The direct impact of a spacecraft into a threatening object may be a feasible NEO deflection method. It relies primarily on ejecting surface material off the object surface, which changes the linear momentum of the hazardous object and its subsequent trajectory. But that ejection process depends on the structure and size of an asteroid and has not been studied in any detail. We have initiated a laboratory, computational and theoretical effort to study that problem. The fundamental physics for the effects of an impact into an asteroid or comet is simply the balance of momentum. If the projectile has mass m and initial velocity, U, relative to the asteroid, the velocity change of the asteroid (of mass M) is âmU/M where â, the momentum multiplication factor, is >1. If a projectile buries itself in the target and no material is thrown out, the event is 'perfectly plastic' and â=1. However, a hypervelocity impact usually blasts out a crater many times the size of the impactor. Material is ejected at fairly large velocity, with a substantial component normal to the local surface. Therefore the total impulse imparted to the target body has two parts: the "primary" component from stopping the projectile, with â=1, and the additional component from the ejected material, which gives â>1 (normal to the surface). Depending on the mass and speed of the ejecta, the total transferred momentum can be significantly greater than the direct momentum of the projectile, i.e. â can be significantly larger than 1.

Laboratory experiments can be used to measure â directly, but only for small targets and impact velocities ranging to 6-7 km/s. Therefore, one must extrapolate (scale) the experimental results, which can be done using the scaling theory of hypervelocity impacts. The theory shows that â is proportional to the impact velocity, U, raised to an exponent that depends primarily on the porosity of the target material. The exponent ranges from 0 (independent of impact velocity) to 1 (â increases linearly with U). The theory also shows a power-law relation between â and the surface gravity g, with the exponent ranging from -3/8 to 0. Given that a deflection mission would involve a much larger U and smaller g than typical laboratory experiments, the value of â in the mission could be much larger than measured in the lab. We will summarize the initial results of our experiments and code calculations with the objective of measuring â for various geological materials under a variety of impact conditions and to test the scaling theory.

Yarkovsky and YORP effects are important factors of asteroid evolution. Yarkovsky effect changes orbits of asteroids and drives them to resonances, thus responding for renewal the population of near-Earth asteroids. YORP effect changes rotational state of asteroids, leads them to disruption limit, thus forming distribution of small asteroids over sizes and shapes. As well by changing rotation rate and orientation of rotational axis YORP influences Yarkovsky drag. In addition to its theoretical implications Yarkovsky effect has an important practical implementation: it can be used to change orbits of dangerous asteroids by changing properties of their surfaces. A common approach to Yarkovsky and YORP effects calculations is integrating of light pressure recoil forces over surface of an asteroid. Due to limited resolution of available models, these integrations account only for large-scale structures, commonly larger than several meters. In our report we concentrate on influence of intermediate-scale structures, roughly centimeter to decimeter-sized stones. They appear to qualitatively change the picture seen from large scales. Different mean temperatures of eastern and western sides of stones create a tangential YORP drag, so that even a spherical asteroid evenly covered with spherical stones can undergo YORP acceleration. Yarkovsky effect is also influenced by intermediate-scale surface structures, so that smoothing of the surface under modeling is not entirely justified. We study these effects in a simplified model for asteroid surface and discuss possible output.

Of the two missions to successfully land spacecraft on asteroids (NEAR and Hayabusa), both landings occurred with a large uncertainty in position. Indeed, the landing position of NEAR has been estimated with several different models and results differ by hundreds of meters (Veverka, 2001). Such inherent inaccuracy in positioning can be directly attributed to localization and navigation techniques, and to technology that lacks the required accuracy.

