Title: A Mission Template for Exploration and Mitigation of Hazardous Near-Earth Asteroids
1A Mission Template for Exploration and Mitigation
of Hazardous Near-Earth Asteroids
- D. C. Hyland
- Professor of Aerospace Engineering, College of
Engineering - Professor of Physics, College of Science
- Dr. M. Rao
- DuPont Fellow and Lavoisier Distinguished
Scientist - And Aggies Against Apophis
- R. Margulieux, J. Doyle, J. Sandberg, B. Young,
X. Bai, B. Kuehner, J. Lopez, A. D. Ravidas, N.
Satak, S. Bourgeois, T. Elgohary, - Z. Levin, V. Vasiraju, L. McElduff, R.
Rangarajan, and S. Ge
2 we live in a neighborhood bustling with
fragments (asteroids, comets) left over from the
formation of the solar system
Some of these (Near Earth Asteroids, or NEAs)
follow orbits close to the Earths orbit. The
particularly hazardous Aten class of NEAs have
Earth-crossing orbits that periodically take them
near our planet. There is much evidence that NEA
impacts have caused a sequence of mass extinction
events
2
3History of our project for Asteroid 99942 Apophis
- Learning Through Research (LTR) educational
program at Texas AM gave rise to the project - Fall 07 inaugurated a sequence of three
undergraduate classes addressing the design of
the Apophis Preliminary Exploration Platform
(APEP). Objectives were - Rendezvous with Apophis in 2013, act as a beacon
to enable precise tracking, - Measure physical characteristics most relevant to
improved accuracy of orbit prediction - In spring of 2008, King Abdulaziz City for
Science and Technology (KACST) offered to
participate and provide the science instruments. - AM design team presented results at NASA-ARC
Oct. 2008. Our three-party collaboration was born
4History of our project (cont.)
- Fall 08 design class was challenged to develop a
system that can significantly modify the orbit of
Apophis before its close approach in 2036 - The Deflect Apophis System (DAS) was intended to
be a follow-on to the APEP mission - Lessons Learned
- Too late for 2012 Launch of APEP
- DAS showed that mitigation mission requires
little added complexity beyond the APEP
exploration mission - Aggies Against Apophis (AAA) research group
formed to define the Apophis Exploration and
Mitigation Platform (AEMP) mission for launch in
2020-2022. - Preliminary design presented at NASA-Ames in
April, 2009 - Proposal submitted in response to the National
Research Councils AO - Overall results
- 1. Strategy for executing necessary exploration
and tracking and performing mitigation within a
single mission - 2. Initial use of a relatively temporary
mitigation method that requires minimum
scientific data, followed by (once there is
sufficient scientific data) - 3. A permanently acting mitigation method
5Apophis 99942 Background Information
- 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, occurring roughly every 7
years. When Apophis passes by in 2029, it could
pass through a gravitational keyhole which could
could swing Apophis in a collision with Earth in
2036 - As an Aten class NEA, Apophis rated as high as 4
on the Torino scale but has since been downgraded
to zero. - Apophis Greek name for the Egyptian god Apep,
the god of chaos and destruction
Source http//neo.jpl.nasa.gov/apophis
6
6Apophis 99942 Physical Data
Orbital Elements at Epoch 2454200.5 (2007-Apr-10.0) TDB Ref JPL 139 (heliocentric ecliptic J2000 Orbital Elements at Epoch 2454200.5 (2007-Apr-10.0) TDB Ref JPL 139 (heliocentric ecliptic J2000 Orbital Elements at Epoch 2454200.5 (2007-Apr-10.0) TDB Ref JPL 139 (heliocentric ecliptic J2000
Element Value Units
e 0.191059415
a 0.922261415 AU
q 0.746054689 AU
i 3.331314642 deg
node 204.4591523 deg
peri 126.3855713 deg
M 307.3630785 deg
tp 2454247.800689 (2007-May-27.30007) JED
period 323.5038741 d
n 1.112815112 deg/d
Q 1.098468142 AU
Physical Characteristics Physical Characteristics Physical Characteristics
Vimpact 12.59 km/s
Vinf 5.87 km/s
