AAS 12-128

CONCEPTUAL DESIGN AND ANALYSIS OF PLANETARY DEFENSE TECHNOLOGY (PDT) DEMONSTRATION MISSIONS

George Vardaxis,∗ Alan Pitz,† and Bong Wie‡

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 explo- sions, 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 demonstra- tion mission may become necessary to verify and validate the overall effectiveness and robustness of such various nuclear options and the associated space technolo- gies. 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 veloc- ity 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 clas- sifications: $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. †Graduate Student, Asteroid Deflection Research Center, Dept. of Aerospace Engineering, Iowa State University. ‡Vance 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 pay- load can be used as an alternative option. However, an actual NED would be the preferable experimental payload, to verify and validate the overall effectiveness and robustness of a space sys- tem 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 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 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 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: , 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 calculat- ing 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 con- figuration 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 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. Mission Design Comparisons for 2002AT4 Orbit Parameters Don Quijote PDT Demo Mission Parking Orbit 300 km 185 km v∞ 2.26 km/s 3.46 km/s Mass into Escape Orbit 790 kg TBD Launch Vehicle TBD proportions of the bi-propellant fuel. The two common types of bi-propellant used for the OTV are / (LOX/LH2) and nitrogen tetraoxide/hydrazine (N2O4/Hydrazine). Based on the choice of fuel type, the mass and capability of the fuel can be calculated. From there, the HNIS plus OTV configuration is then checked against the launch vehicle fairing sizes to see if the entire payload can fit inside. If the HNIS or HNIS/OTV configuration does not fit within the specified class of launch vehicles’ fairings, then a new class of launch vehicles will need to be specified for analysis. If the HNIS does fit within one of the launch vehicle fairings, then the algorithm has found a possible solution. With a set of solutions obtained by the HNIS/OTV design algorithm, each solution has to be analyzed to ensure its viability. The user enters the design-loop at this point, deeming a solution as either acceptable or not, and potentially restarting the entire design process if necessary. If a viable design is found from the resulting set of solutions then it can be taken and used to design the corresponding mission to a specified target NEO. Before getting into the specifics for the target NEOs for the missions described in this paper, a case study of 2002 AT4 is carried out to test the HNIS/OTV design algorithm. Table2 shows some of the orbital parameters of the ESA’s Don Quijote mission, along with the equivalent orbital parameters for the HNIS/OTV configuration mission - all other parameters are assumed to be the same. The sections marked to be determined (TBD) are left that way intentionally because the informa- tion that would be placed there is dependent on the mission budget class, and would be retrieved using the design algorithm discussed previously in this section. The payload and preliminary mis- sion parameters for the 2002 AT4 target using a baseline HINS/OTV configuration will presented in the next section, along with a slightly more in-depth discussion of the algorithm subroutines.

BASELINE HNIS/OTV DESIGN AND SELECTION OF LAUNCH VEHICLES

Mission Cost Classifications

A baseline HNIS/OTV mission design is considered for three different classes of mission costs: i) a $500M mission, ii) a $1B mission, and iii) a $1.5B mission. Each mission has an associated nuclear payload. The lowest budget class mission will be carrying a 300-kg nuclear explosive device (NED), while the middle and upper budget class missions will have a 1000-kg and 1500-kg NED payload, respectively. Previous studies assumed that a 1500-kg NED yields approximately 2 Mt of energy, a 1000-kg NED yields about 1 Mt, and a 300-kg NED yields about 300 kt.7 A substantial portion of the mission cost for each mission has been allocated to the launch vehicle to be used for the mission, and based on the total allowable monetary value associated to the missions, launch vehicles or classes of launch vehicles have been attached to each budget class.

5

Figure 3. A Baseline HNIS Consisting of Two Spacecraft Connected by Optional Deployable Booms.6

Table 3. NED Payload and HNIS Mass Summary. Mission Class Total Mass of HNIS (kg) 300-kg NED 1843 1000-kg NED 4251 1500-kg NED 5720

Baseline HNIS and OTV Design

Before discussing a case study mission to 2002 AT4 and the specifics of the various PDT demon- stration missions to other target asteroids, a brief discussion will be provided here on the design of the HNIS/OTV that would carry the associated nuclear payload. HNIS Design The HNIS design draws inspiration from the design of the space- craft, as well as similar spacecraft, with some noteworthy differences. A notable design character- istic is the HNIS configuration itself. The HNIS is made of two distinct sections (an impactor and a follower) connected via extendable booms, as illustrated in Figure3. This figure depicts a baseline HNIS that would be carrying a 300-kg NED to a target NEO. One purpose for having the design as such is due to its scalability for the two other mission scenarios, carrying a 1000-kg and 1500-kg NED to their respective targets. While the three spacecraft would be similar with respect to their design, they are unique in the sense that they are specifically designed to carry, house, and protect their NED . Table3 shows the wet mass of the HNIS for each designated NED payload, including the NED mass and spacecraft mass margin. The impactor spacecraft of the HNIS is designed as a blunt body because it will be the first part of the HNIS to impact the target asteroid at hypervelocity. Four-meter booms separate the impactor from the follower spacecraft during the terminal guidance phase and provide enough of a time delay from kinetic impact to NED detonation for the follower spacecraft to enter the crater created by the impactor on the asteroid surface so that the NED fuzing mechanism would not be damaged and allow for a safe subsurface explosion. The follower spacecraft will contain the NED payload