Unfortunately, many asteroid mitigation techniques require the precise landing of some kind of vehicle, or precision targeting on the surface of an asteroid. A brief listing includes surface or subsurface detonation (Barbee, 2007), (Weaver, 2010), mass-driver landers (Olds, 2007), and even gravity-tractors (Basart, 2009). To date, many of these mission concepts have assumed positioning as a given and do not consider the challenges of navigation without GPS or other external references. Furthermore, any asteroid fragmentation technique, if performed inaccurately, could result in pieces traveling through dangerous keyholes for impact opportunities in the future. With the current methods used for surface positioning, it is likely that these methods might fail when needed most or might not be as efficient as required. As motivation, we attempt to survey the effects of ill-positioning upon these techniques, which leads to the conclusion that improvements in small-body surface navigation must be acquired.

In this paper we consider methods to improve the navigation and localization of both orbiters and landers on a small-body object. We employ the use of radio-ranging through an active transponder (or surface vehicle) and show that stochastic methods can reduce the position uncertainty to several meters even at a relatively large range. Specifically we propose an improvement on navigation through the addition of accelerometers and an active transponder. Unlike previous radio-ranging methods, we remove the assumption that the spacecraft must be in a circular orbit and allow it to hover or move toward the surface. The addition of accelerometers (either on the orbiter/lander or on a surface explorer) provides additional inertial information to help characterize the absolute position of the vehicle and radio-ranging provides relative position information with respect to the asteroid.

To illustrate these methods, we use performance parameters characteristic of the Hayabusa mission and demonstrate a landing scenario with a single and multiple transponders. We extend this illustration to examine the use of this technology for high-speed collisions (impactors or driven explosive charges) or constant orbiting (gravity tractor). This illustration leads to recommendations of technologies to be used for future asteroid mitigation missions.

By combining known asteroid mitigation techniques with improved navigation technology, we can better assure the success of planetary defense missions. This technology must be considered when planning mission concepts and cannot be forgotten as a given. The methods provided in this paper use present-day technology and have shown in simulation to significantly enhance positioning accuracy. This enhanced positioning should help ensure the safety of the Earth.

NEO missions for deflection (impact or slow-push) or characterization (in-situ or sample return, robotic or manned) present different objectives and constraints. Lessons learnt in different studies during last decades are presented for application to the design of different types of NEO missions. Interplanetary trajectory performances are critical for mission design and feasibility analysis. Not all scientifically interesting asteroids can be reached with reasonable flight time (<10 years) and moderate delta-V (few km/s). Chemical propulsion is the conventional option but multiple gravity assists are usually needed. We will be show that for robotic missions electric propulsion increases significantly the mission flexibility (launch date, reacheable NEO) and the final arrival mass. For trajectory design, sample return missions are more challenging since impose more constraints such as stay time, re-entry velocity. In addition, there are more manoeuvres that accumulate in the delta-V budget. Manned missions have similar constraints than robotic sample-return missions. Crew imposes tighter constraints on the system and on the mission duration and safety, which drive the design. The trajectory optimization shall consider emergency situations and navigation aspects. Human missions can reach a smaller set of NEOs as will be shown after systematic computation of human missions to all asteroids in JPL database (over 400.000) within a 30-year launch window. In the approach phase, the detectability of the target is one of the critical aspects, above all in impact and manned missions. It shall be considered in the asteroid selection process. The approach strategy and the Guidance, Navigation, Control (GNC) system must be carefully designed to attain the objectives (impact, close station insertion). The irregular shape of small NEOs produces variables in the light-curve that introduces large uncertainty in the visibility conditions. Therefore, the approach strategy, the image processing algorithm and the GNC system shall be robust against such uncertainties. The proximity operations phase is critical because of the large uncertainties about NEO properties. Different sub-phases shall be considered to achieve different science objectives. The schedule shall be designed to obtain the required knowledge before proceeding to riskier operations. In this phase, GNC aspects in conjunction with ground control procedures are critical to assure the safety of the mission and the achievement of the scientific objectives. In addition, we will show some design options that permit the increase of the autonomy periods, reducing the operational costs without impact on the safety. The different design options for diverse mission and technology constraints are assessed in a high-fidelity simulator. A sequence of proximity-operations modes is presented that fulfils all the safety and scientific requirements within the boundaries of the system and technological constraints. The proximity operations modes include (1) far station keeping (without detailed knowledge of asteroid properties), (2) self-stabilized terminator orbit (best option for radio-science experiments that will use the knowledge gained in previous station keeping phase), (3) closed controlled orbits (including transfers), (4) close station keeping, and (5) descent and landing (D&L). These modes can be designed to accommodate different mission objectives and might be re-ordered and tuned in-flight in order to increase the science return or the mission success probability. One of the most relevant proximity operations is the D&L. In this mode, autonomous operations are mandatory below a certain point. Given the typical small size of NEO and the uncertainty in the surface characteristics, pin-point landing is envisaged (several meters). The results of high-fidelity simulations will be summarized to show the capabilities of the GNC system (critical during autonomous operations).

