foresight: designing a radio transponder mission to … and foresight − apophis, ... and total...

Copyright 2009, SpaceWorks Engineering, Inc. (SEI) | www.sei.aero1

Foresight: Designing a Radio Transponder Mission to Near Earth Asteroid Apophis

Mr. A.C. CharaniaPresident, Commercial Division | SpaceWorks Engineering, Inc. (SEI) | [email protected] | 1+770.379.8006

Dr. John R. OldsCEO/Principal Engineer | SpaceWorks Engineering, Inc. (SEI) | [email protected] | 1+770.379.8002

Mr. Jesse KoenigSystems Engineer | Sierra Nevada Corporation, Inc. | [email protected]

Version A | 28 April 2009 | 2009 IAA Planetary Defense Conference, 27-30 April 2009 (Granada, Spain)


F I R M

SpaceWorks Engineering, Inc. (SEI) | www.sei.aero3

FIRM AREAS OF ENGAGEMENT


SAMPLE CONCEPT STUDIES


SPACEWORKS SOFTWARE PRODUCT LINE


PREVIOUS PLANETARY DEFENSE WORK

- Schaffer, M. G., Charania, A., Olds, J. R., "Evaluating the Effectiveness of Different NEO Mitigation Options," AIAA-2007-P2-1, 2007 Planetary Defense Conference, Washington, D.C., March 5-8, 2007.

- Olds, J. R., Charania, A., Schaffer, M. G., "Multiple Mass Drivers as an Option for Asteroid Deflection Missions," AIAA-2007-S3-7, 2007 Planetary Defense Conference, Washington, D.C., March 5-8, 2007.

- Charania, A., Graham, M., Olds, J. R., "Rapid and Scalable Architecture Design for Planetary Defense," AIAA-2004-1453, 1st Planetary Defense Conference: Protecting Earth from Asteroids, Orange County, California, February 24-27, 2004.


I N T R O D U C T I O N


THE PLANETARY SOCIETY’S APOPHIS MISSION DESIGN COMPETITION

The SpaceWorks Engineering / SpaceDev team thanks The Planetary Society (its directors and its members and specifically Dr. Louis D. Friedman and Mr. Bruce Betts) for the opportunity to present the Foresight design and increase public awareness of the potential planetary threat from Near Earth Objects (NEOs) through the Apophis Mission Design Competition (Foresight: 1st place overall). Special thanks are extended to Mr. Dan Gerachi for his leading financial support for this endeavor.


OVERVIEW

- Foresight spacecraft is a small satellite mission design to orbit Near Earth Object (NEO) Apophis

- Primary mission: reduce future orbital uncertainty of Apophis- Over a span of 300 days reduces the ±3σ error ellipse of Apophis' trajectory ("keyhole" or b-

place encounter) in 2029 to 6.0 kilometers by 2017 (from 4500 km today)- Purpose-designed to meet minimum requirements of The Planetary Society’s 2007 Apophis

Mission Design Competition (1st place overall international winner)- Characteristics

- Small orbiter spacecraft with minimal instruments and complexity- Foresight’s Encounter Spacecraft (ES): 220 kg (wet mass), ~85 cm cube (stowed)- Total launch mass with Propulsive Transfer Vehicle (PTV): 1,608 kg (wet mass with payload)

- Lean, low risk small satellite approach to design and manufacture- Foresight uses heritage components, instruments, and flight proven technologies- Proven mission approach with heritage from NEAR (Eros) and Hayabusa (Itokawa) missions- Low cost launch vehicle (Minotaur IV baseline, other options available)- Total life-cycle cost estimated to be under $131M USD (including launch)

- Flexible- Multiple launch windows between 2012 and 2014 (extended mission option)


FORESIGHT ANIMATION

Foresight: A Radio Beacon Mission to Asteroid ApophisLINK: http://www.youtube.com/watch?v=HRAo1dN7zMc

CHANNEL: SpaceWorksEng


M I S S I O N O V E R V I E W

Copyright ©2008, SpaceWorks Commercial, A Division of SpaceWorks Engineering, Inc. (SEI) All Rights Reserved12

CURRENT AND MISSION OUTPUT ERROR ELLIPSE FOR APOPHIS

Note: not to scale

There is a 600 m long “keyhole” somewhere in the current 4500 km position ellipse. If Apophis goes through this region of space during its close approach in 2029, in 2036 it will hit the Earth.

