American Institute for Aeronautics and Astronautics
ATHOS DEFLECTION MISSION ANALYSIS AND DESIGN
Ralph Kahle*, Gerhard Hahn†, Ekkehard Kührt‡
German Aerospace Center, Institute of Planetary Research, Rutherfordstr. 2, 12489 Berlin, Germany
and
Stefanos Fasoulas§ Technische Universität Dresden, Institute for Aerospace Engineering, 01062 Dresden, Germany
ABSTRACT
An asteroid deflection mission is analyzed and designed based on a fictitious threat scenario. The mission objective is to prevent the collision of the virtual binary asteroid Athos with Earth on February 29, 2016. Two alternative techniques are investigated both aiming on the diversion of Athos’ trajectory rather than on its destruction. These techniques are an inflatable solar collector and a kinetic energy projectile. Based on the results of a precursor mission to Athos and the outcomes of technology feasibility studies, one technology will be selected for mitigation. Mission constraints are identified as the time of collision, the physical and orbital properties of Athos, and the technology readiness level of the envisaged mitigation techniques and underlying launch system and spacecraft technology. The last point is of high priority because of the short time span from time of detection (February 22, 2005) to launch date of the deflection spacecraft. Detailed mission schedules, ∆V analysis, mass budgets, payload analysis, and cost estimates are derived to assess the feasibility of both mitigation techniques. We show that Athos can be deflected with non-nuclear concepts. Further, the threat posed by Athos’ satellite DeWinter, which might separate during mitigation, is assessed in terms of Earth atmosphere entry analysis.
INTRODUCTION
For the 2004 Planetary Defense Conference, four fictitiously defined threat scenarios are posed,10 where
one comet and three asteroids will strike Earth. The present work deals with one out of those four near-Earth objects (NEOs), which we believe to be the most likely scenario. The candidate is a 200 m binary asteroid named Athos, which without mitigation will collide with Earth on February 29, 2016.
In this paper we first briefly introduce numerical tools that underlie mission analysis, e.g. orbit computation and trajectory design. Next, the threat is characterized and a mission objective is defined, which demands for the deflection of Athos from its collision path with Earth. Based on the given physical properties of Athos, mission requirements (∆V and impulse required for diversion) and constraints (timeline) are defined. Thereafter, we characterize the deflection mission, where two applicable mission concepts are identified. These comprise a solar collector and kinetic energy projectiles.
We show that Athos can be deflected with non-nuclear concepts, which are investigated and evaluated in terms of detailed mission schedules, spacecraft ∆V analysis, mass budgets, payload analysis, and cost estimates. Further, atmospheric entry analysis is carried out for DeWinter, which might separate from Athos during mitigation. Here, we estimate the energy released by an airburst in the upper Earth atmosphere, which can be compared to the 1908 Tunguska explosion.
MISSION ANALYSIS TOOLS
The analysis of mission requirements for an asteroid deflection demands especially a careful analysis of orbital parameters of the collision object. Naturally, one would start to compute the objects’ orbit based on the observation data. In case of an impact risk, observers would immediately start to refine orbit information based on further observation campaigns. Finally, the orbit would be well characterized and the threat could be evaluated. These statements refer to a short
* PhD student, Section Physics of Small Bodies, corresponding author, contact: [email protected] † Scientific staff member, Section Physics of Small Bodies. ‡ Section leader, Section Physics of Small Bodies. § Professor, Space Systems and Utilization, Member AIAA.
2004 Planetary Defense Conference: Protecting Earth from Asteroids<br>23 - 26 February 2004, Orange County, California
American Institute for Aeronautics and Astronautics
integration scenario of a few decades. Otherwise uncertainties in numerical treatment arise from the modeling of poorly characterized effects such as the thermal re-radiation of solar energy absorbed by the asteroid.