H 19.7
Diameter 0.27 km
Mass 2.78E10 kg
Energy 4.00E02 MT
- LL-chondrite composition
- Estimated rotational period of 30.57 h
- Keyhole Event April 13, 2029
- Results in resonant return and possible impact in
April 2036
Sources http//ssd.jpl.nasa.gov/sbdb.cgi?sstr999
42orb1cov0log0discovery
Chesley, Milani, Vokrouhlicky, Icarus 148,
118138 (2000)
7
7Best Fit Solutions From Tracking Data To-Date
- Close encounter on Friday, April 13, 2029
- Latest (optical) data extends to August 16, 2006
- Current best estimate solution using the
Standard Dynamical Model (SDM) has a nominal
Earth-centered Approach of 5.96 ? 0.08 R? - Close encounter on April 13 (Easter Sunday),
2036 - Current best SDM solution shows a distant 0.34
AU passage. - However, the ? 3? orbit set extends through 720
of heliocentric longitude, intersecting Earths
orbit ?2.4 ? from the center of the probability
distribution. - Note Existing dynamical uncertainties can
produce very large perturbations away from the
SDM solutions. The probability of a 2036 impact
is likely to be small (1 in 45000), but not zero
8Why consider Apophis?
- Apophis appears somewhat representative of the
mid-range (20 billion kg) of hazardous NEAs
able to cause regional destruction (400 MT of
energy) - Apophis is relatively easy to get to in the
not-too-distant future - There are favorable (low energy) launch windows
every 7yrs coinciding with Apophis close
approaches to Earth - Weve missed the 2012-2014 window, but can still
try for 2020-2022 - Although Apophis closely approaches Earth (esp.
in 2029 and then in 2036) it is very unlikely to
impact Earth - Good The Apophis mission is intended to be a
dress rehearsal, not the real thing. - The aim is to flight-validate an archetypal
exploration/mitigation operation - The 2021 2023 mission allows us to measurably
change the orbit (thus verifying the
technologies), but without producing
significantly harmful orbit perturbations. - Once our technologies are proven, we can be ready
for the real thing
9Systematic exploration and mitigation of NEAs A
gradualist approach
- 1. Start early, combine object exploration with
mitigation. - 2. Begin a mission with precision tracking,
tagging the asteroid with the spacecraft, and
combine this with science measurements of the
gravity field, material composition, and thermal
properties. - 3. Cross-correlate the tracking data found in the
initial exploration phase with SDM predictions
and resolve the modeling and parameter
uncertainties. - 4. As early as possible, perform some initial
mitigation technique that depends on the least
data (mass distribution, total mass,
center-of-mass location, geometric envelope, spin
state). - 5. Combine the initial mitigation phase with
continued observation to map the albedo, model
the thermal properties, model solar pressure, and
the Yarkovsky effect. - 6. Finally, having obtained sufficiently precise
model data, apply some permanent mitigation
technique that can eventually retire the threat
completely. - 7. Above all Do no harm!
10Combined Exploration/Mitigation Mission Timing
- Launch windows for minimum ?V rendezvous
approximately coincide close approach events
occurring every 7yrs. - If Apophis passes through a gravitational keyhole
in 2029, no technology exists capable of
forestalling Earth impact in 2036 - Mitigation should begin at least 6 to 7 years
previously. - Therefore the most appropriate launch window is
January March 2021, with arrival at Apophis in
September, 2021
11Dry mass 415 kg, wet mass 570 kg. Launch
mass to Earth escape 1100 kg, (includes
booster stage for rendezvous) Max power mode
Grav Tractor 1kW Payloads science instruments
and albedo change mechanism. Development time
5yrs. Total cost 350 M
12Exploration/Mitigation Mission Launch to
Rendezvous Station-Keeping
- Direct launch-to-transfer orbit using a Falcon-9
vehicle. - Star 30BP was chosen to be the kick motor.
