thermal Insulation 1.0
Structure, Thermal, and Power 35.6 Batteries 14.8
Total 45.3 Total 35.6
Total (with 30% safety factor) 58.9 Total (with 30% safety factor) 46.3
47
Andrei Kolomenski – Group LeaderDanielle FitchCory PhillipsKris KeiserScott Southwell
Budget and Scheduling
48
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
49
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.
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:
54
Parametric Cost Budget
RDT&E Cost, 1st Unit Cost and SubTotal Cost values are expressed in thousands of dollars
55
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
56
References1. Roos, Achim, and Patrick Schmid. "Flash SSD Update: More Results, Answers." 14 Jan. 2008. Web. 28
Oct. 2009. <http://www.tomshardware.com/reviews/ssd-hard-drive,1968-11.html>.
2. "A Small, High-Torque Reaction/Momentum Wheel." Goddard Space Flight Center-Innovative Partnerships Program Office. NASA, 11 Apr. 2005. Web. 02 Nov. 2009. <http://ipp.gsfc.nasa.gov/ft-tech-reaction-moment-whl.html>.
3. "Reaction control system." Wikipedia, the free encyclopedia. 29 Sept. 2009. Web. 02 Nov. 2009. <http://en.wikipedia.org/wiki/Reaction_control_system>.
4. "Pegasus." Orbital Sciences Corporation. 2009. Web. 02 Nov. 2009.<http://www.orbital.com/SpaceLaunch/Pegasus/>.
5. Wertz, James R., and Wiley J. Larson, eds. Space Mission Analysis and Design, 3rd edition (Space Technology Library) (Space Technology Library). 3rd ed. New York: Microcosm, 1999. Print.
6. Antenna, Transceiver, GPS – spacequest.com
7. VHF downlink / UHF uplink transceiver - http://www.cubesatshop.com/index.php?page=shop.product_details&category_id=5&flypage=flypage.tpl&product_id=10&option=com_virtuemart&Itemid=1&vmcchk=1&Itemid=1
8. “Reliable Glass Capacitor Chosen by NASA for More Than 50 Years,” Microwave Product Digest<http://www.mpdigest.com/issue/Articles/2005/june2005/avx/Default.asp>
9. “Glass Capacitor Chosen by NASA for Over 50 Years”, Channel E: Magazine for Electronicshttp://www.channel-e.biz/design/articles/glasscapacitors.html
10. Camera - University of Leicester CubeSat Project . <http://cubesat.wikidot.com/lenscamera>
11. Camera – Malin Space Science Systems. <http://www.msss.com/camera_info/index.html>
12. "Available Subsystems." Surrey Space Technologies LTD. N.p., 2008. Web. 4 Dec. 2009. <http://www.sstl.co.uk/Products/Subsystems/Available_Subsystems>.
13. Stavast, Vann M., et al. Adapter Ring for Small Satellites on Responsive Launch Vehicles . N.p., n.d. Web. 4 Dec. 2009. <http://aeweb.tamu.edu/aero489/426.Fall.09/ESPA%20Adapter%20Ring.pdf>.
14. Holemans, Walter. The Lightband as Enabling Technology For Responsive Space . N.p., n.d. Web. 4 Dec. 2009. <http://aeweb.tamu.edu/aero489/426.Fall.09/Lightband%20separation%20systems.pdf>.