Here, we deal with a fictitious short-term scenario and no further observation results are available. Anyway, to enable a precise numerical treatment, we utilize the information about the state vector of Athos at time of impact, which is given in Earth-fixed rotating coordinates (tab. 1).10 After a coordinate transformation towards the heliocentric ecliptic reference frame the heliocentric orbit computation is carried out with the 15th order Gauss-Radau numerical integrator.5 Here, Earth and Moon are treated separately and inner and outer planets are included. Further, a subroutine is implemented, which enables us to examine atmospheric entry effects, e.g. braking, flight path, ablation, object deformation and fragmentation (air burst). The analysis follows the model proposed by Chyba et al.3
Mission opportunities to Athos are analyzed in the Lambert approach, e.g. launch and arrival dates are varied within a certain time frame to find optimal trajectories assuming impulsive maneuvers (chemical propulsion rather than electrical propulsion is
considered). Besides simple elliptical transfers, multiple-encounter transfers as well as gravity-assist trajectories are analyzed.
THREAT CHARACTERIZATION
Preliminary properties of binary asteroid Athos are extracted from the impact scenario document10 and are summarized in table 1. The potential damage due to an ocean impact of the 200 m rubble pile asteroid is not accessed within this work. Possible consequences could range from a high energy explosion in Earth’s atmosphere (airburst) to an ocean impact that could result in a hazardous tsunami. Because fatal damage cannot be ruled out, we have to investigate how to prevent such an impact. Because of the large uncertainties regarding the dynamical and physical properties of the binary asteroid a precursor mission is proposed in order to reveal those properties that are significant for the particular mitigation techniques.
Precursor mission
The precursor spacecraft has to support a multi-spectral stereo imaging camera in order to determine shape, spin state, surface mineralogy, and eventually topography. Further, radio science investigations are required in
Athos Athos satellite (DeWinter) Time/date of detection 22 Feb 2005
00:00:00 UT 22 Feb 2009 00:00:00 UT
Expected date of impact 29 Feb 2016 06:07:42.332 UT
Same as Athos
Impact location Pacific Ocean, close to California coast 42.1°N, 233.8°E
Within 200 km of Athos’ impact location
State vector of impactor in Earth-fixed rotating coordinates at time of impact
x = -2927.238530 km y = -3999.564261 km z = 4000.938620 km x& = 3.777952 km/s y& = 5.984562 km/s z& = -9.690825 km/s
Same as Athos
Spin period, inertial direction (α, δ), and geographic pole location (λ, φ) at time of detection
Period = 1.0 days ± 1 hr (α, δ) = unknown (λ, φ) = unknown
Physical model S-type rubble pile size: 200 x 180 x 170 m mass: 1.1x1013 g ± 50% density: 3.5 ± 1.5 g/cm3
S-type rubble pile size: ~70 m mass: 6.3x1011 g ± 50% density: 3.5 ± 1.5 g/cm3
Tab. 1: Mission scenario for Athos and its satellite DeWinter, extract.10
3
American Institute for Aeronautics and Astronautics
order to determine mass and gravity harmonics. A possible precursor concept could be the 120 kg Simone spacecraft.1 If development and production of such a novel spacecraft could not be accomplished within the restricted time frame, a precursor spacecraft should be built based on known and successfully flown spacecraft technology instead, for example the NEAR spacecraft.
+ + + ++
+
+++
++++++++++++++++++++++++
-2-1.5
-1-0.5
00.5
11.5
X [AU]
1.5
-1
-0.5
0
0.5
1
1.5
2
Y [AU]
-0.2-0.15-0.1-0.0500.050.10.15Mars
Earth
Athos S/C
Launch from GTO 11 Jan 2008
Earth Gravity-Assist 1 Mar 2008
Athos Rendezvous 17 Oct 2008
Deep Space Maneuver
Fig. 1: Precursor spacecraft (S/C) rendezvous trajectory to Athos.
An optimized rendezvous mission opportunity for the precursor is depicted in figure 1. The launch from geostationary transfer orbit (GTO) is foreseen on January 11, 2008. The required ∆V for the transfer from GTO to the Earth-escaping hyperbola is estimated as
787 m/s. After accomplishment of a deep space maneuver, the precursor swings by Earth on March 1, 2008. The spacecraft encounters Athos on October 17, 2008. The rendezvous velocity equals 1896 m/s yielding a total-∆V of 2690 m/s, which is well within ∆V-budgets of conventional exploration missions.
It is assumed that the in-situ characterization of Athos can be completed after a three months survey. The mission timeline is depicted in figure 9. The results of the precursor mission, especially improved data on mass and orbital elements, are decisive for determining deflection mission requirements, e.g. total impulse for diversion.