- Main engine single propellant Hydrazine engine
manufactured by AeroJet. - Attitude control sensors comprise a star tracker,
an IMU, 6 sun sensors, 4 reaction wheels and
Aerojet 22 N thrusters. - Upon insertion into the orbit of Apophis, the
spacecraft is to take up a 2 to 3 km stand-off
position, making periodic close approaches to
collect science data
Launch date Feb 19-2021
Rendezvous Date Sep 14 2021
Time of flight 208 days
C3 4.3 km2/s2
?Vf 3 km/s
13 Why is Exploration Needed Before
Mitigation?(J. D. Giorgini, L. A. M. Benner, S.
J. Ostro, M. C. Nolan, and M. W. Busch.
Predicting the Earth encounters of (99942)
Apophis, Icarus 193 (2008) 1-19.)
Errors in the Standard Dynamical Model (SDM) can
produce tens of Earth radii positional errors
over the period 2029 - 2036
Uncertainties in Apophis thermal emission
parameters (e.g. bond albedo) can produce tens of
Earth radii errors over the period 2018-2036
14Exploration/Mitigation Mission Initial
Exploration Phase (precedes Gravity Tractor Phase)
- The first actions to be performed are designed to
achieve the following science objectives - 1) Determine the trajectory of Apophis 99942
with sufficient accuracy to establish the minimum
trajectory change that can guarantee no Earth
impact from the start of on-orbit operation
through the close approach of 2036 - (2.a) Study physical characteristics to refine
the orbit propagation models - (2.b) Study physical characteristics to refine
intervention procedures. - On-board Science Instruments
- Comm. system employed as a beacon, operating in
conjunction with ground-based tracking stations,
to determine the trajectory of Apophis - Laser range finders (LRFs) determine the
spacecraft-to-Apophis relative position (for
tracking purposes), as well as map the surface
geometry. - Navigational cameras and the LRFs establish the
spin state. Combined with tracking data during
close approach spacecraft maneuvers, the surface
geometry model determines the gravitational
model. On-board sensors also map the albedo of
the asteroid and accumulate spectroscopic data to
establish its material composition. - Slide 16 relates the science objectives to the
instruments and techniques used to operate them.
15Science Instruments
Framing Cameras (Dawn)
NSSDC ID 2007-043A-01
Resolution 0.3 m at a distance of 3 km
Mass 10.0 kg (5.0 kg per camera)
Average Power Consumption 24.0 W (12.0 W per camera)
Micro-Bolometer (from THEMIS, Mars Odyssey)
NSSDC ID 2001-014A-01
Spectral Range 6.2-15.5µm (190 - 410K)
Mass 11.2 kg
Power 14 W
Laser Range Finder (NEAR-Shoemaker)
NSSDC ID 1996-008A-04
Range 50km
Mass 5 kg
Power 16.5 W
16Exploration Objectives Mapped to Instruments and
Measurement Techniques
Required Science Instrument(s) Technique(s)
Track spacecraft to determine trajectory in ecliptic frame and relative to Earth Radio Science Range rate, DSN Interferometry for angular position, Doppler range rate measurements
Determine SC to Apophis relative position Laser Range Finder, OpNav Camera, Star Tracker Correlate image sequences and LRF measurements
Determine mass of Apophis Laser Range Finder, OpNav Camera, IMU, Radio Science Correlate LRF and acceleration measurements over multiple revolutions compute inertial estimates
Determine surface geometry OpNav Camera, LRF Correlate images of Apophis at different angles of regard
Determine bulk size and density OpNav Camera, LRF Correlate a series of images and LRF topography Density indicates solid body vs. rubble pile
Determine gravity field OpNav Cam, LRF, Radio Science, IMU Correlate imager and range data, inertial accelerations
Map albedo over the surface of the asteroid Op/Nav camera, Micro-bolometer Imagery
Determine average bond albedo OpNav camera, Micro-bolometer Model from Imagery at varying phases and correlate with surface mapping
Map surface temperature Micro-bolometer Correlate high resolution infrared images to hot and cold zones of Apophis
Determine the spin axis OpNav camera Correlate images to determine spin orientation
Determine surface composition Spectrometer Correlation imagery with known materials
Impart impulse upon Apophis LRF, OpNav Cam, Low-thrust propulsion Station keep at specified distance for specified time
Modify average bond albedo of Apophis Albedo Change System Deploy albedo change material over the surface
21
17Tracking and Mitigation Pose Conflicting
Requirements
Whereas, if we launch and do tracking just before
a close approach, our uncertainty will be smaller
but there will be no time for mitigation.