6 Table 4. Bipropellant Fuel Characteristics.

Propellant Characteristics LOX/LH2 N2H4/Hydrazine

Isp (sec) 451 339 Fuel-Oxidizer Ratio 1:6 1:1.34 Oxidizer Density (kg/m3) 1140 1450 Fuel Density (kg/m3) 71 1008

(symbolized by the red cylinder in Figure3), protected by the gray conical shield to ensure that the detonation fuzing mechanism would be safe from the high-temperature plasmas/debris generated by the kinetic impact of the impactor. The HNIS also has a thruster system to help with targeting control, solar arrays on the surface of the follower’s body, and a one meter high-gain antenna for communication purposes during the final 24-hr terminal guidance phase. A detailed discussion of the baseline HNIS design can be found in Reference 6. OTV Design The design of a baseline OTV is needed for three mission cost classifications. If an OTV is deemed necessary for a particular mission, it would be nothing more than another spacecraft bus and motor carrying propellant, to be used for the purposes of orbit injection and/or trajectory correction maneuvers, to be attached to the HNIS bus. According to the desired ∆V , the mass of the fuel and oxidizer will be calculated and tanks will be sized to fit the corresponding amounts of propellant. There are two main types of fuel being considered for the OTV: i) LOX/LH2 and ii) N2H4/Hydrazine. To calculate the amount of propellant needed to provide a certain amount of ∆V , a simplified version of the standard equation is used to find the amount of propellant needed by the OTV, as follows: m0 ∆V = g0Isp ln (1) mf 2 where ∆V is the change in velocity of the spacecraft, g0 is the gravitational acceleration of 9.8 m/s , Isp is the specific impulse of spacecraft’s engine, m0 is the initial mass of the spacecraft, and mf is the final mass of the spacecraft. Given that the necessary change in velocity and final mass of the HNIS/OTV are known, Equation (1) is solved for the final mass. Therefore, the difference between the initial and final spacecraft masses is the propellant mass. Since both fuel options are bipropellant fuel sources, the amount of fuel and oxidizer needs to be calculated so that the appropriate size tanks can be used onboard the OTV. Given the propellant mass needed for the OTV, the Fuel-Oxidizer ratio listed in Table4 for each type of fuel, can be used to find how much of that mass is the fuel and how much is the oxidizer. Using the associated density values for both the fuel and oxidizer, the volume taken up by each propellant component can be determined for the desired change in velocity.

Propellant Tank Sizing

With limited payload mass and fairing space, it is important to find the appropriate number and size of propellant tanks. While those two components should be kept separate from each other, too few or too many tanks could waste valuable space within the or fuel. Each launch vehicle has a finite volume for its payload. The primary piece of payload to be fit into the launch vehicle fairing is the HNIS, while the OTV and its associated fuel tanks would have to be designed based on the remaining space, while harboring all the propellant needed to fulfill its duties. A bulky OTV design that does not fit with the HNIS within the fairing, or has too much mass is as useful as

7 not having an OTV on a mission that mandates one. The design and arrangement of the OTV and its tanks could be crucial to a successful mission. Given the amount of propellant needed, the process of finding the correct size tanks may warrant several iterations of the design-loop before a solution is finally found. The key variables include the desired ∆V, the mass of the propellant, the size of the tanks, and the number of tanks, that all need to be kept track of in that stage of the design process. There exists a fixed relationship between the ∆V and mass, shown by Equation (1), and between the number of tanks to hold the propellant and the size of those tanks, so the initial design variables to be changed, if necessary, would be the number and size of the propellant tanks. If the size and number of tanks do not fit within the fairing of the specified launch vehicle, the solution set would need to be revised, either by changing the number of tanks in the OTV design, or as a last alternative, pick a different launch vehicle or target NEO, in order to come up with a suitable solution.