DEFLECTION MISSION REQUIREMENTS
The change in Athos’ orbital velocity ∆V required to deflect it by one Earth radius (this corresponds to a grazing encounter) and additionally to deflect it to a safe encounter at a surface distance of 10,000 km are estimated as a function of the epoch of interception. This analysis follows the numerical procedure recently proposed by Carusi et al.2 The results are depicted in figure 2. Clearly, it would be advantageous to apply the deflecting impulse within Athos’ direction of flight rather than in opposite direction. To illustrate the problem a reference plane is defined perpendicular to Athos’ velocity vector at time of collision with Earth. When projecting Earth’s center into that plane, the distance of impact location from Earth center can be
1 E-02
1 E-01
1 E+00
1 E+01
1 E+02
1 E+03
01 Jan2003
01 Jan2004
01 Jan2005
01 Jan2006
01 Jan2007
01 Jan2008
01 Jan2009
01 Jan2010
01 Jan2011
01 Jan2012
01 Jan2013
01 Jan2014
01 Jan2015
01 Jan2016
01 Jan2017
Epoch of deflection
dV [m
/s]
negative dV for safe encounterpositive dV for safe encounternegative dV for grazing encounterpositive dV for grazing encounter
Fig. 2: ∆V for Athos deflection as a function of interception epoch.
4
American Institute for Aeronautics and Astronautics
monitored. Without mitigation, Athos will impact outside of Earth’s center when viewing from Sun. Thus, a sufficient deflection can be achieved with less effort by acceleration (this corresponds to a raising of Athos’ semi-major axis) rather than deceleration (lowering). From figure 2, the minimum accelerating ∆V’s (positive) are found to be 0.035 m/s and 0.067 m/s for an interception in April 2012 for deflection into grazing and safe encounter, respectively. Further local minima exist at epochs of perihel passage. Interception opportunities have to be identified for the time span from year 2010 to 2014. Otherwise ∆V requirements increase substantially, which might prevent sufficient deflection. A significant increase of several orders of magnitude appears during the final orbit before impact on February 29, 2016.
The total impulse required for deflection equals the product of Athos’ mass and required ∆V at a particular interception epoch. For example, a minimum impulse of 3.87·108 Ns ± 50% or 7.37·108 Ns ± 50% would be necessary for deflection into a grazing or safe encounter, respectively for an interception in April 2012. The large uncertainty in impulse stems from the poor knowledge of Athos’ mass. Although it seems more likely that a rubble pile object has a lower density than 3.5 g/cm3 (table 1), it cannot be ruled out that the object is intact and denser instead. Thus, the full range of possible impulse requirement has to be considered. Only after accomplishment of the precursor mission the exact mass of Athos will be known and more precise impulses can be estimated.
DEFLECTION TECHNOLOGIES
In parallel to the precursor operation, the concept exploration and the detailed development of deflection technologies should be performed. Ideally, feasibility studies for individual mitigation techniques would be carried out in advance of a real threat. This preparatory work could save precious time in case of emergency.
Extensive literature review and critical system analysis have shown that only three techniques exist, which might be feasible for purposes of NEO mitigation: nuclear explosive, kinetic energy projectile and solar collector.6 In the following we are concentrating on the non-nuclear techniques. Based on the results of feasibility studies a mitigation technique can be selected for development at the end of year 2005 (figure 9).
Solar collector
The basic idea of this technology is to concentrate solar radiation onto Athos’ surface, where depending on duration and intensity of illumination, the material within the spot will be heated up and vaporized delivering a low but continuous thrust.11
The inflatable space structure that underlies the solar collector technique has been developed by L’Garde Inc. under NASA’s in-space technology experiments program. The technology readiness level (TRL) is evaluated as high, because a 14 m diameter inflatable antenna experiment (IAE) was successfully deployed in an operational orbit in May 1996 (Space Shuttle flight STS-77, Spartan 207). It is assumed that the further-development towards a larger collector can be accomplished within a development time of three years.
The mitigation spacecraft has to support the inflatable solar collector. During the interplanetary travel towards the target asteroid, the collector is stowed in a comparably small box less than 1 m3 in volume.14 This enables a mission performance with only one launch of a moderate spacecraft. When rendezvousing with the target object, the collector is deployed and inflated to a 170 m diameter structure. This size is a compromise on payload capability and collector performance, which
SUNS/C
Athos‘trajectory
DeWinter
Incident Sunbeam
Spot
Reflector
Clear canopy
Fig. 3: Solar collector working principle in vicinity of Athos and its satellite (true to scale).