Sources: "How Dangerous are Near-Earth Asteroids?," Clark R. Chapman, Southwest Research Institute Boulder, Colorado, USA, 2007 Space Weather Workshop Reception, After-Dinner Talk, UCAR Center Green Campus, Bldg. 1, 25 April 2007.


APOPHIS AND FORESIGHT

− Apophis, a 270 meter wide near-Earth asteroid, will pass within the range of geostationary satellites during its close approach to the Earth in April 2029

− According to S. R. Chesley in 2005, the 2029 Apophis-Earth encounter distance is predicted to be 5.89 Earth radii, ± 0.35 Earth radii, (3σ)

− The 2029 close approach will significantly alter Apophis’ orbit− There is a small possibility that the asteroid will pass through an approximately 600 m

wide area of space called a “keyhole”, which would cause Apophis to impact the Earth in 2036

− More precise measurements of the orbit of Apophis can confirm or deny this possibility

− Today, we have a 99.7% confidence (±3 sigma error) that in 2029 Apophis will be within a 4500 km window (with the 600 m “keyhole” somewhere in this window)

− Results in a 1 in 45,000 chance of Apophis impacting the Earth in 2036

− After 300 days of orbiting Apophis, the Foresight mission will reduce the size of this window to approximately 6 km window (Goal in the Planetary Society Competition was 14 km)

− This will help determine whether Apophis will pass through the keyhole in 2029 and subsequently impact the Earth when it comes back in 2036


FORESIGHT SPACECRAFT CONFIGURATION

ENCOUNTER SPACECRAFT (ES)ENCOUNTER SPACECRAFT (ES) WITH PROPULSIVE TRANSFER STAGE (PTV)


FORESIGHT MISSION OVERVIEW

− Foresight spacecraft: concept design for radio tagging mission to Near Earth Asteroid (NEO) Apophis

− Designed to meet requirements of 2007 Planetary Society Apophis Mission Design Competition)− Goal: Apophis tracking accuracy must be adequate to reduce the long dimension of the ±3σ error

ellipse to 14 kilometers by 2017, for reference, this translates approximately to a 10% impact probability if the keyhole is right in the middle of the 14-kilometer error ellipse.

− Joint team design with SpaceWorks Engineering, Inc. (SEI) and SpaceDev, Inc.

− Low-cost, low-risk, robust, minimal science mission to obtain accurate tracking information− Leverages off the shelf technologies, incorporating leaner approaches to spacecraft

design

− Launch on Orbital Sciences Corporation (OSC) Minotaur IV (Wallops Island, Virginia USA)− Five launch windows have been identified spanning years 2012 to 2014− Chemical propulsive transfer vehicle to perform outbound burn to Apophis (3,600 m/s) with Foresight

encounter spacecraft performing portion of Earth departure, and Apophis capture burn (total less than 2,400 m/s)

− Foresight spacecraft mass is 220 kg (propulsive transfer vehicle of 1,387 kg)− Foresight orbiting spacecraft powered by solar arrays augmented by rechargeable batteries (280.6 W

EOL); transfer vehicle is powered by onboard batteries − The Spacecraft has two main instruments, a multi-spectral imager and laser altimeter− Over a span of 300 days reduces the ±3σ error ellipse of Apophis' trajectory ("keyhole"

or b-place encounter) in 2029 to 6.0 kilometers by 2017− The total cost for this mission is estimated to be $130.9 M ($87.9 M for spacecraft and instrument

development and acquisition, $21 M for operations, and $22 M for the launch vehicle)− Overall system reliability is estimated to be 90.2%


