The mission scenario AsteroidSQUADS/iSSB was developed in response to several of the Recommendations from the 1st IAA Planetary Defense Conference which addressed the need for deflection-related testing and campaign design, studies of momentum transfer in impulsive deflection techniques, and the development of protocols and responsibilities within existing space situational awareness and civil defence infrastructures on a global scale. Several more recommendations put dangerous objects smaller than the current threshold definition for Potentially Hazardous Objects (PHO) in focus. Throughout, the need for increased international participation was emphasized.
AsteroidSQUADS/iSSB is intended to enable Serendipitous Quantitative Understanding and Assessment of Deflection Strategies. The advantages and efficiency of modern small-satellite-derived design philosophies evolved and improved for interplanetary spaceflight are highlighted by using the DLR Kompaktsatellit programme's Standard Satellite Bus kit as a study baseline. This spacecraft platform draws strongly on the experiences gained and lessons learned from the DLR small satellites BIRD and TET. It also has been the baseline of choice in many studies at DLR's Bremen Concurrent Engineering Facility, and it is currently used for the AsteroidFinder spacecraft under development.
A number of circumstances in today's commercial and scientific spaceflight environment are on their own widely regarded as detrimental or unpleasant situations: For some time now, an uneasy struggle has developed between the test flight requirements of the heavy launch vehicle sector, related costs and risks, commercial and schedule pressures, public relations and insurance contracting concerns, and the choice and motivation of payloads for such development flights. Also, realistic testing in particular of geostationary payload launch vehicles carries a significant risk of polluting the most vital regions of Earth-orbital space with large targets that invite the escalation of space debris collisional cascading (Kessler syndrome). In the planetary science sector, it has always been difficult to obtain funding for missions towards less prestigious target objects in the solar system. For most such missions, target selection was severely constrained by the need to cover as many fields of science as possible within the given launch budget. Resulting spacecraft designs push the launcher performance limit and require gravity-assists from the nearest planets. Accordingly, rather small probes experience extended interplanetary cruise phases, causing high radiation doses on sensitive components and high operational cost. In planetary defence, with few exceptions, lesson-learning has so far been restricted to paper exercises. Though NEO surveys have generally made good progress given the resources assigned, even the basic methods of deflection are hardly explored beyond lab experiments. The AsteroidSQUADS/iSSB mission scenario seeks to benefit from several opportunities which are presented by these situations when their mere co-existence is turned into a synergetic advantage for all potential participants, including all branches of the planetary defence community. It employs a flotilla of simple multi-role spacecraft directed at a suitable sub-PHO size practice target for a brief but intense integrated deflection campaign exercise in real space.

Missions to Near-Earth Asteroids (NEAs) offer a wide range of possibilities for space exploration, scientific research, and technology demonstration. In particular, manned missions to NEAs provide a unique opportunity to be the first human expedition to an interplanetary body beyond the Earth-Moon system and represent the perfect environment to gain experience in deep-space operations, which is an indispensable prerequisites for human missions to Mars. As a starting point for the analysis of such missions, the objectives of this study are to identify target asteroids and evaluate possible transfer trajectories as well as the associated launch windows. Subsequently, it is investigated whether different missions to the identified targets exhibit commonalities in terms of out- versus inbound flight times. Lastly, the possibilities of mission abort are examined. The list of accessible asteroids is narrowed down by taking dynamical and structural properties such as size and rotation rate into account. An accessibility model for NEAs is developed allowing pre-selection of asteroid targets for human missions. For this model, a novel approach is taken which assesses the accessibility of a NEA not by considering its orbital parameters separately. Instead, accessibility is determined by evaluating the combination of all orbital parameters only limited by mission duration (less than 365 days) and round-trip ∆v (less than 10 km/s). In order to verify the reliability of the model, mission architectures for missions departing from low-Earth orbit are investigated and transfers to 1947 NEAs in the time frame from 2020 to 2040 are simulated. 71 asteroids are found to be accessible for human missions under the given boundary conditions and are observed to nicely fit the model developed. 32 of these remaining 71 asteroids can be reached with a ∆v iü 7.5 km/s, seven of which allow mission durations of less than 180 days. 81 launch windows are found for these 32 asteroids between 2020 and 2040. Launch opportunity analysis shows that most years in the given time frame offer more than one launch opportunity for missions with durations of less than 365 days and ∆v iü 7.5 km/s. Commonalities or patterns are observed in missions to different asteroids. It can be seen that 41 launch windows only permit missions with a longer outbound than inbound flight (pattern I) while 26 launch windows offer missions for which the duration of the inbound flight exceeds the duration of the outbound flight (pattern II). During each of the remaining 14 launch windows, it can be freely chosen whether to fly the long or the short leg first for the same overall mission duration. All 32 NEAs can be reached on pattern I missions. 20 asteroids can in addition be visited on pattern II missions. Missions with an outbound trajectory of a 365 day period offer free return possibilities, however, only few such missions to very few asteroids exist in the 2020 to 2040 time frame. Mission abort with an anytime return scenario is more promising in terms of safety but causes ∆v penalties. With these results, the mission analysis of the interplanetary part of human missions to asteroids is concluded, setting mission-specific requirements and boundary conditions required for subsequent spacecraft design.