Launch Vehicles The preliminary mission design studies conducted in this paper considered three classes of launch vehicle: i) Delta II, ii) Delta IV, and iii) V. Due to the payload capacity and launch costs, the 300-kg NED mission will exclusively look at Delta II class launch vehicles, the 1000-kg NED mission could be handled by any launch vehicle from the Delta IV and classes, and the 1500-kg NED mission will likely need a Delta IV Heavy launch vehicle due to the large amount of mass designated to the NED special payload. The launch vehicles will be carrying specially designed HNIS, comprised of an impactor and follower spacecraft, with the NED payload contained within the follower spacecraft. An OTV option is considered, that would accompany the HNIS on its mission, to provide extra ∆V for trajectory correction maneuvers (TCMs) or orbital insertion, as needed. Delta II Launch Vehicles The Delta II launch vehicles have a 98% reliability record, with ca- pabilities to launch from either the East or West coast. The vehicles can be configured with two or three stages with up to nine strap-on graphite-epoxy motors, and two sizes of payload fairings. The versatility and the low cost of the Delta II launch vehicles makes it ideal for the requirements of a $500M mission.8 The types of Delta II launch vehicles considered are the Delta II 732X, Delta II 742X, Delta II 792X, and Delta II 792XH, where X can be 0 (no third stage), 5 (-48B third stage), or 6 (STAR-37FM third stage). The major differences between the two and three stage configurations of the Delta II are the payload and orbit injection capabilities. Two-stage Delta II can really only take payloads into low Earth orbits (LEOs), while the three-stage configurations have the ability to inject payloads into hyperbolic C3 orbits. With payload mass being such an important variable for the Delta II class mission, a trade study was conducted between the two-stage configuration Delta II to LEO plus OTV and the three-stage configuration Delta II to the desired C3 energy orbit, to find which configuration would be more accommodating for the HNIS/OTV design. As mentioned, the three-stage configuration of the Delta II launch vehicle is capable of placing its payload into a C3 orbit - an Earth escape trajectory. The C3 value is defined by the energy of the orbit that the payload is placed in, expressed as 2 C3 = v∞ (2) where v∞ is the hyperbolic excess speed of the spacecraft. As shown in Table2, the escape speed would have to be about 3.46 km/s in order to arrive at 2002 AT4 with a relative speed of 9 km/s.

8

Figure 4. Delta II Launch Vehicle Configurations.8

Figure 5. C3 Plots of Three-Stage Delta II Capabilities.

Based on Equation (2), the required C3 orbit for the Delta II rocket would have to be about 11.94 km2/s2. Using data taken from the Delta II Payload Planner’s Guide,8 the following diagram was constructed to compare the C3 and payload mass capabilities of the three-stage Delta II configura- tions. The green vertical line in this figure denotes the required C3 value to go from the designated parking orbit into the desired hyperbolic escape trajectory necessary to impact asteroid 2002 AT4. The individual colored curves are polynomial fits of points extracted from the Delta II Payload Planner’s Guide curves, in order to have analytical expressions for C3 in terms of payload mass and vice versa. The intersection between the green line and a colored curve gives the maximum allowable payload mass that can be injected into the desired C3 orbit using that specific launch

9 Table 5. Payload Capabilities of Two-Stage Delta II Launch Vehicles to 185-km LEO. Delta II Launch Vehicles Payload Mass (kg) 7320 2809 7420 3195 7920 5030 7920H 6097

Table 6. OTV Characteristics for Delta II based mission. 300 kg NED mission 200 kg NED mission HNIS mass 1843 kg 1843 kg OTV structure mass 300 kg 300 kg Desired ∆V 3.6 km/s 3.6 km/s Fuel type N2O4/Hydrazine N2O4/Hydrazine Total Fuel mass 1787.75 kg 1704.33 kg Total Oxidizer mass 2395.59 kg 2283.80 kg Number of Fuel tanks 2 2 Number of Oxidizer tanks 2 2 Fuel tank dimensions 1.2446 m x 1.1582 m 1.9050 m x 0.9020 m Oxidizer tank dimensions 1.9050 m x 0.9020 m 1.9050 m x 0.9020 m Fuel tank volume 1.6695 m3 1.5499 m Oxidizer tank volume 1.5499 m3 1.5499 m Total OTV mass 4483.34 kg 4288.13 kg Total Delta II Payload mass 6326.34 kg 6031.13 kg vehicle. A baseline HNIS, to be launched onboard a Delta II rocket, has a total wet mass of 1843 kg, far too much mass to be carried into the desired C3 orbit by any individual three-stage Delta II launch vehicle. Indicating that either the mass of the HNIS would have to be trimmed down, or an alternative method to reach the necessary orbit must be found. Since the two-stage configuration of the Delta II class launch vehicles can take payloads of vary- ing size to LEO, shown in Table5, an orbital transfer vehicle is needed in this case to provide the necessary push to a hyperbolic escape trajectory. The design for the OTV would be rather simplistic, a structure containing a number of propellent tanks and an engine capable of performing the desired orbital insertion, using the systems from the HNIS bus to control the performance of the OTV. As- suming an OTV structure mass of 300 kg and payload mass of 1843 kg for the HNIS, Equation (1) can be used to find the propellant mass needed to obtain a v∞ of 3.6 km/s. The resulting OTV design is summarized in Table6. Based on the total Delta II payload mass of 6326.34 kg and the payload capabilities of the Delta II class launch vehicles presented in Table5, the Delta II 7920H launch vehicle cannot carry the baseline HNIS/OTV to the specified parking orbit. With those results, a decision needs to be made of whether to put the HNIS on a larger launch vehicle, or reduce the seize of the payload within the Delta II launch vehicle fairing. The HNIS mass contains a 30% mass margin, so the total payload mass is taken to be a maximum. So, the mass of the NED is reduced to 200 kg and the design algorithm is run again. The results come up with a payload mass of 6031.13 kg, within the Delta II 7920H’s capacity to take into LEO. Delta IV and Atlas V Launch Vehicles The Delta IV class of launch vehicles are much larger launch vehicles, capable of not only taking large payloads to LEO but directly injecting them into higher C3 orbits. The variety within the Delta IV class allows for a rocket from this group to be