5
American Institute for Aeronautics and Astronautics
will be discussed later. The collector consists of an inflatable torus that houses the inflatable parabolic reflector (figure 3). In order to achieve high concentration, ratios an off-axis collector geometry is applied.7 Inflatable struts connect the collector with the spacecraft. The envisaged 170 m diameter collector is expected to have a mass of 1000 kg.14
Figure 3 depicts a possible arrangement of the deployed collector in vicinity of the binary asteroid. The spacecrafts orbit is not stable and requires therefore a continuous balancing of disturbing forces, which are mainly due to solar pressure (~ 104 mN), attractions of Athos (~ 15 mN) and DeWinter (max. ~ 0.5 mN), and pressure of vapor and dust. Additionally, the close residence of Athos’ surface within the foci of collector ellipse has to be ensured. Therefore, Athos’ motion around the center of gravity of the binary asteroid and topographical deviations from Athos’ ellipsoidal shape have to be taken into account, too. Here, a thrust of ~ 5 mN is estimated, yielding a total thrust (worst case) of 130 mN, which has to be contributed by the spacecraft’s orbit maintenance thrusters. For example, 30 days of collector operation on a 2 ton spacecraft equipped with N2H4-resistojets (specific impulse of ~
300 s) would require 112 kg of fuel. We will show later that no more than 30 days are necessary for mitigation. Concerning the positioning of the spacecraft, the collector operation seems to be feasible, albeit huge efforts are involved with the orbit and attitude control system.
To estimate the amount of force that can be generated with the 170 m collector, model assumptions are necessary for the process of evaporation. Because of Athos’ affiliation to S-type, the surface could consist of Pallasite and/or Mesosiderite.12 Both minerals consist of a Nickel-Iron-matrix containing Olivine in the case of Pallasite and silicates (Pyroxene, Plagioclase, Olivine) in the case of Mesosiderite. Further, we assume that for increasing surface temperature (due to collector illumination) the first mineral to evaporate is Nickel-Iron. Based on the Clausius-Clapeyron equation, empirical equations are derived for the vapor pressures of Fe and Ni as a function of surface temperature. A condensation temperature of 1200°C at 10 Pa pressure is applied for calibration.12 Further assumptions are made for latent heat (9.1 MJ/kg), thermal conductivity (2 W/m/K), specific heat (~ 1000 J/kg/K),6 and
Fig. 4: Evolution of surface temperature, heat fluxes, evaporation rate and directed vapor pressure for energy input of 1 MW/m2. The spot size is about 6.5 x 5 m (25.5 m2).
Fig. 5: Evolution of surface temperature, heat fluxes, evaporation rate and directed vapor pressure for energy input of 2 MW/m2. The spot size is about 3.5 x 2.9 m (8 m2).
6
American Institute for Aeronautics and Astronautics
emissivity (0.95).
A thermal model that takes into account the energy input due to the collector (which mainly depends on solar distance, spot size, and concentration ratio), one-dimensional thermal conduction into the core, heat radiation, and evaporation of surface material is applied to analyze the generation of vapor. A total efficiency of 0.5 is assumed, which is the product of efficiencies for light-passing through canopy twice (0.92 = 0.81), film reflectivity (0.85), assumed asteroid surface reflectivity (0.9), and assumed topographical shadowing (0.8).
Figures 4 and 5 depict the results obtained from thermal analysis. Figure 4 corresponds to an off-focus surface distance of ± 1.5 m, where an ellipsoidal spot of 6.5 x 5 m is illuminated on Athos’ surface with an average energy input of 1 MW/m2. At beginning of collector operation, Athos’ surface temperature is assumed as 200 K. After about 10 s of illumination the production of vapor starts and a recoil impulse is transferred onto Athos. For in-foci illumination, 2 MW/m2 are beamed into a 3.5 x 2.9 m spot, where evaporation begins after 2.5 s (figure 5). Off-foci surface positioning results from Athos’ topography as well as spacecraft orbit control. For deviations larger than ± 3 m, the increasing spot size limits the energy input to < 0.6 MW/m2 (figure 6), which disables a sufficient vapor production.
