Deflection Missions Research Articles

Hazardous near Earth asteroid mitigation campaign planning based on uncertain information on fundamental asteroid characteristics

Given a limited warning time, an asteroid impact mitigation campaign would hinge on uncertainty-based information consisting of remote observational data of the identified Earth-threatening object, general knowledge of near-Earth asteroids (NEAs), and engineering judgment. Due to these ambiguities, the campaign credibility could be profoundly compromised. It is therefore imperative to comprehensively evaluate the inherent uncertainty in deflection and plan the campaign accordingly to ensure successful mitigation. This research demonstrates dual-deflection mitigation campaigns consisting of primary (instantaneous/quasi-instantaneous) and secondary (slow-push) deflection missions, where both deflection efficiency and campaign credibility are taken into account. The results of the dual-deflection campaign analysis show that there are trade-offs between the competing aspects: the launch cost, mission duration, deflection distance, and the confidence in successful deflection. The design approach is found to be useful for multi-deflection campaign planning, allowing us to select the best possible combination of missions from a catalogue of campaign options, without compromising the campaign credibility.

Closed loop terminal guidance navigation for a kinetic impactor spacecraft

A kinetic impactor spacecraft is a viable method to deflect an asteroid which poses a threat to the Earth. The technology to perform such a deflection has been demonstrated by the Deep Impact (DI) mission, which successfully collided with comet Tempel 1 in July 2005 using an onboard autonomous navigation system, called AutoNav, for the terminal phase of the mission. In this paper, we evaluate the ability of AutoNav to impact a wider range of scenarios that a deflection mission could encounter, varying parameters such as the approach velocity, phase angle, size of the asteroid, and the attitude determination accuracy. In particular, we evaluated the capability of AutoNav to impact 100–300m size asteroids at speeds between 7.5 and 20km/s at various phase angles. Using realistic Monte Carlo simulations, we tabulated the probability of success of the deflection as a function of these parameters and find the highest sensitivity to be due to the spacecraft attitude determination error. In addition, we also specifically analyzed the impact probability for a proposed mission (called ISIS) which would send an impactor to the asteroid 1999RQ36. We conclude with some recommendations for future work.

Rapid design and optimization of low-thrust rendezvous/interception trajectory for asteroid deflection missions

Asteroid deflection techniques are essential in order to protect the Earth from catastrophic impacts by hazardous asteroids. Rapid design and optimization of low-thrust rendezvous/interception trajectories is considered as one of the key technologies to successfully deflect potentially hazardous asteroids. In this paper, we address a general framework for the rapid design and optimization of low-thrust rendezvous/interception trajectories for future asteroid deflection missions. The design and optimization process includes three closely associated steps. Firstly, shape-based approaches and genetic algorithm (GA) are adopted to perform preliminary design, which provides a reasonable initial guess for subsequent accurate optimization. Secondly, Radau pseudospectral method is utilized to transcribe the low-thrust trajectory optimization problem into a discrete nonlinear programming (NLP) problem. Finally, sequential quadratic programming (SQP) is used to efficiently solve the nonlinear programming problem and obtain the optimal low-thrust rendezvous/interception trajectories. The rapid design and optimization algorithms developed in this paper are validated by three simulation cases with different performance indexes and boundary constraints.

Accurate analytical approximation of asteroid deflection with constant tangential thrust

We present analytical formulas to estimate the variation of achieved deflection for an Earth-impacting asteroid following a continuous tangential low-thrust deflection strategy. Relatively simple analytical expressions are obtained with the aid of asymptotic theory and the use of Peláez orbital elements set, an approach that is particularly suitable to the asteroid deflection problem and is not limited to small eccentricities. The accuracy of the proposed formulas is evaluated numerically showing negligible error for both early and late deflection campaigns. The results will be of aid in planning future low-thrust asteroid deflection missions.

Development of a handbook and an on-line tool on defending Earth against Potentially Hazardous Objects

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. A project overview and status are presented.

Indirect Optimization of Asteroid Deflection Missions with Electric Propulsion

Indirect Optimization of Asteroid Deflection Missions with Electric Propulsion

Ion Beam Shepherd for Asteroid Deflection

We present a novel concept to impart a continuous thrust to an Earth threatening asteroid from a hovering spacecraft without need for physical attachment nor gravitational interaction with the asteroid. The concept involves an ion thruster placed at a distance of a few asteroid diameters directing a stream of quasi-neutral plasma against the asteroid surface resulting into a net transferred momentum. As the transmitted force is independent of the asteroid mass and size the method allows deflecting subkilometer asteroids with a spacecraft much lighter when compared to a gravity tractor spacecraft of equal deflection capability. The finding could make low-cost asteroid deflection missions possible in the coming years.

Mission feasibility analysis on deflecting Earth-crossing objects using a power limited laser ablating spacecraft

This paper analyzes several mission capabilities to deflect Earth-crossing objects (ECOs) using a conceptual future spacecraft with a power limited laser ablating tool. A constrained optimization problem is formulated based on nonlinear programming with a three-dimensional patched conic method. System dynamics are also established, considering the target ECO’s orbit as being continuously perturbed by limited laser power. The required optimal operating duration and operating angle history of the laser ablating tool are computed for various types of ECOs to avoid an Earth impact. The available final warning time is also determined with a given limited laser power. As a result, detailed laser operating behaviors are presented and discussed, which include characteristics of operating duration and angle variation histories in relation to the operation’s start time and target object’s properties. The calculated durations of the optimal laser operation are also compared to those estimated with first-order approximations previous studies. It is discovered that the duration of the laser operation estimated with first-order approximations could result in up to about 50% error if the operation is started at the final warning time. The laser operation should be started as early as possible because an early start requires a short operating duration with a small operating angle variation. The mission feasibility demonstrated in the present study will give various insights into preparing future deflection missions using power limited spacecraft with a laser ablation tool.