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Figure 6. Delta IV (left) and Atlas V (right) Launch Vehicles.9,10

Figure 7. Payload Mass vs. C3 Plots for Delta IV and Atlas V Launch Vehicles. picked either the 1000-kg NED mission or the 1500-kg NED mission. The Atlas V launch vehicles are also rather powerful rockets, with comparable if not better payload capabilities to that of their equivalent Delta IV rocket counterparts, with the exception of the Delta IV Heavy launch vehicle. Due to their versatility and lack of a heavy launch vehicle, the Atlas V launch vehicles are only being considered for the 1000-kg NED mission. The currently available launch vehicles for use for either of those missions are: Delta IV Medium, Delta IV M+(4,2), Delta IV M+(5,4), Delta IV Heavy, Atlas V 401, Atlas V 431, and Atlas V 551. Most of these launch vehicles are too powerful to simply inject the payload into LEO, so instead the payloads will be placed directly into a C3 orbit of 11.94 km2/s2.

11 Table 7. Payload Capabilities of Delta IV and Atlas V Launch Vehicles for C3 = 11.94 km2/s2 Launch Vehicle Payload Mass (kg) Delta IV Medium 2586.13 Delta IV M+(4,2) 3679.75 Delta IV M+(5,4) 4377.67 Delta IV Heavy 8637.07 Atlas V 401 2683.41 Atlas V 431 4453.53 Atlas V 551 5239.31

Table 8. List of Potential Experimental Target NEOs.4 Target NEO Estimated Diameter (m) Departure ∆V (km/s) Departure Date C3 (km2/s2) 2003 GA 300 3.519 12/03/2015 6.50 2006 SJ198 1200 4.595 03/17/2015 32.01 2009 TB3 300 3.600 01/27/2018 8.35 2007 FS35 620 3.473 02/04/2015 5.46 2003 QC 400 4.479 01/01/2015 29.14 2004 GY 480 4.354 06/30/2015 26.09 2001 SX269 280 3.572 05/02/2019 7.70 1998 SB15 330 3.335 05/05/2017 2.37 2004 KE1 240 4.539 02/08/2017 30.62 2011 BX10 1000 3.948 01/01/2015 16.39

The allowable payload mass to the required C3 orbit for each launch vehicle is shown in Table7. Without the need of an OTV for either of these missions, the total payload inside the fairing of the launch vehicles will be that of the HNIS including NED payload. The designs for the HNIS for the 1000-kg and 1500-kg NED missions have a total wet mass of 4251 kg and 5720 kg, respectively. Thus, for the case of the 1000-kg NED mission, a Delta IV M+(5,4), Atlas V 431, or Atlas V 551 launch vehicle is capable of reaching the desired C3 orbit. Taking into account the desire to have more space in the fairing, to allow for more freedom in the HNIS design, and available payload capacity, the Atlas V 551 launch vehicle would be used for the 1000-kg NED mission. For the 1500-kg mission the Delta IV Heavy launch vehicle is the only choice given its tremendous payload capabilities. The wet masses of the two HNIS designs are not completely finalized, but the choice of launch vehicles for each of the missions also takes into account a mass margin for the HNIS, in case the nominal mass of the spacecraft or NED needs to be changed.