0,0
0,5
1,0
1,5
2,0
2,5
-4 -3 -2 -1 0 1 2 3 4
Off-foci distance [m]
Ene
rgy
inpu
t [M
W/m
²]
Fig. 6: Concentrated energy input as a function of surface off-foci distance. By courtesy of Hans Krüger.
The resulting pressure normal to Athos’ surface (bottom of figures 4 and 5) is estimated as the product of evaporation rate, vapor velocity, and a factor that accounts for non-directed spreading (0.5).11 Finally, the thrust is estimated as the product of directed pressure and spot area.
But, before estimating thrust, the rotation of Athos has to be considered, too. This is very important, because the surface moves beneath the spot, where the velocity
depends on geographic spot location. In a worst case, the spot would be directed onto Athos’ equator, which corresponds to a surface speed of 0.053 m/s. Within the time required for establishing a sufficient evaporation rate (5 to 20 s) the surface would move 0.25 to 1 m for 2 to 1 MW/m2 of energy input, respectively. This actually lowers the generated thrust. After applying an additional safety factor, we estimate an average continuous thrust of 400 N. To achieve the required total impulse of 3.87·108 Ns ± 50% (grazing encounter) or 7.37·108 Ns ± 50% (safe encounter), the collector has to be operated for the duration of 11.2 ± 5.6 days or 21.3 ± 10.7 days, respectively.
Because of the close distance between collector and Athos’ surface, vapor will start to condensate on the canopy and dust particles might penetrate through the canopy and eventually the reflector membrane. Both effects cause a reduction of canopies efficiency, which directly affects the amount of concentrated energy and thus the generated thrust. At the moment, we do not know the evolution of canopy efficiency. For that, laboratory or space-environment investigations are necessary and should be accomplished within the collector development phase. Besides, more sheltered collector-designs should be investigated, e.g. a Cassegrain-type arrangement, in which the concentrated beam is reflected by a secondary mirror through a hole in the center of the collector.11
-2
-1,5
-1
-0,5
0
0,5
1
1,5
2
-2 -1,5 -1 -0,5 0 0,5 1 1,5 2
X [AU]
Y [A
U]
Mars
Earth
Athos
S/CLaunch from GTO12 Mar 2009
Athos Rendezvous16 Jan 2012
Fig. 7: Solar collector rendezvous trajectory, projection into ecliptic plane.
7
American Institute for Aeronautics and Astronautics
Development and production phase of collector and carrying spacecraft are assumed to finish at the end of year 2008 (figure 9). Because of the temporal location of the minimum ∆V-requirement, the spacecraft should rendezvous Athos no later than early 2012. Otherwise mitigation impulse requirements increase and complicate mission success. No optimal Earth- or Mars-Gravity-Assist trajectories are identified that coincide with this transfer window. But, a direct rendezvous mission opportunity is found, which foresees the launch of the collector spacecraft from GTO on March 12, 2009 (∆V = 1561 m/s). The rendezvous with Athos is scheduled for January 16, 2012 (∆V = 1278 m/s), as depicted in figure 7.
Assuming the launch with a Delta IV Heavy or Ariane 5 and further the use of chemical propulsion (specific impulse of 350 s), a maximum spacecraft mass of 4 tons is available at time of arrival at Athos, which gives us a safety factor of 2 concerning the aspired collector carrying spacecraft of 2 tons.
In case the identified operation problems cannot be solved, the kinetic energy projectile has to be chosen for mitigation after accomplishment of the feasibility study instead. The same applies in case the precursor mission reveals the unfavorable nature of Athos’ surface, e.g. a dense iron-nickel composition (high thermal conductivity) or a ragged topography that prevents an operation close to the foci of ellipse.