FORESIGHT MISSION PROFILE


LAUNCH WINDOW (1)

0

1,000

2,000

3,000

4,000

5,000

6,000

7,000

8,000

9,000

10,000

11,000

12,000

1/2011 7/2011 1/2012 7/2012 1/2013 7/2013 1/2014 7/2014 1/2015

Departure Date

Del

ta V

[m/s

]

Total Delta VDeparture Delta VArrival Delta V

Launch Window 1

Launch Window 2

Launch Window 3

Launch Window 4

Launch Window 5

Total Delta V Limit

Departure, Arrival, and Total Delta-V for Minimum Total Delta-V Trajectories from LEO to Apophis.


LAUNCH WINDOW (2)

0

500

1,000

1,500

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

6,500

7,000

7,500

8,000

-5 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 100

105

110

115

120

125

130

135

140

145

150

155

160

165

170

175

180

Departure Date

Del

ta V

[m/s

]

Total Delta V

Departure Delta V

Arrival Delta V

Arrival Delta V Limit

Total Delta V Limit

53 Days 13 Days 43 Days 18 Days 7 Days

PrimaryWindow

Alternate Window A

SecondaryWindow

Alternate Window B

Alternate Window C

Optimal Launch Dates

Nominal Mission Launch Date

4/17

/201

2

5/9/

2012

6/8/

2012

12/1

8/20

12

12/2

5/20

12

12/3

0/20

12

4/5/

2013

4/20

/201

3

5/17

/201

3

2/3/

2014

2/10

/201

4

2/20

/201

4

4/10

/201

44/

13/2

014

4/16

/201

4

Departure, Arrival, and Total Delta-V for Minimum Total Delta-V Trajectories fromLEO to Apophis for Specified Launch Windows


FORESIGHT LAUNCH OPPORTUNITY AND TIMELINE

Minimum Total Delta-V Earth-Apophis Transfer Trajectories for each Launch Window

No. Mission PhaseDate

(initial)

1. Launch 5/9/2012

2. Earth Departure: PTV Maneuver 5/9/2012

3. Stage Separation 5/9/2012

4. Commissioning 5/9/2012

5. Cruise 6/8/2012

6. Trailing Capture Maneuver 3/15/2013

7. Initial Survey 3/15/2013

8. Apophis Capture Maneuver 3/25/2013

9. Observation 3/25/2013

10. Apophis Withdraw Maneuver 4/24/2013

11. Tracking 4/24/2013

12. Extended Mission 2/18/2014

Mission Timeline for Primary Launch Date


FORESIGHT DELTA-V BUDGET: ENCOUNTER SPACECRAFT AND PROPULSIVE TRANSFER VEHICLE

3,600 m/s3,600 m/s3,600 m/s3,600 m/s3,600 m/s

90 m/s90 m/s 90 m/s 90 m/s 90 m/s

533 m/s 638 m/s 224 m/s

1,231 m/s

1,851 m/s

2,047 m/s 766 m/s

2,144 m/s

814 m/s

176 m/s513 m/s

1,156 m/s

192 m/s 515 m/s

0

500

1,000

1,500

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

6,500

Primary Alternate A Secondary Alternate B Alternate C

Launch Window

Del

ta V

[m/s

ES RemainingES Arrival BurnES Departure BurnPTV ReservePTV Departure Burn

Spacecraft and Propulsive Transfer Vehicle (PTV) Delta-V Budget for Optimum Launch Dates within each Launch Window


FORESIGHT DELTA-V BUDGET: ENCOUNTER SPACECRAFT

533 m/s638 m/s

224 m/s

1,231 m/s

1,851 m/s

2,047 m/s 766 m/s

2,144 m/s

814 m/s

60 m/s

60 m/s

60 m/s

60 m/s

60 m/s

100 m/s

100 m/s

100 m/s

100 m/s

100 m/s

16 m/s

353 m/s

996 m/s

32 m/s

355 m/s

0

250

500

750

1,000

1,250

1,500

1,750

2,000

2,250

2,500

Primary Alternate A Secondary Alternate B Alternate C

Launch Window

Delta

V [m

/s]

