Aas 12-128 Conceptual Design and Analysis of Planetary

AAS 12-128 CONCEPTUAL DESIGN AND ANALYSIS OF PLANETARY DEFENSE TECHNOLOGY (PDT) DEMONSTRATION MISSIONS George Vardaxis,∗ Alan Pitz,y and Bong Wiez When the warning time of the impact threat of a near-Earth object (NEO) is short, the use of nuclear explosives may become necessary to safeguard the Earth. A variety of nuclear options, such as standoff, surface contact, and subsurface explosions, for mitigating the impact threats of NEOs have been proposed and studied in the past two decades. Eventually in the near future, an actual flight demonstration mission may become necessary to verify and validate the overall effectiveness and robustness of such various nuclear options and the associated space technologies. This paper presents the conceptual mission architecture design of such flight validation missions with a consideration of three mission cost classifications (e.g., $500M, $1B, and $1.5B). INTRODUCTION Given the past occurrences of asteroids and comets colliding with the Earth, it is necessary to prepare a global plan on how to mitigate the threat of a near-Earth object (NEO) on an Earth- impacting trajectory. During the past several years, research activities at the Iowa State Asteroid Deflection Research Center (ADRC) have focused on various nuclear options, such as standoff, surface contact, and subsurface explosions.1;2;3 The most effective approach is to use a penetrated subsurface explosion to deliver a considerable amount of energy to a small depth (< 5m) resulting in the possible total disruption of the target NEO. Depending on the mission lead time, a timely execution of a real NEO deflection/disruption mission can be a challenging task. When the warning time is short, the use of nuclear explosive devices (NEDs) will be the only option for generating a sufficient impulsive velocity change or to impart sufficient disruption energy to the threatening NEO. Such a last-minute intercept mission will result in a closing arrival velocity of more than 10 km/s. Because the current nuclear fusing mechanisms are limited to surviving impact speeds of less than 300 m/s, a hypervelocity nuclear interceptor spacecraft (HNIS) concept was conceived especially for penetrated subsurface explosions providing much more effective frag- mentation and dispersion of the target NEO.1;2;3 It is envisioned that eventually in the near future, planetary defense technology (PDT) demonstration missions will be considered seriously by an in- ternational space community in order to validate the overall effectiveness and robustness of various nuclear options and the associated space technologies. The PDT flight demonstration mission concepts studied in this paper fall into three budget classifications: $500M, $1B, and $1.5B. The ADRC’s mission design software tools have been utilized ∗Graduate Student, Asteroid Deflection Research Center, Dept. of Aerospace Engineering, Iowa State University. yGraduate Student, Asteroid Deflection Research Center, Dept. of Aerospace Engineering, Iowa State University. zVance Coffman Endowed Chair Professor, Asteroid Deflection Research Center, Dept. of Aerospace Engineering, Iowa State University. 1 to conduct a search for several target NEOs as well as perform preliminary mission designs.4 The required characteristics for target NEOs, to test the capabilities of the proposed HNIS, consist of: i) at least 100 meter diameter asteroids from the Amor group, ii) low mission ∆V requirements, and iii) hypervelocity intercepts. During proposed flight demonstrations, a small explosive device or a representative nuclear payload can be used as an alternative payload option. However, an actual NED would be the preferable experimental payload, to verify and validate the overall effectiveness and robustness of a space system to be employed in a real situation. Before getting into discussions of the target asteroids and the flight demonstration mission design, a previously proposed NEO deflection mission, known as the Don Quijote mission to asteroid 2002 AT4, will be briefly discussed in the next section. DON QUIJOTE MISSION CONCEPT To expand our knowledge on NEOs, the European Space Agency (ESA) decided to endorse six space mission proposals in July 2002.5 Of those mission proposals, the one given the highest priority was the Don Quijote mission concept, although the concept has not been realized later as an actual mission project. Mission Concept The Don Quijote mission was comprised of two satellites, an Orbiter and Impactor to be launched separately, to rendezvous with a target NEO. The Orbiter spacecraft, named Sancho, would be launched first and placed into an orbit around the NEO to precisely determine the orbital elements of the asteroid both before and after the impact by the second, impacting spacecraft. The Impactor, named Hidalgo, would be launched about two months