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Apophis Mitigation Technology Flight Test Design Review

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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

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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.

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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

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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

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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.

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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.

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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)

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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

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Frans Ebersohn – Group LeaderAustin BondBrannen ClarkJoshua HempelJulianne LarsonAgustin MaquiAndrew Schaeperkoetter

Albedo Change Demo

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SPADE – Static Preliminary Albedo Demonstration Experiment

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Experiment Mission Profile

1. Orient spacecraftV

2. Charge test surface then remove power supply

3. Initiate tribogun and spray test surface

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Experiment Mission Profile (cont)

4. Allow particles time to cure

5. Observe, record, and transmit data

Data

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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

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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

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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

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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

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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

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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

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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

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Minimum Imaging Requirements

Image Rate (fps)

Total Frames

During Spray

1 276

During Curing (30 mins)

0.0167 30

Mission Total 306

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Travis Jacobs – Group LeaderDarkhan AlimzhanovJonathan EllensPatrick HarringtonCathy SpohnheimerBarrett Wright

Telecommunications and Instrumentation

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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°

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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

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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)

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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)

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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

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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

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Total Requirements

Mass 887 g

Power 6.4 W

Operating Temp -20C to 60C1

1does not include computer and camera

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Nathan Jones – Group LeaderKenneth BarnesChristina DaughtreyBrandt KingBrian KuehnerTimothy LoweryJared Wissel

Propulsion and Attitude Control

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Attitude SensingSensor Parameters

4 Sun Sensors Mass = 1.2 kgPower < 1.2 W

Magnetometer Mass = 0.14 kgPower < 300 mW

Sun Sensor

Magnetometer

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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

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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

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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

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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

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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.

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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

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Erica Furnia – Group LeaderBrian AtteberryWesley FiteJason YorkStephen OehlerJennifer WellsAltay Zhakatayev

Structures, Thermal and Power

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Spacecraft Structure

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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

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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

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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

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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)

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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

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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.

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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

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Andrei Kolomenski – Group LeaderDanielle FitchCory PhillipsKris KeiserScott Southwell

Budget and Scheduling

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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

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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.

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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:

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Parametric Cost Budget

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

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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

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