Spacecraft formation flying for Earth-crossing object deflections using a power limited laser ablating

A formation flying strategy with an Earth-crossing object (ECO) is proposed to avoid the Earth collision. Assuming that a future conceptual spacecraft equipped with a powerful laser ablation tool already rendezvoused with a fictitious Earth collision object, the optimal required laser operating duration and direction histories are accurately derived to miss the Earth. Based on these results, the concept of formation flying between the object and the spacecraft is applied and analyzed as to establish the spacecraft’s orbital motion design strategy. A fictitious “Apophis”-like object is established to impact with the Earth and two major deflection scenarios are designed and analyzed. These scenarios include the cases for the both short and long laser operating duration to avoid the Earth impact. Also, requirement of onboard laser tool’s for both cases are discussed. As a result, the optimal initial conditions for the spacecraft to maintain its relative trajectory to the object are discovered. Additionally, the discovered optimal initial conditions also satisfied the optimal required laser operating conditions with no additional spacecraft’s own fuel expenditure to achieve the spacecraft formation flying with the ECO. The initial conditions founded in the current research can be used as a spacecraft’s initial rendezvous points with the ECO when designing the future deflection missions with laser ablation tools. The results with proposed strategy are expected to make more advances in the fields of the conceptual studies, especially for the future deflection missions using powerful laser ablation tools.

Optimal trajectories for the impulsive deflection of near earth objects

With the current technological level it is possible to deflect an asteroid such as 99942 Apophis in case its orbit refinement reveals an unacceptable collision risk. Thanks to the solution of an ad hoc defined global optimisation problem we show that a pre-keyhole deflection is possible even with a simple spacecraft and with a very late launch in 2026. We also give a mathematical description of the trajectory optimisation problems that have to be faced in the case of a more generic asteroid deflection mission defining analytically the objective function, that is the displacement in the encounter b -plane of the asteroid uncertainty ellipsoid.

Optimal deflection of NEOs en route of collision with the Earth

Recently, a method for the n-body analysis of the velocity change required to deflect a hazardous near-Earth object (NEO) was presented by Carusi et al. [Carusi, A., Valsecchi, G.B., D'Abramo, G., Boattini A., 2002. Icarus 159, 417–422]. We extent this method in order to optimize the velocity change vector instead of its along-track magnitude. From an application of both methods to a fictitious NEO we find Carusi's parallel approach to be reasonable for phases of unperturbed two-body motion. But, for orbit phases inhering third-body perturbations, i.e., for planetary close approaches or prior to a collision, the results obtained from the new method show the radial component of deflection impulse to play a major role. We show that a fivefold greater efficiency can be achieved by a deflection impulse being non-parallel to orbital velocity. The new method is applied to two possible 99942 Apophis impact trajectories in order to provide constraints for future Apophis deflection mission analysis.

3차원에서의 순간적인 속도변화에 의한 ECO의 최적궤도변경

순간적인 속도변화에 의한 ECO(Earth-Crossing Object)의 케도변경을 최적화하는 알고리즘을 개발하였다. 이를 통해, ECO의 궤도변경을 위한 속도변위를 계산할 때, 기존연구에서 간과되었던 궤도평면에 수직인 방향의 속도 변화를 살펴보았다. 이러한 3차원의 최적화 문제를 풀기위해서 순간적인 속도변화를 계산하기 위한 순간추력 근사법이 적용되었으며, ECO의 지구 접근 시에는 지구중력 효과를 고려한 부분적 궤도근사법을 사용하였다. 지구와 충돌천체의 상대적인 위치와 속도에 따라 ECO의 궤도변경을 위한 최적해가 달라지며, 그러한 최적해는 순간추력시간에 대한 최적속도변화나 최적비행각으로 표현될 수 있다. 순간추력시간이 작을 때, 궤도평면에 수직인 방향의 속도 변화를 무시할 수 없는 경우도 발견되었다. ECO의 궤도가 지구의 궤도와 비슷할수록 더 많은 최적속도변화가 필요로 하였으며, 순간추력시간이 충돌 순간에 가까워질수록 궤도변경에 필요한 최적속도변화의 크기가 지수함수적으로 증가하였다. 이러한 연구결과는 실제 ECO의 우주임무를 설계하는데 중요한 지침이 될 것이다. Optimization problems are formulated to calculate optimal impulses for deflecting Earth-Crossing Objects using a Nonlinear Programming. This formulation allows us to analyze the velocity changes in normal direction to the celestial body's orbital plane, which is neglected in many previous studies. The constrained optimization in the three-dimensional space is based on a patched conic method including the Earth's gravitational effects, and yields impulsive <TEX>${\Delta}V$</TEX> to deflect the target's orbit. The optimal solution is dependent on relative positions and velocities between the Earth and the Earth-crossing objects, and can be represented by optimal magnitude and angle of <TEX>${\Delta}V $</TEX> as a functions of a impulse time. The perpendicular component of <TEX>${\Delta}V $</TEX> to the orbit plane can sometimes play un-negligible role as the impulse time approaches the impact time. The optimal <TEX>${\Delta}V $</TEX> is increased when the original orbit of Earth-crossing object is more similar to the Earth's orbit, and is also exponentially increased as the impulse time reaches to the impact time. The analyses performed in present paper can be used to the deflection missions in the future.