POTENTIAL NEO TARGETS The class of asteroids analyzed as potential targets in this study are the Amor asteroids, whose orbits do not cross the path of the Earth and have no possibility of impacting the planet. From the over 3000 Amor asteroids, 10 of the most suitable asteroids were identified in Reference 4 as appropriate targets for at least one of the three mission classes, as stated in the premise of the study, shown in Table8. From the list of 10 candidate target NEOs, primary and secondary targets are picked for each of the three mission types. In order to pick the best targets for each mission, the asteroids were grouped according to their estimated diameter and C3 value. In regards to their estimated diameters, the 10 asteroids seem to fall into three distinct groups: i) below 400 meters, ii) between 400 meters and 1000 meters, and iii) above 1000 meters. These groupings were made based on the relative size

12 of the NEDs, assuming that the 300-kg NED would be used on a mission to a smaller asteroid, the 1000-kg NED used on a mid-sized asteroid, and the 1500-kg NED used on a larger asteroid. Similarly, the asteroids fell into three separate C3 energy groups: i) below 10 km2/s2, ii) between 10 km2/s2 and 30 km2/s2, and iii) above 30 km2/s2. The C3 energy groupings were formed in relation to the capabilities of the launch vehicles assumed to be capable of attaining the orbit with the corresponding HNIS. Delta II, mid-sized Delta IV or Atlas V, and Delta IV Heavy would be capable of handling the respective C3 energy group. With the asteroids grouped according to both evaluation criteria, targets began to be selected and eliminated. However, a third evaluation criteria was used to help distinguish primary and secondary targets, launch date. A primary target NEO could not have a launch date after that of the secondary target. Unfortunately, there was not always an ideal choice for each mission class, so comprises on either size or energy had to be made. Using this selection methodology, the primary and secondary selections for each mission have been made as follows: i) for the 300-kg NED mission: 1998 SB15 and 2009 TB3 ii) for the 1000-kg NED mission: 2003 QC and 2004 GY iii) for the 1500-kg NED mission: 2011 BX10 and 2006 SJ198 With these selections and the previously discussed HNIS design, the design flowchart shown in Figure2 was applied to each primary asteroid and mission pair to find the appropriate launch set up and mission design needed to reach each target.

MISSION TO 1998 SB15

From the data provided in Table8, we can see that asteroid 1998 SB15 has an estimated diameter of 330 meters, a required C3 value of 2.37 km2/s2 and an associated launch date on May 5, 2017 in order for the HNIS to arrive with a relative speed of approximately 10 km/s, a perfect target for the inexpensive mission class. Figure8 shows the mission trajectory of the HNIS and the orbit tracks of the Earth and the asteroid over the transfer time. The mass of the HNIS configuration to be used for the 300-kg NED mission to 1998 SB15 becomes 1843 kg, with NED payload mass and the added mass margin included. From Figure5, the three-stage configuration Delta II launch vehicles are incapable of taking such a massive payload to a C3 orbit of 2.37 km2/s2. Therefore, a two-stage Delta II launch vehicle with additional OTV needs to be utilized in order to facilitate the mission requirements. The results from the launch vehicle and OTV design algorithm are shown in Table9. From the given C3 energy, the associated departure ∆V is approximately 3.33 km/s (Table8). Therefore, airing on the side of caution and carrying some extra fuel, the OTV would carry 3.35 km/s worth of fuel in order to inject the HNIS/OTV configuration into the appropriate C3 orbit to impact 1998 SB15. The resulting trajectory has a transfer time of 159 days, having the HNIS and OTV arriving at the asteroid on October 11, 2017. Figure9 shows a preliminary design of the HNIS/OTV configuration, within the fairing of a Delta II rocket, that would be launched to impact asteroid 1998 SB15. The drawing on the far left shows the OTV sitting in the fairing of a Delta II launch vehicle, in order to provide a reference for its size. The drawings in the middle and right show two different views of the HNIS and OTV mounted inside the Delta II fairing, to show how they would fit together and inside the launch vehicle fairing.

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Figure 8. Mission Trajectory from Earth to 1998 SB15.

Table 9. Mission Design Results for 1998 SB15. Launch Vehicle Total Payload Mass (kg) Departure ∆V (km/s) Departure Date Arrival Date Delta II 7920H 5868.2 3.35 05/05/2017 10/11/2017

Figure 9. A Baseline HNIS and OTV Configuration within a Delta II Two-Stage Fairing

MISSION TO 2003 QC

Asteroid 2003 QC has an estimated diameter of 400 meters, a required C3 of 29.14 km2/s2, and a launch date on January 1, 2015, an appropriate target for the 1000-kg NED mission. Based on the C3 energy requirement, it can be assumed that the amount of energy needed to be provided to the

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Figure 10. Mission Trajectory from Earth to 2003 QC.