NEO hazard mitigation is currently an area of interest and a number of different deflection strategies (nuclear interceptor, kinetic impactor, mass driver, low-thrust propulsion, solar or laser ablation, gravity tractor, etc.) have been proposed. Deviation methods such as nuclear interceptor or kinetic impactor which physically interact with the asteroid and solar concentrators which volatilize surface materials are strongly dependant on not only the physical characteristic but also the chemical constitution and properties of the asteroid. It is therefore crucial that a sufficiently comprehensive analysis of the chemical properties of target asteroids is performed to enhance any hazard mitigation mission. Although ground-based or space-based NEOs observations are capable of identifying basic physical properties of asteroids (mass, size, mean density, porosity, albedo, etc), specific chemical compositions -olivine, metal, feldspar, orthopyroxene, etc.- can be accurately determined only by close proximity observations or in-situ sampling and analysis. This however may not always be feasible particularly when lead times for a deflection mission are short. The focus of this paper is therefore to evaluate the robustness of different deviation methods to various NEO chemical compositions. Initially, the composition of Richardton meteorite is used as baseline NEO composition. Evidence Theory is then used to model uncertainties with regards to chemical compositions of the baseline asteroid and three deflection missions based on nuclear interceptor, kinetic impactor, and solar concentrator are modeled. Each deflection strategy is then applied to a set of virtual impactors (i.e. a variety of NEOs with different Keplerian elements). In this paper, the chemical composition of the set of virtual impactors is modified to investigate how NEO composition affects a change in deflection methodologies. Finally, the total asteroid velocity change and deflection distance achieved by each of these strategies are compared to evaluate the robustness of these strategies over a range of different asteroid chemical properties and different physical parameters which have a bearing on the deflection methodology, such as, for example, the rotational state of the asteroid.

The gravity tractor concept is a proposed method to deflect an imminent asteroid impact through gravitational tugging over a time scale of years. In this study, we present mission scenarios and operational considerations for asteroid mitigation efforts involving multiple gravity tractors. Among mission concepts, we investigate the scenario of sending a small precursor gravity tractor to an asteroid suspected of being on course to enter a gravitational keyhole followed by a larger subsequent gravity tractor if an imminent keyhole entry is confirmed. Ground observations can typically only estimate the probability of a keyhole entry or impact but can be improved significantly with spacecraft radio tracking after rendezvous with the asteroid. If a deflection is determined necessary, the precursor can perform gravity tractor duties to deflect the asteroid but might require an additional dedicated gravity tractor to provide sufficient deflection if beyond the capability of the precursor. Multiple gravity tractors are therefore required for mitigation missions in which the required deflection exceeds the capability of a single launch vehicle. We explore additional scenarios requiring multiple gravity tractors such as the need for redundancy to mitigate a high-consequence impact through the use of independent launch providers, spacecraft contractors and international collaboration. We further assess operational capabilities enabled by multiple-spacecraft gravity tractor missions such as accommodating redundant spares in proximity of an asteroid and the dedication of one spacecraft to monitor the deflection through radio tracking. The combined effect of multiple gravity tractors operating simultaneously is described: we find that the total b-plane deflection is the sum of the deflections caused by each gravity tractor if it were acting alone. We also discuss proximate operations for multiple gravity tractors since co-location is desirable to maximize deflection but one must mitigate any chance of collision. One solution requiring autonomous control consists of phasing gravity tractors around a halo orbit centered on the optimal standoff location. We subsequently investigate the relative benefits and drawbacks of mechanically docking gravity tractors in proximity of the asteroid and include a discussion on design considerations that allow docked gravity tractor operations. In the context of gravity tractor mitigation, we conclude that a confirmed asteroid threat would necessitate multiple spacecraft for redundancy and to provide sufficient deflection if beyond the capability of a single gravity tractor. A multiple-spacecraft mission involving international collaboration is also explored for an impact with worldwide consequences. Finally, we discuss operational considerations and utilize 99942 Apophis as an illustrative example.