Tracking measurements
17
18Radio Science for Tracking Uncertainty Estimation
- We need to use of the spacecraft as a beacon, to
be tracked using standard radio science methods. - We have studied in detail the methods for
tracking with radio science, and the estimation
of position uncertainties (up to 2036) given
those measurements and the propagation of
uncertainties using the unscented Kalman filter
(UKF). - Measurements used for tracking with radio
science - Range to the target. Use pseudo-noise ranging,
see DSMS Telecommunications Link Design Handbook
(810-005), available from http//deepspace.jpl.nas
a.gov/dsndocs/810-005/ (A refinement is Delta
Differential One-way Ranging (?DOR) methods, but
these are not assumed here) - Range-rate. Using the 34- or 70-meter DSN dishes
and the Ka-band downlink. Described in detail in
DSN module 202. - Altitude and azimuth angular position. Measured
from the ecliptic. Uses standard radio very-long
baseline interferometry (VLBI) techniques,
determining the phase delay, and thus the angle,
calibrated against background quasar beacons as
described in DSN module 211. - Expected errors
Measurement Uncertainty (3?)
Range 12 meters
Range-rate 0.015 mm/s
Altitude 7 nrad
Azimuth 7 nrad
19Radio Science for Tracking Uncertainty Estimation
- An UKF is run to obtain an optimal estimate of
the asteroid position based on the previous
measurements and the assumed dynamics. - The UKF may not be necessary for the tracking
period, since the tracking measurements
themselves should be sufficient to keep error
growth within the linear regime however, it is
required for the error propagation in the
post-tracking period as the simpler EKF incurs
significant errors. - For the sake of simplicity, at this point, we do
not distinguish between the spacecraft position
and the asteroid position. The required
asteroid-to-spacecraft position accuracies are an
order-of-magnitude better precision than the
Earth-to-spacecraft position, so that the two
problems can be effectively de-coupled. - Also, note that we assume a perfectly well known
dynamic model that includes the sun, Earth,
Jupiter and the Moon, but that does not include
any photonic effects. This is justified by the
fact that the mission should significantly
improve understanding of those effects, and thus
will be modeled very precisely, and will not
contribute significantly to trajectory
uncertainty. - The nominal trajectory was kept centered on the
ephemeris trajectory from JPL HORIZONS, since
these were calculated with a higher-fidelity
model. The time-steps in the simulation are
tightened from 10 days to only a few hours during
the 2029 close approach, in order to properly
capture the effects of the flyby.
20Expected Tracking and Orbit Prediction Accuracy
21Mitigation Techniques Consideration of Impulsive
OptionsNear-Earth Object Survey and Deflection
Analysis of Alternatives, Report to Congress,
pursuant to Public Law No. 109-155, NASA Office
of Program Analysis and Evaluation (PAE), March
2007.
Impulsive Techniques Description Comments
Conventional Explosive (surface) Detonate on impact
Conventional Explosive (subsurface) Drive explosive device into NEO, detonate
Nuclear Explosive (standoff) Detonate on flyby via proximity fuse
Nuclear Explosive (surface) Impact, detonate via contact fuse
Nuclear Explosive (delayed) Land on surface, detonate at optimal time.