Table 10. Mission Design Results for 2003 QC. Launch Vehicle Payload Capacity (kg) Delta IV Medium 1691.52 Atlas V 401 1844.34 Delta IV M+(4,2) 2561.74 Delta IV M+(5,4) 3093.23 Atlas V 431 3215.71 Atlas V 551 3775.26 Delta IV Heavy 6427.22

HNIS far exceeds that which can be provided by a Delta II rocket, in either a three stage or two stage plus OTV configuration. Regardless, both Delta II configurations were tested to see if the asteroid could be reached. With an HNIS mass of 4251 kg, with NED and mass margin included, no three- stage Delta II configuration launch vehicle could carry such a massive payload to the required C3 orbit. The required HNIS/OTV mass to be placed in LEO to reach the target NEO would have to be around 7435 kg, over 1400 kg beyond the maximum carrying capacity of the Delta II 7920H launch vehicle. Therefore, a more powerful rocket must be used to achieve the desired orbital injection and transfer requirements. Figure 10 shows the mission trajectory of the HNIS from the Earth to 2003 QC, and the orbit tracks of both the Earth and asteroid over the transfer time of the spacecraft. For this target asteroid, the HNIS/OTV design algorithm did not consider the Delta II class launch vehicles, since they have been ruled out for this particular mission. Also, since the launch vehicles within the Delta IV and Atlas V classes can support the size of payload to the desired C3 orbit, an OTV is not needed in the design of this mission.

15 The payload capacity for the various launch vehicles to a C3 orbit of 29.14 km2/s2 are shown in Table 10. The only launch vehicle capable of placing the 4251-kg HNIS payload into the desired C3 orbit to impact 2003 QC is the Delta IV Heavy. Now, given that the mass of the HNIS has a mass margin added to it, in the case that the mass decreases after construction and fabrication, a smaller launch vehicle like the Atlas V 551 could potentially be used to perform the same mission. However, given the data from Table3, the Delta IV Heavy is the launch vehicle of choice for the mission to 2003 QC.

MISSION TO 2011 BX10 For the most expensive and complex mission, with the largest size NED and HNIS mass, it was decided that the two best asteroids to target were 2011 BX10 and 2006 SJ198. While specific details on the asteroids can be seen in Table8, the is taken here to highlight a few key characteristics. Both asteroids have relatively large estimated diameters, of 1000 and 1200 meters respectively, and they both have high C3 energies - much more than any configuration of Delta II launch vehicle can handle. The reason that asteroid 2011 BX10 was chosen as the primary target for this mission over 2006 SJ198, despite its smaller comparative size and C3 energy, is the launch date - January 1, 2015 for 2011 BX10 versus March 17, 2015 for 2006 SJ198. In case the launch date for asteroid 2011 BX10 cannot be met, the designed mission must be easily adapted for asteroid 2006 SJ198 by its launch date. With that being said, Figure 11 shows the mission trajectory of the HNIS from Earth to 2011 BX10, and the time elapsed orbit tracks of both bodies.

Figure 11. Mission Trajectory from Earth to 2011 BX10.

Despite the large cost for this mission, given the large NED being carried to the target, all the launch vehicles within the Delta IV and Atlas V classes are analyzed, on the off chance that a launch vehicle smaller than a Delta IV Heavy can be used to complete the mission. Table 11 shows a comparison of the different payload capacities of the analyzed launch vehicles with respect to C3

16 Table 11. Comparison of Launch Vehicle Capabilities for Asteroids 2011 BX10 and 2006 SJ198. Launch Vehicle Payload Capacity to 2011 BX10 (kg) Payload Capacity to 2006 SJ198 (kg) Delta IV Medium 2337.90 1559.07 Atlas V 401 2449.56 1721.15 Delta IV M+(4,2) 3362.58 2401.87 Delta IV M+(5,4) 4022.02 2902.28 Atlas V 431 4109.43 3033.08 Atlas V 551 4816.06 3572.58 Delta IV Heavy 8028.09 6095.76 energies of 16.39 km2/s2 (2011 BX10) and 32.01 km2/s2 (2006 SJ198). The mass of the HNIS, with NED and mass margin included, designated for this particular mission scenario is 5720 kg. And, based on the results displayed in Table 11, the only launch vehicle capable of carrying this payload to such high C3 energy orbits is the Delta IV Heavy.

IMPACT/INTERCEPT APPROACH ANGLE With the launch vehicles, payload configurations, and mission orbits determined, the impact ap- proach angle for each mission is discussed here. The configuration of the HNIS, which was dis- cussed earlier, is such that the impactor portion of the spacecraft must make contact with the aster- oid surface before any other portion of the spacecraft body. In which case, it is important to know the impact (or intercept) approach angle between the spacecraft and the asteroid near the time of impact. The impact approach angle is defined as the angle between the velocity vectors of the aster- oid and spacecraft at the time of target impact (or intercept). Figure 12 shows the impact approach angles and impact velocities of the Deep Impact mission and the Don Quijote mission.

