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A Mission Template for Exploration and Mitigation of Hazardous Near-Earth Asteroids

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Title: A Mission Template for Exploration and Mitigation of Hazardous Near-Earth Asteroids


1
A 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
3
History 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

4
History 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

5
Apophis 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
6
Apophis 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
7
Best 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

8
Why 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

9
Systematic 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!

10
Combined 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

11
Dry 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
12
Exploration/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
14
Exploration/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.

15
Science 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
16
Exploration 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

17
Tracking 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
18
Radio 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
19
Radio 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.

20
Expected Tracking and Orbit Prediction Accuracy
21
Mitigation 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
22
Choice 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.

23
Mitigation 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
24
Exploration/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.

25
Gravity 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
26
Exploration/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
28
Apophis 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
29
Efficacy 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.

30
Surface 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.

31
SATS 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.

32
Basic 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
33
Bounds 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.

34
Bounds 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.
35
Charged 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
36
Cost 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

37
Cost 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

38
Reference 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.
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