Nuclear Explosive (subsurface) Drive explosive device into NEO, detonate
Kinetic Impact High velocity impact
Can be risky, but must be employed when the
threat of impact is urgent and dire. It is
necessary to have detailed and accurate models
such that one is sure that significant ejecta do
not escape. Otherwise, delta-V errors may do
more harm than good. Moreover, once the impulsive
technique is applied, there is no opportunity to
correct errors. Finally, under the criteria for
our first NEA mitigation effort (next slide),
these techniques are too complex and expensive.
22
22Choice of Initial and Permanent (Slow Push)
Mitigation Techniques Criteria for Humanitys
First Effort
- First efforts at asteroid mitigation should
employ the simplest means possible - There is no more than one maneuvering entity
- There are no power-greedy, active components
- Precise and complex pointing maneuvers are not be
attempted, and - Neither landing nor attaching to the asteroid is
be attempted.
23Mitigation Techniques Slow Push
OptionsNear-Earth Object Survey and Deflection
Analysis of Alternatives, Report to Congress,
pursuant to Public Law No. 109-155, NASA Office
of Program Analysis and Evaluation (PAE), March
2007.
Slow Push Options Description Comments
Focused Solar Use large mirror to focus solar energy on a spot, heat surface, boil off material Violate criteria in the previous chart
Pulsed Laser Rendezvous, position spacecraft near NEO, focus laser on surface, material boiled off provides small force Violate criteria in the previous chart
Mass Driver Rendezvous, land, attach, mine material, eject material from NEO at high velocity Violate criteria in the previous chart
Gravity Tractor Rendezvous with NEO, fly in close proximity for extended period, gravitational attraction provides small force. Good choice for initial mitigation. Requires only data obtained at the outset of the exploration phase
Asteroid tug Rendezvous with NEO, attach to NEO, push. Violate criteria in the previous chart
Enhanced Yarkovsky (Altered solar pressure and Yarkovsky effect) Change albedo of a rotating NEO, radiation from sun-heated material will provide small force as body rotates The only choice for permanent mitigation. Required data collection completed late in mission
24
24
24Exploration/Mitigation Mission Gravity Tractor
Phase
- Following the initial exploration phase, there is
sufficient data to determine how to direct the
gravity tractor force to mitigate the close Earth
approach in 2036. - The gravity tractor method is well-known and can
be applied early in the mission since very
detailed knowledge of the physical properties of
the NEO is not required. - Starting in April 2022, the spacecraft takes up
the gravity tractor station-keeping position and
employs its thrusters to maintain a constant (in
the Euler-Hill frame) position relative to the
asteroid. - The thrusters used are xenon Hall Effect
thrusters canted 38 deg. off the thrust axis and
providing a net tractor force of 4 mN. - By maintaining a position 270m from the
center-of-mass of Apophis for 12 months (prior to
the 2029 close approach), we calculate that 3
Earth radii of deflection will be achieved by
2036. - Because of the tracking and orbit prediction
refinements made possible during the exploration
phase, this amount of deflection should be well
above the remaining orbit uncertainty in 2036.
25Gravity Tractor Effectiveness
- This shows the contours of total orbit position
perturbation in 2036 (in Earth radii) in the
plane of the tractoring start date and the
initial spacecraft mass divided by the square of
the distance from Apophis center of mass.
Parameter Tractoring period Average Mass Distance from CM Force imparted
Selected 1 year 560 kg 270 m .014 N
26Exploration/Mitigation Mission Transition to
Permanent Mitigation Method
- During Gravity Tractoring (using Busek BHT-200
engines), continuous tracking and sustained
albedo and thermal emissions monitoring will
build up sufficient data to fully understand the
solar pressure and Yarkovsky effects on the
Apophis trajectory. - The derived models of solar pressure and,
especially, of the Yarkovsky effect will permit
determination of the albedo change required to
alter the 2036 position to efficiently increase
the Apophis-to-Earth close approach distance. - At this point, the gravity tractor standoff
maneuver is ended so that the spacecraft can
approach the asteroid more closely in order to
deploy the required albedo change treatment
27 Yarkovsky Effect
Cooler dawn side
Net force
Solar Radiation
hotter dusk side
Excess radiation Carries away momentum Pphoton
per photon
D. Vokrouhlicky, A. Milani, and S. R.
Chesley. Yarkovsky Effect on Small Near-Earth
Asteroids Mathematical Formulation and
Examples, Icarus 148, 118-138 (2000).