Target Velocity 29.9 km/s Target Velocity 29.83 km/s

15 deg 11.4 deg 7.16 km/s 10 km/s 24.4 km/s Impact Velocity 25.18 km/s Impact Velocity Spacecraft Velocity Spacecraft Velocity

(a) Deep Impact Mission (b) Don Quijote Mission Figure 12. Velocity Vector Diagrams and Impact Approach Angles of the Deep Im- pact and Don Quijote Missions.

In most cases, the speed of the asteroid is greater than that of the approaching spacecraft. In each of the three previously discussed missions, the HNIS would be impacted by the asteroid. In all three cases, the asteroid would be impacting the spacecraft from behind, meaning the velocity vectors of both bodies would be mostly along the same direction with the asteroid’s speed being greater than the spacecraft’s speed. For asteroids 1998 SB15 and 2003 QC, the impact approach angles are around 20 and 16 degrees respectively, while for asteroid 2011 BX10 the angle is about 6 degrees, as illustrated in Figure 13. Prior to a terminal-phase operation that guides the spacecraft into a precision impact at a certain speed and/or angle, the transfer orbit from the Earth to a target asteroid will place

17 Target Velocity 29.52 km/s Target Velocity 29.89 km/s

19.93 deg 15.97 deg 9.96 km/s 9.96 km/s Impact Velocity Impact Velocity 23.12 km/s 27.63 km/s Spacecraft Velocity Spacecraft Velocity

(a) 1998 SB15 Mission (b) 2003 QC Mission

Target Velocity 37.67 km/s

5.88 deg 10.74 km/s 27.45 km/s Impact Velocity Spacecraft Velocity

(c) 2011 BX10 Mission

Figure 13. Velocity Vector Diagrams and Impact Approach Angles of the three HNIS missions. the HNIS in a proper relative position for the asteroid to run into it. Should the need to impact any of these asteroids from a different direction or with a different speed, a terminal-phase guidance system would take trajectory data - relative position and velocity, and use a combination of the spacecraft’s main engines and divert and thrusters to reposition and/or orient the HNIS to impact the asteroid in the desired manner. Detailed discussions of such an active terminal guidance system design for asteroid intercept missions can be found in References 11-12.

MISSION COST ESTIMATION

Mission cost estimation to design and fabricate the missions is an important task necessary for an early assessment of the mission viability and feasibility. The final total cost of each mission is given as a combination of the cost for the launch vehicle, the HNIS/OTV system, and any fuel for the OTV, if utilized. Initially, the maximum costs for the three mission scenarios were assumed as: $250M, $500M, and $1B; however, based on the designs of the HNIS/OTV that would be required for the selected target asteroids, those initial cost estimates have been found to be rather modest. A cost estimation algorithm was developed to determine the costs associated with constructing the HNIS, based on a number of previous spacecraft missions with similar goals and parameters. Spacecraft such as Deep Impact, , and were researched to find the cost of develop- ing their spacecraft and a linear polynomial fit was applied to the data to come up with an analytic formula relating spacecraft mass and cost. Before the results of the cost estimation algorithm are discussed, it is important to note that the mass/cost of the NED was not included when the estima- tions were made. In addition, the total mass margin was left intact when estimating the cost of the HNIS development, in order for the estimate to be thought of as a relative maximum.

18 Table 12. Cost Breakdown of Three Baseline PDT Demo Missions. 1998 SB15 Mission 2003 QC Mission 2011 BX10 Mission Launch Vehicle Delta II 7920H Delta IV Heavy Delta IV Heavy HNIS Mass (kg) 1543 3251 4220 Launch Vehicle Cost ($) 100M 324.98M 324.98M HNIS Cost ($) 411.72M 823.86M 1057.68M OTV Cost ($) 2M 0 0 Total Costs $513.72M $1148.84M $1382.66M 30% Cost Margins $154M $344M $415M Total Mission Costs $668M $1.5B $1.8B

The total mass of these spacecraft for the three different demo missions, without NED payloads, are 1543 kg, 3251 kg, and 4220 kg, respectively. Running these masses through the cost estima- tion algorithm gives spacecraft development costs of approximately $411M, $823M, and $1057M, respectively. Again, these values are only estimates of the cost to fabricate the individual space- craft. Table 12 shows a cost breakdown for each NED mission, along with a total mission cost. The cost of each mission is limited to: i) the launch cost of the specified launch vehicle, ii) the HNIS fabrication, and iii) the OTV fabrication and fuel. A similar cost analysis was also run using NASA’s Advanced Mission Cost Model (AMCM),13 to get a rough-order-of-magnitude approximation for the costs of these three missions. The estimates from the AMCM for each HNIS came out to be $616M, $979M, and $1149M, respectively, in 2004 US dollars. These estimates are really rough, mostly due to the fact that these HNIS designs don’t exactly fit into a single mission category from the available choices. However, the estimates at least verify that the estimates shown in Table 12 are in the appropriate cost range. Given the total cost estimates discussed above and taking the results from the AMCM into consideration, the revised mission costs would be approximately $668M, $1.5B, and $1.8B, respectively, accounting for mission operations costs by adding 30% of the estimated total costs, as shown in Table 12. More detailed discussions on cost estimates as well as technical assessments of a variety of NEO deflection/disruption missions can be found in References 14 and 15.