Nuclear devices have long been discussed as a hazard mitigation option for potentially hazardous objects (PHOs), either for deflection or disruption of the object (e.g. Teller et al. (1995), Ahrens and Harris (1994), and Holsapple (2004)). The technical and geopolitical challenges associated with the potential deployment of a nuclear device for this purpose necessitate a careful analytical exploration of the response of hypothetical PHOs to shocks and radiation, well in advance of an impact threat.

We will present radiation hydrocode models of thermal and kinetic energy deposition onto simulated PHO surfaces and particle transport code models of the absorption of nuclear radiation into similar surfaces. With these models we begin to estimate the momentum imparted to the PHO by the burst, given PHO composition, shape, and height of burst above the PHO surface. We find that, as expected, the thermal radiation is deposited within a few centimeters of the burst-facing surface. Part of this thin surface layer is vaporized and heated to temperatures as high as 10,000 K. This results in vaporization efficiencies between 1-30% of the theoretical maximum possible vapor given the amount of incident energy and the thermodynamic properties of the target (Barnes and Lyon (1988), Clark (1984)). The vaporization efficiency depends on the height of burst and heat capacity of the target material, among other factors.

The nuclear radiation, which mediates about 3% of the energy released by the device, is absorbed at a much greater depth than the thermal radiation. Our initial models of nuclear radiation deposition were calculated using the MCNP Monte Carlo particle transport code, and geometries similar to those used in the hydrocode models. These models predict that neutrons are absorbed down to a much greater depth than the x-rays. The exact depth depends on the nuclear cross-section of the target material, but may be in excess of one meter for the materials considered. This thicker deposition region has the potential to result in a larger amount of material being mobilized from the target surface, both as vapor and solid debris, which increases the transfer of momentum to the object (similar to Walker and Chocron (in press, doi:10.1016/j.ijimpeng.2010.10.026)).

Preliminary models for an 11 kt stand-off burst at a height of 10 m above the surface of a 500-m-diameter target predict that the momentum imparted to the PHO in such a scenario would be on the order of 1.0x10⁹ N·s. We are using these models to place constraints on the amount of momentum that would be imparted to a PHO by a nuclear burst, and how it would be affected by variations in the composition and shape of the PHO, and the positioning of the burst relative to the object.

Recent studies by the US National Academy of Science identified nuclear explosives as the only current technology able to deflect large asteroids (> than 300-500 meters) or to mitigate impacts of smaller bodies when the warning time is short. Previous modeling of inhomogeneous asteroid sized bodies, have shown the ability of either standoff bursts or very low yield surface bursts to impart velocity perturbations sufficient to deflect a body one or more decades prior to impact, but concerns continue to be expressed that nuclear explosives are either to strong risking fragmentation, or insufficient providing too little impulse for very porous bodies. Amid such claims, the Office of Science and Technology Policy (OSTP) released a letter to the US congressional committee on Science and Technology expressing a belief that "significantly more analysis and simulation are needed". We begin to address these concerns by studying the limiting affects of material porosity, and the limits in which fragmentation will occur.