28
28Apophis Trajectory Change from 2/14/2018 to
2/14/2036 Due to Changes in Energy Reflection,
Absorption and Emission(Giorgini, Benner, Ostro,
Nolan, and Busch. Icarus 193 (2008) 1-19)
Case Spin state Absorption, Mass Trajectory change in earth Radii due to a 4 change in (1-A) ?x/?A (Earth radii)
Prograde Min. Absorp., Max. mass 8.0 ?200
Prograde Max. Absorp., Min mass 18.0 ?450
In-plane Min. Absorp., Max. mass ?9.0 225
In-plane Max. Absorp., Min mass ?20.0 500
Retrograde Min. Absorp., Max. mass ?27.0 675
Retrograde Max. Absorp., Min mass ?58.0 1450
29Efficacy of Albedo Change Treatment
- The previous chart summarizes the results in
Icarus 193 (2008) 1-19) for the Apophis
trajectory change from 2/14/2018 to 2/14/2036 due
to changes in energy reflection, absorption and
emission, for various spin states and where A is
the absorptivity (unity minus the bond albedo). - These calculations include both solar pressure
and Yarkovsky and precisely account for the
effect of the 2029 close encounter. - Based upon these results, we can show that a 5
albedo change (beginning in May, 2023), using
less than 25 kg of surfacing material, will
deflect Apophis between 12 and 40 Earth radii by
2036, depending on the spin state. (for nominal
values of absorption and mass) - Note that whereas gravity tractoring is of
limited duration, the albedo change treatment can
operate forever.
30Surface Albedo Treatment Subsystem (SATS)
- SATS is intended to raise or lower the average
albedo by approximately 5 - Preliminary design concept is a dispenser that
shoots fine powder (having the appropriately high
or low albedo) from a supply canister mounted on
the spacecraft towards the asteroid surface. - SATS consists of two pressurized supplies of
surface treatment, a selection valve, a common
piping system and a common spray nozzle. - Each supply canister contains either a very high
albedo powder or a very low albedo powder (both
in liquid suspension) - Either one or the other will be used, depending
on the albedo/thermal emission data and the
tracking/orbit prediction data collected during
the exploration phase of the mission.
31SATS Preliminary Design
- We select a bag on valve (BOV) system, where
the surface treatments, each a suspension of
albedo change particles (ACPs) and an inert
liquid, are separated from the pressurant. - This permits us to separately customize the
surface treatment liquid matrix for miscibility
and viscosity, and the pressurant for its P-V-T
characteristics. - As in standard dispensing technology, the
pressurant is kept at a temperature and pressure
just above its boiling point. - As the albedo treatment bags are evacuated, some
pressurant evaporates, thereby maintaining
pressure. - When the nozzle actuator is exercised, the
surface treatment is expelled into vacuum, where
the suspension is first atomized, and then the
liquid matrix is quickly dissociated, leaving
only the ACPs. - Both the expulsion through the nozzle and the
dissociation process leave the ACPs ionized.
32Basic Strategy of Albedo Change
- Maneuver the spacecraft to within 100m of the
surface on the equatorial plane, and maintain
that station for the duration of the surface
treatment letting the asteroid rotate beneath - Once on-station, the select valve is exercised
and the spray nozzle engaged. - The spray nozzle then directs the stream of
albedo change particles (ACPs) to the surface of
the asteroid until the supply is exhausted. - The ACP deposition process is continuously
monitored by the Navigation Camera aboard the
spacecraft to ensure adequate distribution in the
correct area. - The thrust due to ACP dispensing can sustain the
hover maneuver.