CONCLUSION In this paper, the preliminary designs and analyses of flight demonstration missions have been presented for spacecraft systems carrying nuclear payloads on an interplanetary trajectory from Earth to a target asteroid. The threat of asteroids impacting the Earth is very real and must be taken seriously. While the threat of the Earth being struck by an asteroid of substantial size to warrant the use of such spacecraft and mission designs is relatively small, the day may arise when the Earth is in danger and life on this planet may be in jeopardy. Instead of hoping for a successful mission when the time comes or assuming that day will never come, flight demonstrations missions such as those described in this paper on actual non-Earth-threatening asteroids can provide knowledge and experience that may prove vital, should they be needed in the future.

ACKNOWLEDGMENT This study was supported in part by the NIAC (NASA Innovative Advanced Concept) program of the NASA Office of the Chief Technologist and by the Iowa Space Grant Consortium. The authors would like to acknowledge the help provided by Timothy Winkler at the Iowa State ADRC for the

19 target NEO selection data he has provided as well as the discussions which led to the refinement of this paper.

REFERENCES [1] B. Kaplinger, B. Wie, and D. Dearborn, “Near-Earth Object Fragmentation and Dispersion Modeling and Simulation,” presented at 2011 IAA Planetary Defense Conference, Bucharest, Ro- mania, May 2011 (to appear in Acta Astronautica). [2] B. Wie, “Hypervelocity Nuclear Interceptors for Asteroid Deflection or Disruption,” presented at 2011 IAA Planetary Defense Conference, Bucharest, Romania, May 2011 (to appear in Acta Astronautica). [3] B. Kaplinger and B. Wie, “Comparison of Fragmentation/Dispersion Models for Asteroid Nu- clear Disruption Mission Design,” AAS 11-403, AAS/AIAA Astrodynamics Specialist Conference, Girdwood, AK, August 2011. [4] T. Winkler, S. Wagner, and B. Wie, “Optimal Target Selection for Planetary Defense Technol- ogy (PDT) Demonstration Mission,” AAS 12-226, AAS/AIAA Space Flight Mechanics Meeting, Jan. 30 - Feb. 1, 2012. [5] I. Carnelli, A. Galvez, F. Ongaro. “Learning to Deflect Near Earth Objects: Industrial Design of the Don Quijote Mission,” 57th International Astronautical Congress, 2006. [6] A. Pitz, B. Kaplinger, B. Wie, and D. Dearborn “A Hypervelocity Nuclear Interceptor System (HNIS) for Optimal Disruption of Near-Earth Objects,” AAS 12-225, AAS/AIAA Space Flight Mechanics Meeting, Jan. 30 - Feb. 1, 2012. [7] S. Wagner, A. Pitz, D. Zimmerman, and B. Wie, “Interplanetary Ballistic Missile (IPBM) Sys- tem Architecture Concept for Near-Earth Object Threat Mitigation,” presented at 60th International Astronautical Congress, Daejeon, Korea, October 2009. [8] “Delta II Payload Planner’s Guide,” , Littleton, CO, 2006. [9] “Delta IV Payload Planner’s Guide,” United Launch Alliance, Littleton, CO, 2007. [10] “ Atlas Launch System Mission Planner’s Guide,” Commercial Launch Services, Denver, CO, 2004. [11] M. Hawkins, A. Pitz, B. Wie, and J. Gil-Fernandez, “Terminal-Phase Guidance and Control Analysis of Asteroid Interceptors,” AIAA 2010-8348. [12] M. Hawkins and B. Wie, “Impact-Angle Control of Asteroid Interceptors/Penetrators,” AAS 11-271. [13] “Advanced Missions Cost Model,” NASA JSC Cost Estimating and Models Web Site. 2007. cost.jsc..gov/AMCM.html [14] “Near-Earth Object Survey and Deflection Study Report,” NASA, 2006. [15] “Defending Planet Earth: Near-Earth Object Surveys and Hazard Mitigation Strategies,” Report No. 0-309-14968-1, National Research Council, National Academy of Sciences, 2010.

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