Despite the lack of a known immediate threat from a near-Earth object (NEO) impact, historical scientific evidence suggests that the potential for a major catastrophe created by an NEO impacting Earth is very real. It is only a matter of when, and humankind must be prepared for it. During the past two decades, various concepts and techniques for mitigating such impact threats from NEOs have been proposed. The non-nuclear techniques, such as slow-pull gravity tractors and kinetic energy impactors, require mission lead times much larger than 10 years, even for a relatively small NEO. However, warning time of an asteroid impact with the Earth can be very short. For realistic worst-case mission campaigns with a warning time of less than 10 years, a direct intercept mission employing nuclear explosives becomes the only viable option. Direct intercept missions have approximately 2-30 km/s arrival velocities with respect to target asteroids. A rendezvous mission requiring a large arrival Delta V of 2-30 km/s is impractical. Although a nuclear surface or standoff explosion can be employed for such direct intercept missions, the momentum/energy transfer caused by a subsurface nuclear explosion is about 10-100 times more effective than a surface or standoff nuclear explosion. However, existing penetrator technology limits the impact velocity to less than 300 m/s because higher impact velocities destroy prematurely the detonation electronic equipment. Thus, hypervelocity nuclear penetrator technology needs to be developed for the last-minute intercept missions with impact velocities as high as 2-30 km/s.

This paper presents the preliminary results for assessing the technical feasibility of hypervelocity nuclear interceptors employing various nuclear explosion approaches in the presence of realistic technological constraints as well as physical modeling uncertainties. This paper also examines the uncertainty effects of a wider range of asteroid material properties on the nuclear explosion approaches. In particular, this paper focuses on a nuclear interceptor concept for subsurface explosion missions, which exploits a two-body space vehicle consisting of a fore body (leader) and an aft body (follower). The fore body provides proper kinetic impact crater conditions for an aft body carrying nuclear explosives to make a deeper penetration into an asteroid body. The surface or standoff explosion missions do not require the separate fore body; however, they will require accurate timing of the nuclear explosive detonations near the surface of a target asteroid. This paper also examines the technical feasibility of hypervelocity intercept missions employing a standoff or surface nuclear explosion in the presence of realistic technological constraints as well as mission uncertainties.

It is currently estimated that 16% +/- 4% of the Near-Earth Asteroid (NEA) population is comprised of binary asteroid systems. However, relatively few Near-Earth Object (NEO) mitigation papers have investigated deflection strategies for such systems. Koenig and Chyba have discussed applying a ÄV for breaking, merging, or only deflecting binary systems. Fahnestock et al. have also looked into the control and performance of gravity tractor operations at binary asteroids, concluding on a performance penalty compared to unary objects that is manageable using a simple control algorithm. This paper aims to go a step further providing an analysis of the deflection achieved by a Gravity Tractor (GT) acting at one of the system's components. We find that the net deflection on the binary system's center of mass depends on the relative position and mass of the components. As expected, if the binary is made of a large primary and a small moon, station keeping at the primary yields similar results to a single asteroid system. We present trends for a range of different binary systems, varying mass ratios, component sizes, and dynamical parameters.

We further investigate a GT that continuously pulls the asteroid satellite along its orbital velocity relative to its primary such that the satellite would eventually have enough energy to escape the system. Reaching mutual escape may require a relatively large change in velocity, involving a long period of tractoring. For an Apophis-like system with a hypothetical 30 m diameter moon orbiting 500 m from the primary, it would take a 1000 kg spacecraft less than 10 years to achieve mutual escape if tractoring constantly. We find that even though a large deflection can be delivered to the moon due to its smaller mass, relatively little deflection is delivered to the binary center of mass compared to the strategy of tractoring the primary along its heliocentric velocity vector.

Finally, we investigate the influence of a GT on the binary system dynamics itself, and we present analytical estimates and numerical simulations for a range of binary asteroid systems. During GT operations, we examine the orbital elements and parameters of the binary components changing with time, and the associated change in system energy and stability.