A credible potentially hazardous object (PHO) threat may be identified sometime in the future, requiring actions to be taken to prevent an impact disaster. To negate that threat, mitigation techniques are being proposed where the potential for collision is unacceptably high. The Aerospace Corporation is developing a Handbook for Near Earth Object (NEO) Deflection and a complementary web-based NEO Deflection interactive tool. The purpose of the tool is to aid in the design and understanding of the deflection impulses necessary for defending Earth against threatening objects and in the analysis and comparison of various techniques that might be used to provide those impulses. The Handbook and the associated web-based resource center will provide first-order requirements for effective NEO deflection missions using a variety of deflection concepts. The resources will include educational materials on NEO threats and deflection concepts, as well as examples demonstrating the use of the Handbook and web-based tool. Project overview and status is presented.

A great amount of practical knowledge, both on the scientific and technical side, has been acquired in the past and ongoing NEO mission studies in ESA (Don Quijote, Marco Polo, NEMO, Proba IP, NEOMEX, NEMS, etc). This understanding is very useful to help choose asteroid targets, estimate likely orbital dynamics constraints, identify operational needs, make realistic technology assumptions, and exploit synergies between missions with different goals i.e. technology demonstration, science, exploration, impact risk assessment and impact mitigation, whenever possible. The discussion involves several levels of detail, both at scientific and technical level. The first one is related to the object’s orbital and dynamic characteristics, and how they drive the mission design; in a first order approximation, detailed consideration on the spacecraft system design are not needed, but still some rather generic “guiding principles” can be derived that affect system aspects with key programmatic importance, such as the total mission delta-v and the launcher performance, the propulsion technology options and the mission duration. While some of these principles are common to most interplanetary missions, others are particularly linked to the asteroids’ orbits, such as the variation of heliocentric and geocentric distances during the mission, the constraints on return trajectories (for sample return and crewed missions), or the efficiency of the momentum transfer for e,g, kinetic impactors. In the next level of detail, dynamical characterisation of the object i.e. binarity, rotation period, orientation of spin axis, etc would drive some mission operations aspects, such as the approach to a NEO and the spacecraft ability to move and perform operations in the vicinity of the body. Issues like precise gravity field and shape models would then help refine orbiting or co-flying strategies, and derive the spacecraft guidance and navigation requirements and configuration constraints, among other. Some of these factors also need to be addressed early in the planning stages e.g. to initiate the related technology development activities. A last layer might be defined in relation to the type of information that the mission is expected to gather, which can be linked to that specific mission and its objectives (e.g. solar system science, etc) but also to more generic practical knowledge that would be useful for other future NEO missions. The discussion is in this case would be focused on the spacecraft instruments required to gather such generic knowledge and to obtain lessons from the environment (topography, thermal, radiation, dust etc) solely for operational purposes. This analysis also includes the technology demonstration opportunities linked to NEO missions. All these aspects would need to be considered in a programme using an incremental capability development approach to the exploration of these bodies or to the built-up of confidence on impact mitigation systems. In summary, we will discuss some examples on how our current physical understanding of the NEOs and of past technical trade-offs results can greatly help define the future trade spaces, and support the selection of baseline mission options. This is an analysis with a high degree of multidisciplinary and which is key to decision-making n the definition of NEO mission campaigns.