ReserveManeuveringStation-keepingArrival BurnDeparture Burn

Encounter Spacecraft Delta-V Budget for Optimum Launch Dates within each Launch Window


FORESIGHT SUN-EARTH ANGLES


ORBIT DETERMINATION METHOD

− Internal software tools were developed by the Foresight team to accurately propagate Apophis’s orbit state and predict the uncertainty in Apophis’ state as a function of number of measurements and time between measurements

− An 8th/9th order n-body numerical propagator with a variable step size was used to propagate the actual and dispersed orbits of Apophis forward from a given state and epoch

− The Sun, all of the planets and the Earth's moon are considered in the gravitational model. The perturbing effects of the large asteroid-belt asteroids Ceres, Pallas, and Vesta are also included. Solar pressure and the Yarkovsky effect are not modeled, but their associated uncertainties are addressed in the analysis

− For a given starting condition, the propagator’s step-wise integration tolerances were set so that results for position accuracies were on the order of a few meters in 2029

Range Measurements, ρmeas

Calculate Objective Function, f

Propagate X0_p

Set New Initial State, X0_p

Guess Initial State, X0_p

Converged?

X0 = X0_p

Yes

No

Optimizer

Range Measurements, ρmeas

Calculate Objective Function, f

Propagate X0_p

Set New Initial State, X0_p

Guess Initial State, X0_p

Converged?

X0 = X0_p

Yes

No

Optimizer


DECREASE IN ERROR ELLIPSE WITH OBSERVATION TIME

0.1

1

10

100

1000

4/21/2013 6/20/2013 8/19/2013 10/18/2013 12/17/2013 2/15/2014 4/16/2014Date

Max

imum

202

9 Er

ror E

llips

e D

imen

sion

(km

)

14 km Target

300

Day

Dur

atio

n

Apophis Error Ellipse Reduction for Target Mission (With Fine Monte Carlo)


FORESIGHT MISSION: EFFECT ON OVERALL ERROR ELLIPSE

-8

-7

-6

-5

-4

-3

-2

-1

0

1

2

-12 -11 -10 -9 -8 -7 -6 -5 -4 -3 -2 -1 0 1 2

ξ-axis Position (Earth radii)

ζ-ax

is P

ositi

on (E

arth

radi

i)

InitialAfter MissionEarth

+/- 3σ = 6 km

-7.54000

-7.53500

-7.53000

-7.52500

-7.52000

-7.51500

-7.51000

-7.50500

-7.50000

-7.49500

-7.49000

-1.50500 -1.50000 -1.49500 -1.49000 -1.48500


ζ-ax

is P

ositi

on (E

arth

radi

i)

-8

-7

-6

-5

-4

-3

-2

-1

0

1

2

-12 -11 -10 -9 -8 -7 -6 -5 -4 -3 -2 -1 0 1 2


ζ-ax

is P

ositi

on (E

arth

radi

i)

InitialAfter MissionEarth

+/- 3σ = 6 km

-7.54000

-7.53500

-7.53000

-7.52500

-7.52000

-7.51500

-7.51000

-7.50500

-7.50000

-7.49500

-7.49000

-1.50500 -1.50000 -1.49500 -1.49000 -1.48500


ζ-ax

is P

ositi

on (E

arth

radi

i)

Initial and Final Position Error in 2029 after 300 Days of Tracking (B-Plane Error Ellipse Comparison)(Assuming no Additional Earth Observations after 2012)

Location Probability


S P A C E C R A F T

O V E R V I E W


FORESIGHT ENCOUNTER SPACECRAFT (ES)