after Sancho has successfully been placed into its observation orbit. The objectives of the mission were: i) to impact the NEO with the impactor spacecraft and be able to determine the object’s change in momentum after the impact, by measuring the mass, size and bulk density, and the variation in the asteroid’s center of mass orbital elements and rate of rotation, and ii) to perform multi-spectral mapping of the asteroid using an Autonomous Surface Package Deployment Engineering eXperiment (ASP-DEX). In addition, Sancho must be capable of measuring the deflection of any of the above characteristics to at least 10% accuracy and back up the guidance, navigation, and control systems of Hidalgo.5 Target Selection Based on the set of NEO characteristics defined for the Don Quijote mission, two potential target asteroids were selected - 2002 AT4 (baseline) and 1989 ML (back up). Their orbits are illustrated in Figure1. From the stand point of the Orbiter design, 1989 ML would be more accessible, but due to its larger mass it would be more difficult to perturb. Table1 shows the characteristics of both target asteroids. It was decided that the 2002 AT4 mission scenario would be used to size the Orbiter spacecraft, while 1989 ML would be used to size the Impactor spacecraft, therefore allowing for a more robust design that could be adapted to other targets.5 PRE-MISSION DESIGN PROCESS FOR INTEGRATED HNIS/OTV TRADEOFFS A multi-purpose, scalable configuration design of a baseline HNIS architecture is being per- formed at the Iowa State ADRC.6 A baseline HNIS architecture basically consists of its bus system 2 5 Figure 1. Orbits of 2002AT4 and 1989ML, Once Considered for Don Quijote Mission. Table 1. Characteristics of Target Asteroids of Don Quijote Mission.5 2002 AT4 1989 ML Orbital Period (yr) 2.549 1.463 e 0.447 0.137 i (deg) 1.5 4.4 Mission ∆V (km/s) 6.58 4.46 Orbit Type Amor Amor MOID large large Diameter (m) 380 800 and its NED payload. It may consist of two separable spacecraft: a leader spacecraft (impactor) and a follower spacecraft carrying NED payload for a penetrated subsurface explosion mission.2;6 The integrated HNIS/OTV (orbital transfer vehicle) design tool takes into account several parameters to decide the necessity of an OTV for the mission: launch vehicle, tank sizing, and fairing fit to produce a baseline mission architecture that would be suitable and applicable to a chosen target and mission. With the detailed design of the HNIS taken out of the mission design loop, there are a few less variables to deal with, but constrains the solution to work with the specific design. The pre-mission design software tool is comprised of several functions and subroutines calculating several OTV and preliminary design variables. Using information about the masses of the HNIS bus and NED payload, mission ∆V or C3 needed to reach the target NEO, and class of launch vehicles to be analyzed, the algorithm begins the process of calculating the payload capacity of the launch vehicles, the propellant mass of the OTV, size of the propellant tanks, if the payload configuration will fit in the fairing, and analyzing the solution. A flowchart of the pre-mission design process is provided in Figure2. The beginning of the design algorithm takes inputs about the HNIS bus system and NED payload, then requires additional data on the target NEO and mission parameters, and launch vehicles to be 3 HNIS and NED parameters Target NEO parameters Mission parameters Launch Vehicle YES NO OTV? Propellant Mass Calculation HNIS/ NO OTV fit? Viable YES NO Solution? YES Proceed to Mission Design Figure 2. Flowchart Illustration of the Pre-Mission Design Process. considered for the mission. With the pre-prescribed design of the HNIS, consisting of an impactor and a follower with NED payload, the program will ask whether the mission is a direct C3 injection orbit or if there would be an applied ∆V from the 185 km altitude circular parking orbit. If the indication is a C3 orbit, the program will ask for the class of launch vehicles to be analyzed for use in the mission. For the C3 orbit missions, if Delta II class launch vehicles are chosen, only the three-stage Delta II launch vehicles are considered because of their C3 payload capabilities. If a ∆V is to be applied, it tells the program that an OTV is planned on being used for the purposes of the mission, and the program will ask for the amount of ∆V required of the OTV. With all the given inputs, the program looks to see if the parameters indicate the need of an OTV. If not, the HNIS mass and dimensions are analyzed against the fairing sizes of the launch vehicles to ensure that it will fit inside the fairing and can be carried to the specified orbit. If there is a need for an OTV, the amount of ∆V needed enables the program to calculate the mass and 4 Table 2.

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