Kinetic energy (KE) projectile
This technique is based on the momentum transfer from an impacting spacecraft on the hazardous object. The impacting process goes along with the ejection of crater material, where the total momentum change of the target object is the momentum of the escaping ejecta plus the momentum carried with the projectile. For non-porous targets the ratio between ejecta momentum and projectile momentum can be as large as 13, whereas for porous targets this could be decreased to 0.2, yielding a momentum enhancement factor of 14 and 1.2, respectively.8
In the following a momentum enhancement factor of 3 is assumed, which was deduced from the physical properties of Athos and experimental factors for analogous materials that were determined by Tedeschi et al. For example, they reported factors of >2.94, 3.87, and 4.76 for the impact of a 7.6 to 8.4 km/s aluminum projectile into a silicate rock, a rock with 10% porosity,
and an aerogel (very-high-porosity silica structure), respectively.13
To achieve similar diversion results as with the solar collector interaction an impulse of 3.87·108 Ns (grazing encounter) or 7.37·108 Ns (safe encounter) has to be applied. Here, we do not explicitly discuss the +50% margin since for denser and less porous objects a larger momentum enhancement factor applies, because less impact energy is absorbed in crushing of pores.
Because of the required enormous total impulse, more than one launch is necessary. The state-of-the-art launch system arsenal comprises two rockets, which could both be utilized for the task of KE interaction. These are the American Delta IV Heavy and the European Ariane 5 capable of launching 12.4 and 12 tons (planned for 2006) into GTO, respectively.
A KE mission opportunity is identified, which maximizes the product of spacecraft mass and relative velocity (in Athos’ direction of flight) when impacting on Athos in April 2012. The launch from GTO is scheduled for December 27, 2011. A velocity change of 787 m/s, applied during perigee passage, is required to get onto an interplanetary orbit. After a low demanding deep space maneuver (4 m/s) the spacecraft swings by Earth on February 15, 2012. This gravity-assist enables the high energy impact with Athos on April 5, 2012. The transfer trajectory is depicted in figure 8.
+
+
++
+
++ + + + ++
-2-1.5
-1-0.5
00.5
11.5
X [AU]
-1.5
-1
-0.5
0
0.5
1
1.5
2
Y [AU]
-0.2-0.15-0.1-0.0500.050.10.15
Mars
Earth
Athos
S/C Launch from GTO
27 Dec 2011 Earth Gravity-Assist
12 Feb 2012
Athos Impact 5 Apr 2012
Deep Space Maneuver
Fig. 8: Kinetic energy impact trajectory.
Based on the estimated velocity budget and the known launch capability, the total mass of spacecraft at time of arrival at Athos is determined as 9574 kg (the specific impulse of spacecraft propulsion is assumed as 350 s). The relative velocity within Athos’ direction of flight
8
American Institute for Aeronautics and Astronautics
amounts to 8670 m/s. After applying the momentum enhancement factor of 3, the impact of a single spacecraft evokes a total impulse of 2.49·108 Ns. Thus, for deflection into a safe encounter, three spacecrafts are necessary, which equals a total impulse of 7.47·108 Ns.
Nevertheless, the launch of four identical spacecrafts is scheduled instead. After impact of all four spacecrafts the post-interaction tracking will reveal the correctness of the enhancement factor assumption. In a worst case scenario, it would turn out that the enhancement factor is minimal (1.2). In this case, the applied impulse (3.98·108 Ns) would still be sufficient for the deflection into a grazing encounter. On the other hand, for larger enhancement factors the encounter distance would be even larger than required for safe encounter.
Since launch systems cannot be started within short intervals, the spacecrafts are to be launched into geostationary transfer parking orbits, beginning one year in advance of interplanetary transfer to Athos (figure 9).
Last but not least it has to be ensured that the projectile shelling will not cause fragmentation of Athos into large boulders that possibly remain on collision course with Earth. The shattering threshold for silicate objects is difficult to assess. Assuming a specific energy of 100 J/kg,9 an energy of 1.1·1012 J ± 50% would be required for disruption of Athos. The kinetic energy implied in each of the proposed projectiles equals 3.6·1011 J. Hence, the staggered impact of the
spacecrafts should pose no serious danger of fragmentation.
The TRL of KE projectiles is evaluated as high. No technological problems are expected. Nevertheless, possible development efforts could deal with shape and composition of a KE spacecraft.
POST-INTERACTION ASSESSMENT
After the interaction with solar collector or projectiles is finished in April 2012, the outcomes have to be monitored. Here, a ground-based post-interaction tracking of Athos and possibly of larger fragments from the projectile shelling should be performed. Further, Athos’ satellite DeWinter demands for careful tracking in case of being released from Athos. If not feasible with ground based telescopes, existing space-born telescopes should be utilized instead.