FORESIGHT SUBSYSTEM COMPONENT SPECIFICATIONS


Component Name Manufacturer

No. on ES

No. on PTV Specifications

Propulsion S/C Main Engine22 Aerojet 445 Aerojet 1 0 Thrust (vac): 445 N, Isp: 309 s, T/W: 24.39

S/C RCS23 Aerojet 2 Aerojet 6 0 Thrust (vac): 2 N, Isp: 265 s, T/W: 0.75

PTV Main Engine24 R-40B Aerojet 0 1 Thrust (vac): 4000 N, Isp: 293 s, T/W: 56.4

PTV RCS25 Aerojet 21 Aerojet 0 4 Thrust (vac): 21 N, Isp: 285 s, T/W: 3.81, Quad configuration

Thermal Control

Heaters26 Kapton Heaters Minco 16 0 -200 to 200ºC range, Kapton/FEP material

Power

Batteries21 VES 180 Saft 6 2 Li-Ion space technology, specific energy: 165 Wh/kg, storage: 180 Wh each

Solar Array22 Triple Junction Spectrolab 2 0

GaInP2/GaAs/Ge, BOL power: 289 W/m2, BOL efficiency: 22.5%, EOL power: 256 W/m2, 4% degradation per year

Distribution PDU SpaceDev 1 1

16 5-Amp high side relays, Integrated 200-W Li-Ion battery charger, 96 12-bit ADCs, Digital solar array peak power tracking

Command and Data Handling

CPU27 PowerPC 750 FX IBM 1 0

RISC Microprocessor, 1856 MIPS at 800 MHz with 256 MB RAM, RS-422 / USB / Ethernet compatible

Memory28 16 GB SSD Samsung 2 0 NAND-based SSD, read rate: 57 MBps, write rate: 32 MBps

SCC29 8051 Silicon Labs 0 1 1000 MIPS @ 100 MHz, 128 KB Flash, 8448 bytes data RAM, 8 12-bot ADCs, 2 12-bit DACs

Attitude Determination and Control Sun Sensor30 MSS AeroAstro 12 0 60º FOV, accuracy ±1º

Star Tracker31 HE-5AS Terma 2 0 22º FOV, <1 arcsec cross-track accuracy, 5 arcsec boresight accuracy

Reaction Wheel32 MicroWheel 1000 Dynacon 4 0

Produce 30 mNm torque, hold 1000 mNm angular momentum, mounted with 1 each on XYZ axes and 1 on skew axis

IMU33 LN-200S Northrop Grumman 1 1

Fiber Optic Gyro, silicon accelerometers and electronics

Communications Low Gain Antenna34 Custom Ball Aerospace 2 0 S-Band, 50 bps data rate

High Gain Antenna34 Custom Ball Aerospace 1 0 X-Band, 17 kbps data rate at 0.5 AU, SNR: 3, efficiency: 55%

X-Band Transponder35 SDST

General Dynamics 1 0

DSN Compatible, X-Band transmit and receive, 2.0 dB Noise Figure, -157.7 dBm Receiver Threshold, 10 ns Ranging Delay Variation, 0.5 ns Carrier Delay Variation

S-Band Transceiver36

Multi-Mode S-Band Transceiver

General Dynamics 1 0

DSN Compatible, S-Band transmit and receive, < 2.5 dB Noise Figure, Delay Variation, 0.5 ns Carrier Delay Variation


MASS STATEMENT: FORESIGHT ENCOUNTER SPACECRAFT (ES)


No. Name Element Mass [kg]

Subsystem Mass [kg]

1.0 Structures and Mechanisms 26.0 Primary Structure 14.8 Secondary Structure 5.5 Fuel Tank 2.3 Oxidizer Tank 2.3 Pressurant Tank 0.7 Solar Array Support Structure 0.2 Solar Array Actuators 0.2

2.0 Propulsion 7.0 Main Engine: Aerojet 445 1.9 Main Engine Feed Lines 1.9 Maneuvering Engines: Aerojet 2 (x6) 1.6 Maneuvering Engine Feed Lines 1.6