The NASA Exploration Systems Mission Directorate received a presidential mandate to send humans to explore a Near-Earth Asteroid (NEA). In studying potential architectures, NASA convened a "NEA User Team" to provide definitive guidance to the human spaceflight community, by identifying the means by which humans can productively meet exploration goals and objectives. Beginning with preliminary goals and objectives, generated by attendees of a NASA-sponsored workshop, the NEA User Team developed concepts of operations for a set of preliminary design reference missions. The concepts of operations provided the foundation for establishing that the preliminary goals could be met, identifying drivers for meeting those goals, and identifying the information that would be required about these objects to inform mission design and planning and to meet safety requirements and decrease risk. One design reference missions included a 30 day stay at the asteroid and utilized a dedicated space exploration vehicle to allow humans to directly interact with the surface. At the other end of the spectrum was a minimalistic design reference mission that presumed a 14 day stay and no direct human interaction with the surface of the NEA.

The NEA User Team subsequently used this filtering process to identify the critical knowledge required prior to sending humans to any NEA, and particularly for the first human mission. For example, navigation near an asteroid and attachment to the surface present unique challenges in an extremely low gravity environment. To maximize the probability of mission success, the NEA User Team determined initial estimates for the acceptable spin rate and tumble characteristics for viable NEA targets. Precursor robotic exploration was identified as the required means to validate the basic pre-qualification of any potential destination NEA. At the same time, the NEA User Team worked to identify those known asteroids that could be accessible within certain mission parameters. Sorting known objects by a set of initial filters, the resulting sets were analyzed for total mission delta-V, mission duration, mass to orbit, and physical characteristics related to the object itself.

The paper will summarize the tasks completed by the NEA User Team, and the results obtained, including information essential to qualify any object as appropriate for the first human mission to a Near-Earth Asteroid.

We present a novel concept (*) to impart a continuous thrust to an Earth threatening asteroid from a hovering spacecraft with no need for physical attachment nor gravitational interaction with the celestial body. The spacecraft, placed at a distance of a few asteroid diameters, directs a highly collimated stream of quasi-neutral plasma against the asteroid surface using an ion thruster. The impact of the ions with the asteroid surface results into a net transmitted momentum slowly deflecting the asteroid. At the same time, a second propulsion system is employed to compensate for the ion engine reaction and keep the distance between the asteroid and the shepherd satellite constant throughout the deflection phase. After describing the concept in details we optimize the system design for minimum spacecraft mass and estimate its deflection capability for asteroids of different size and orbits. As the force exherted is independent of the asteroid mass and size the method allows deflecting subkilometer asteroids with a spacecraft more than one order of magnitude lighter when compared to a gravity tractor spacecraft of equal deflection capability. This improvement in performance has the potential of making collisionless asteroid deflection missions feasible in the near future.

(*) Bombardelli, C. and Pelaez J. "Ion Beam Shepherd for Asteroid Deflection"; submitted to the Journal of Guidance, Control and Dynamics (July 2010).

Hayabusa spacecraft finally came back to the earth on June 13, 2010. It has a quite dramatic story and we had a lot of experiences about planetary mission. And now we have started Hayabusa follow-on mission, Hayabusa-2. It is an asteroid sample return mission again, but the type of the target asteroid is C-type, which is different from the target of Hayabusa, Itokawa (S-type). The target asteroid is 1999 JU3, which is one of the NEOs. The estimated size is about 900m in diameter. So we will get another data of small NEO. The scale of the Hayabusa-2 spacecraft is similar to Hayabusa, but many parts will be modified so that we will not have the troubles that we experienced in Hayabusa. Also the spacecraft has new equipment, which is called impactor. The impactor will make an artificial crater on the surface of the asteroid, and we will try to get the sample inside the crater. Then we can get much fresh material. The planned launch year is 2014 or 2015, arriving at the target asteroid 1999 JU3 in 2018, and coming back to the earth 2020. The budget of Hayabusa-2 has been approved by the Japanese government, so we can start manufacturing it from April 2011.