VE
H
rC
RA
33
33Bounds on Albedo Change Particle (ACP) Size
- The area covered by ACPs per unit mass is
inversely proportional to their size. Thus
smaller particles are better. - From the material presented on electrostatic
interactions and particle levitation in P. Lee,
Dust Levitation on Asteroids, Icarus
124,181-194 (1996). , in order to prevent ACPs
from levitating and escaping, the ACP diameter,
DP, must satisfy DPgt100 ?m - Thus, from what has been noted for coverage
efficiency, the APC size should not be much
greater than this.
34Bounds on ACP Dispensing Speed
The relatively slow dispensing is advantageous
for the Yarkovsky effect. Assuming that the
spacecraft will hover over Apophis in its
equatorial plane as it rotates below, the SATS
will paint a wide ribbon extending over the
circumference. This will provide a continuous
alteration in the Yarkovsky effect.
35Charged APCs Role of electrostatics (P.Lee,
Dust Levitation on Asteroids, Icarus 124,
181-194 (1996))
- On the sun-facing side, the surface has a net
positive charge, - But above the surface is a cloud of ejected
electrons. - The SATS nozzle is designed to impart a negative
charge to the ACPs and dispensing is performed
solely on the sun lit side. - This means that the ACP descent is slightly
retarded when moving through the relatively
diffuse, negatively charged plasma layer over the
surface, but the particles are much more strongly
attached to the surface. - This effect further ensures that the particles
will be quickly bound to the surface and will not
rebound or levitate into an escape condition.
VE
RA
36Cost Modeling
- Since the proposal submittal, an updated costing
exercise was performed in April 09, at Ames
Mission Design Center. - Spacecraft and Payload cost estimated using
parametric cost estimation models - Spacecraft cost modeling SSCM, and Analogy
NEAR, Dawn, DS-1 - Payload cost estimation MICM, NICM system, SOSCM
- Albedo payload cost modeled after propulsion
system - Other costs modeled as percentage wrap-factors
- Phase E/F costs based on historic, interplanetary
Discovery class missions
37Cost Estimates
Payload Price (US Dollars)
Dawn Framing Camera 25420
NEAR Laser Range Finder 10135
Microbolometer 15182
Albedo Change 10210
Total 60947
Element Amount in 2009 (US Dollars, Thousands)
PM/SE/MA 24500
Payload 61000
Spacecraft 102500
Hardware total 163500
MOS.GDS 18000
Development Cost, No Reserves 206000
Development reserves 61800
Total Development costs 267800
Phase EF costs 28000
Phase EF Reserves 8400
Total mission costs, no launch Vehicle 304200
Launch Vehicle/services 50000
Total mission 354200
- Modifications additions to the proposal cost
estimates - More advanced Op Nav camera
- Microbolometer added
- Development cost and reserves
added - Launch vehicle/services added
- Phase E F costs added
38Reference Mission Advantages
- Precise and complex pointing maneuvers are not
attempted. - Neither landing nor attaching to the asteroid are
attempted. - Only the spacecraft itself is the active agent.
- Regarding gravity tractor positioning,
simulations have shown simple station-keeping
controls to be quite effective Near-Earth
Object (NEO) Analysis of Transponder Tracking and
Gravity Tractor Performance, Reort to the B612
Foundation. Final Report, JPL Task Plan No.
82-120022, September 22, 2008. - For the final, permanent mitigation technique
- Only equipment on board the spacecraft is used.
There are no complex maneuvers or landing/docking
attempts. - The spacecraft station-keeps, dispensing the
albedo change treatment while the asteroid
rotates under it. - Moreover, the reaction force produced by the
dispenser can do double duty as the propulsion
device used for station-keeping. - The albedo change dispenser itself is entirely
passive, with few moving parts, requires
negligible power, and relies on very familiar,
highly developed technology with heritage that
includes both commercial aerosol dispensers and
tankage and nozzles associated with space-borne,
cold-gas propulsion. - In summary, our proposal is to develop a model
procedure for exploring and mitigating hazardous
near-Earth asteroids that reliably guarantees
mitigation success, and eventual threat
elimination, while using the simplest possible
technological means.