Anyway, if everything goes well with the interaction during April 2012, Athos will be deflected from its collision path with Earth. The post-interaction orbit analysis shows that Athos will flyby Mars on its new path in July 2012 at a distance of 0.072 AU. Thereafter, it will not approach closer to Earth than 0.02 AU for the next century.
The fate of DeWinter
So far, we have not discussed the consequences of mitigation efforts on Athos’ 70 m satellite DeWinter. Without examination, we think that for the case of solar collector mitigation DeWinter will remain on its orbit
2005 2006 2007 2008 2009 2010 2011 2012 2013 2014 2015 2016 2017 2018Ground-based telescope survey detection follow-up observation post-interaction trackingPrecursor mission development + production transfer to Athos in-situ characterizationDeflection technologies feasibility study1) Solar collector development + production launch + transfer to Athos concentrator operation or failure2) Kinetic energy projectile production of 4 KE spacecrafts launch into parking orbit transfer + staggered Athos impactEvacuation of coastal area
colli
sion
with
Ear
th o
n Fe
brua
ry 2
9, 2
016
in c
ase
of h
azar
dous
DeW
inte
r air
burs
t
Fig. 9: Athos deflection mission timeline. Note that mitigation is accomplished either with solar collector or KE projectiles. If Athos’ satellite remains on collision course with Earth, local evacuation might become necessary.
9
American Institute for Aeronautics and Astronautics
around Athos, and thus would be deflected too. However, for the projectile interaction we expect separation of the weakly bound objects. For the case that DeWinter remains on the collision path with Earth we analyze atmospheric entry. Model assumptions are made for objects strength (107 Pa) and heat of ablation (8·106 J/kg).
Figure 10 depicts outcomes of the entry analysis. Note that DeWinter’s initial radius is 30 m instead of 35 m as given in table 1. This is because the entry body is treated as a cubic cylinder rather than a sphere.3 The velocity of DeWinter relative to Earth increases substantially due to gravitational attraction. When penetrating denser atmosphere (beginning at an altitude of about 30 km), atmospheric braking prevails and the object decelerates. No significant ablation is determined. At an altitude of 15 km the dynamic pressure exceeds objects strength and DeWinter starts to fragment. Atmospheric entry is analyzed up to the point where the radius of fragment cloud exceeds the double initial radius. Here, an airburst is estimated at an altitude of 9 km above the Pacific Ocean. The released energy (kinetic energy at time of air burst) is estimated as 4.3·1016 J, which corresponds to 10 Mt TNT (high explosive equivalent). About the same amount of energy was probably released by the 1908 Tunguska explosion, where an area of 2200 km2 was devastated.3 The potential damage due to an equivalent energy release close to the Pacific coast is not accessed within this work. Eventually, coastal residents have to be evacuated.
9500
10000
10500
11000
11500
12000
12500
1 10 100 1000 10000
Altitude [km]
Rel
ativ
e ve
loci
ty [m
/s]
20
30
40
50
60
70
80
Rad
ius
[m]
Radius
Velocity
Fig. 10: Relative velocity and radius of DeWinter during Earth atmospheric entry.
MISSION COST EVALUATION
In the following we give only a rough estimate of mitigation mission cost in order to determine its
magnitude. Athos deflection mission expenses arise to a moderate extend from Earth-based telescope operation for follow-up observations and post-interaction tracking of Athos and DeWinter as well as from feasibility studies of mitigation techniques. Development and production costs of the precursor spacecraft are estimated to $50 million. Additional cost arises from the precursor launch, which might be accomplished as a piggyback payload for the case of utilizing the Simone spacecraft. Development and production costs for the solar collector are estimated to $30 million,4 whereas no significant development cost is expected for the KE technology. Production cost of the mitigation spacecrafts (one for solar collector or four for KE interaction) should be significantly cheaper than the total cost of the Rosetta mission, which is about $1 billion. One launch with the Ariane 5 into the GTO is typically $150-200 million. Thus, the total costs are expected to be less than $2 billion, very much lower if compared to the damage that could result from an impact. Cost for evacuation is not assessed.