3.0 Thermal Control 5.0 Reflective Foil 1.0 Multi-Layer Insulation 3.8 Heaters 0.2

4.0 Power 12.1 Batteries: Saft VES 180 (x6) 6.7 Solar Array: Spectrolab Triple Junction 2.1 Power Distribution Unit 1.7 Power Cabling 1.6

5.0 Command and Data Handling 4.9 CPU: PowerPC 750FX 0.1 Memory: Samsung 64 GB Solid State Drive (x2) 0.1 Electronics Module 1.0 Wiring 3.7

6.0 Attitude Determination and Control 12.9 Sun Sensors: AeroAstro MSS (x6) 0.4 Star Sensors: Terma HE-5AS (x2) 4.4 Reaction Wheels: Dynacon MicroWheel 1000 (x4) 6.6 Inertial Measurement Unit: LN-200S 1.5

7.0 Communications 8.9 High Gain Antenna 3.0 Low Gain Antenna (x2) 0.7 Small Deep Space Transponder 2.9 Multi-Mode S-Band Transceiver 2.3

8.0 Margin (20%) 13.6 9.0 Dry Mass 90.2

10.0 Consumables 120.1 Fuel: MMH 45.1 Oxidizer: NTO 74.4 Pressurant: He 0.7

11.0 Wet Mass 210.3

12.0 Payload 10.0 Advanced Imagery Mechanism (AIM) 5.0 Laser Altimeter Device (LAD) 5.0

13.0 Gross Mass 220.3

No. Element Name Mass [kg]

1.0 Structures and Mechanisms 26.0

2.0 Propulsion 7.0

3.0 Thermal Control 5.0

4.0 Power 12.1

5.0 Command and Data Handling 4.9

6.0 Attitude Determination and Control 12.9

7.0 Communications 8.9

8.0 Margin (20%) 13.6

9.0 Dry Mass 90.2

10.0 Consumables 120.1

11.0 Wet Mass 210.3

12.0 Payload 10.0

13.0 Gross Mass 220.3

MASS BREAKDOWN STATEMENT: ENCOUNTER SPACECRAFT (ES)


MASS STATEMENT: PROPULSIVE TRANSFER VEHICLE (PTV)


No. Name Element Mass [kg]

Subsystem Mass [kg]

1.0 Structures 133.6 Primary Structure 18.8 Secondary Structure 13.9 Payload Adapter 5.5 Fuel Tank 22.5 Oxidizer Tank 22.7 Pressurant Tank 50.2

2.0 Propulsion 31.8 Main Engine: Aerojet R-40B 6.8 Main Engine Feed Lines 6.8 RCS Engines: Aerojet 21 (x16) 9.1 RCS Engine Feed Lines 9.1

3.0 Thermal Control 12.1 Reflective Foil 2.4 Multi-Layer Insulation 9.7

4.0 Power 6.2 Batteries: Saft VES 180 (x2) 2.2 Power Distribution Unit 0.9 Wiring 3.1

5.0 Command and Data Handling 1.6 Spacecraft Control Computer 0.1 Electronics Module 1.0 Wiring 0.5

6.0 Attitude Determination and Control 1.5 Inertial Measurement Unit: LN-200S 1.5

7.0 Margin (20%) 37.4 8.0 Dry Mass 224.2 9.0 Consumables 1,163.1

Fuel: MMH 426.1 Fuel Reserves / Residuals 10.5 Oxidizer: NTO 702.7 Oxidizer Reserves / Residuals 17.3 Pressurant: He 6.5

10.0 Wet Mass 1,387.3 11.0 Payload 220.3

Foresight Spacecraft 220.3 12.0 Gross Mass 1,607.6

No. Element Name Mass [kg]