Out of all the NEO's found to date the asteroid 99942 Apophis appears to be the most likely to impact Earth. An impact from Apophis appears unlikely, with an estimated impact probability of approximately four-in-a-million. On April 13, 2029 Apophis will pass by the Earth within geostationary orbit. If Apophis passes through a relatively small 600 meter keyhole, impact will occur on April 13, 2036. The orbit and physical attributes of Apophis could be further refined through either a robotic or human exploration mission. Therefore, the purpose of this paper is to perform the mission design for robotic and human exploration mission to Apophis, using software developed by Asteroid Deflection Research Center (ADRC). Possible launch windows, trajectories, and accompanying delta-V's for both robotic rendezvous and human piloted return missions prior to the April 13, 2029 Earth-Apophis close encounter will be analyzed. In addition mission analysis and design will be performed for robotic and human piloted missions for a scenario in which Apophis passes through the keyhole, resulting in and impact on April 13, 2036.
Using the system capabilities of the Interplanetary Ballistic Mission (IPBM) architecture [IAC-09.D1.1.1] a total of 16 launch windows were found from mid-2011 up the April 13th, 2029 close encounter. For the early robotic missions only direct 0-revolution transfers were considered. This allowed a significant number of launch windows with low minimum delta-V's. The possible launch windows occur approximately every 7 years, corresponding to periods when Apophis is closest to the earth. The launch windows found range from 4-142 days in lengths with delta-V requirements from 2.268 km/s up to the 4 km/s limit. The fictional post 2029 mission is over a 7 year period, which requires mission designs other than 0-revolution direct transfer missions. For this part of the study multiple revolutions around the sun were considered in an effort to open up more launch windows. A a total of 18 launch windows were found during the 7 year period [AIAA 2010-8375]. The delta-V's required for these missions ranged from 2.169 km/s up to the limiting 4 km/s, with launch windows ranging from 5 to 166 days in length. For the pre and post-2029 mission the limit of 4 km/s (from GTO) allows for nearly continuous launch windows for the direct intercept mission.
The mission analysis for a return human exploration mission shows that launch windows are only possible near Earth-Apophis close encounters, with the minimum delta-V missions occurring near the 2029 close encounter and just prior to the 2036 impact. In general two minimum delta-V missions are found for return missions, one returning near the close encounter date and the other departing near the close-encounter date. Analysis indicates that a one year mission results in the minimum required ÄV of 6.373 and 6.208 km/s for the early and late 2029 launch windows, close to that of a lunar mission. Shortening the mission to 180, which is the shortest mission considered in recent NASA studies, increases the delta-V required to approximately 12 km/s. In this situation a one year mission to Apophis could be used as a stepping stone between the first human NEO exploration mission and a mission to Mars. For the human piloted deflection mission launch dates only occur just prior to 2036. In this case only the launch window corresponding to the early launch window in the previous mission is feasible. Again analysis indicates that the minimum delta-V mission occurs when a mission length of 1 year is selected. In this case the required delta-V is increased to 8.436 km/s, still feasible with 2 Ares V launches. The Departure date is Approximately 1 year prior to impact with arrival at Apophis occurring 1 month and 10 days prior to impact. Recent studies appear to indicate this is sufficient time to significantly reduce the impact threat from using a high energy nuclear deflection method [AIAA 2010-8374].

The study of planetary defense has drawn wide interest in various research communities. Although the probability of large-scale impact event is small, the consequences of such an event would be disastrous. Study of the strategies available for protection against such occurrences provides insight into what scale of near earth object (NEO) we can hope to deflect using cutting-edge and near-future technologies. This study also pinpoints the bottlenecks that limit that scale. This paper discusses the use of a nuclear thermal rocket (NTR) as a propulsion device for delivery of thermonuclear payloads to deflect or destroy a long-period comet on a collision course with earth. We determine the worst plausible scenario for the available warning time (approximately ten months) and comet approach and use empirical data available to make an estimate of the payload necessary to deflect such a comet (as a function of its mass). Optimization of the trade off between early interception and large deflection payload, as well as mission launch date, establishes the ideal trajectory for an interception mission to follow. We also examine the potential for multiple missions launched at different times to contribute tothe deflection of the comet. Comparison of various propulsion technologies for execution of this mission shows that NTR outperforms all other candidates substantially. Finally, we discuss the size of comet it would be feasible to deflect using NTR propulsion, given currently available launchcapabilities into low earth orbit (LEO).