CONCLUSIONS
The proposed mission design for the deflection of Athos offers a high probability of success at moderate expenses. A precursor mission is foreseen to reveal the most critical properties for mitigation interaction. Based on its results the solar collector or KE is chosen as mitigation technique. This selection also depends upon the results of preceding feasibility studies. These studies are highly recommended especially for the solar collector technology. Here, significant problems are identified, which have to be solved. Anyway, the kinetic projectile interaction seems to be feasible, although it demands for larger efforts (four launches instead of one). For both techniques mission opportunities are identified for an interception during period of minimum mitigation impulse requirement.
The projectile shelling will probably cause the separation of satellite DeWinter, which could remain on its collision path with Earth. In that case, an airburst is estimated, releasing an energy of about 10 Mt TNT. This corresponds to the 1908 Tunguska explosion and would probably demand for evacuation of coastal areas.
ACKNOWLEDGEMENTS
R. Kahle would like to thank the German Aerospace Center (DLR) for a research fellowship to study NEO deflection problems.
10
American Institute for Aeronautics and Astronautics
REFERENCES 1 Ball, A., et al., 2002, SIMONE: near-Earth asteroid rendezvous microsatellites with solar-electric propulsion. In: Proceedings of Asteroids, Comets, Meteors 2002, 87-90. 2 Carusi, A., Valsecchi, G.B., D’Abramo, G., Boattini, A., 2002, Deflecting NEOs in Route of Collision with the Earth. Icarus, 159, 417-422. 3 Chyba, C.F., Thomas, P.J., Zahnle, K.J., 1993, The 1908 Tunguska explosion: atmospheric disruption of a stony asteroid. Nature, 361, 40-44. 4 Dornheim, M.A., 1999, Inflatable structures taking to flight. Aviation Week & Space Technology, 150, 60–62. 5 Everhart, E., 1985, An efficient integrator that uses Gauss-Radau spacings. In Carusi, A. & Valsecchi, G.B. (eds.), Dynamics of comets: their origin and evolution, 185-202. 6 Gritzner, C., Kahle, R., 2004 (in press), Mitigation technologies and their requirements. In Belton, M.J.S. et al. (eds.), Mitigation of Hazardous Impacts due to Asteroids and Comets, Cambridge University Press. 7 Grossman, G., Williams, G., 1990, Inflatable concentrators for solar propulsion and dynamic space power. Journal of Solar Engineering, 112, 229-236. 8 Holsapple, K.A., 2002, The deflection of menacing rubble pile asteroids. In: Extended abstracts from the NASA workshop on the scientific requirements for mitigation of hazardous comets and asteroids, 48-51. 9 Holsapple, K.A., Giblin, I., Housen, K., Nakamura, A., Ryan, E., 2002, Asteroid impacts: laboratory experiments and scaling laws. In Bottke, W., et al. (eds.), Asteroids III. University of Arizona Press, 443-462. 10 Lynch, D.K., Peterson, G.E., 2003, Athos, Porthos Aramis & D’Artagnon: Four planning scenarios for planetary protection. 11 Melosh, H.J., Nemchinov, I.V., Zetzer, Yu.I., 1994, Non-nuclear strategies for deflecting comets and asteroids, in Gehrels, T. (ed.), Hazards due to Comets and Asteroids. University of Arizona Press, 1111-1132. 12 Schultz, L., 1993, Planetologie – Eine Einführung, Birkhäuser Verlag.
13 Tedeschi, W.J., Remo, J.L., Schulze, J.F., Young, R.P., 1995, Experimental hypervelocity impact effects on simulated planetesimal materials. International Journal of Impact Engineering, 17, 837-848. 14 Thomas, M., 1992, Inflatable Space Structures: Redefining aerospace design concepts keeps costs from ballooning. IEEE Potentials Magazine, 29-32. http://www.lgarde.com/people/papers/structures.html
![Page 1: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/1.jpg)
![Page 2: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/2.jpg)
![Page 3: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/3.jpg)
![Page 4: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/4.jpg)
![Page 5: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/5.jpg)
![Page 6: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/6.jpg)
![Page 7: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/7.jpg)
![Page 8: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/8.jpg)
![Page 9: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/9.jpg)
![Page 10: [American Institute of Aeronautics and Astronautics 2004 Planetary Defense Conference: Protecting Earth from Asteroids - Orange County, California ()] 2004 Planetary Defense Conference:](https://reader031.documents.pub/reader031/viewer/2022020615/575095271a28abbf6bbf5d50/html5/thumbnails/10.jpg)