1.0 Structures 133.6

2.0 Propulsion 31.8

3.0 Thermal Control 12.1

4.0 Power 6.2

5.0 Command and Data Handling 1.6

6.0 Attitude Determination and Control 1.5

7.0 Margin (20%) 37.4

8.0 Dry Mass 224.2

9.0 Consumables 1,163.1

10.0 Wet Mass 1,387.3

11.0 Payload 220.3

12.0 Gross Mass 1,607.6

MASS BREAKDOWN STATEMENT:PROPULSIVE TRANSFER VEHICLE (PTV)


FORESIGHT ENCOUNTER SPACECRAFT (ES) CONFIGURATION


FORESIGHT SPACECRAFT DIMENSIONS AND LAUNCH VEHICLE PACKAGING

309 cm

462 cm

549 cm

205 cm

Foresight Encounter Spacecraft and PTV

Foresight Encounter Spacecraft and PTV in Minotaur IV Payload Fairing


COST AND RELIABILITY


Cost Element Name DDT&E [$US FY2007]

Acquisition Cost [$US FY2007]

Total Cost [$US FY2007]

Spacecraft Stages $57.85 M $23.74 M $81.59 M

Foresight Encounter Spacecraft $23.51 M $14.77 M $38.28 M

Propulsive Transfer Vehicle $34.34 M $9.27 M $43.61 M

Scientific Instruments (2) ----- $6.30 M $6.30 M

Operations ----- $20.99 M $20.99 M

Launch Vehicle: Minotaur IV ----- $22.00 M $22.00 M

Total $57.85 M $73.03 M $130.88 M

5.2%

0.8%

3.2%

0.5%

0% 2% 4% 6% 8% 10% 12%

LOM

Contribution of Architecture Elements (%)

Launch VehiclePropulsive Transfer Vehicle (PTV)Encounter Spacecraft (ES)Instruments

Contribution of Architecture Elements to Loss of Mission (LOM)

0

50

100

150

200

250

88.6% 89.0% 89.3% 89.6% 90.0% 90.3% 90.6% 90.9% 91.3% 91.6%

Probability of Mission Success (%)

Freq

uenc

y

Mean: 90.2 %90th: 89.4 %

Histogram of Reliability Results(20,000 Monte Carlo Trials)

Life Cycle Cost Statement


FORWARD WORK AND THOUGHTS

− Potential future analysis− Refinement of baseline concept (alternative tank designs, ∆V split, etc.)− Update to Apophis knowledge since initial design− Analysis of alternative instrument suite− Alternative target− Integration with Sierra Nevada Corporation (SNC) low cost bus approaches− Examine commonalties and alternatives for NASA Ames common modular bus− Potential small satellite technology advancement (funding: NASA SBIR/STTR, etc.)

− Other thoughts− ESPA ring designs may offer potential cost savings for asteroid missions, but schedule

and launch integration issues may affect overall benefits− Transponder missions do not have to be technology demonstration missions− Potential leveraging of multiple funding sources (multiple end users, the U.S. –

DoD/NASA/DHS) for actual mission− Consider use of Falcon 1e: 1,010 kg to LEO for $11.27 M (FY2009, Q4)

− LEO: 185 km circular orbit launched due east (9.1 deg)− 2009 IAA Planetary Defense Conference White Paper Notes

− Planetary defense is multilayered response− Low cost non-science focused reconnaissance missions are one part of an overall

response strategy (multiple types of missions: transponder, thermal IR, etc.)− Potential need to prioritize order of data required for reconnaissance missions (radio

science before thermal IR for instance or in what combination)− Leverage existing global small satellite community (already in progress in various

ways)


w w w . s e i . a e r o

Business Address:SpaceWorks Engineering, Inc. (SEI)

1200 Ashwood Parkway, Suite 506, Atlanta, GA 30338 U.S.A.Phone: 1+770.379.8000 | Fax: 1+770.379.8001 | www.sei.aero | [email protected]

foresight: designing a radio transponder mission to … and foresight − apophis, ... and total...

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