Multiple Asteroid Retrieval Mission from Lunar Orbital Platform-Gateway Using Reusable Spacecrafts

Gustavo Gargioni David Alexandre Department of Aerospace and Ocean Engineering Department of Aerospace and Ocean Engineering Virginia Tech Virginia Tech Blacksburg, VA 24060 Blacksburg, VA 24060 [email protected] [email protected] Marco Peterson Kevin Schroeder, Ph.D. Department of Aerospace and Ocean Engineering Department of Aerospace and Ocean Engineering Virginia Tech Virginia Tech Blacksburg, VA, 24073 Blacksburg, VA, 24073 [email protected] [email protected]

Abstract—This paper describes the results of a study to ﬁnd pos- 1. INTRODUCTIONAND MOTIVATION sible Near Earth Asteroids (NEA) capable of being captured using upcoming rocketry for the purposes of space-based mining, According to the Federal Aviation Administration, only USD combining reusable rockets such as SpaceX’s Big Falcon Rocket 1.1 billion of the USD 344.5 billions of dollars were spent in (i.e. BFR) and refueling capabilities. This work introduces a space as private investment, excluding satellite and govern- relatively low-cost option with higher Delta V, ∆V , compared ment contracts [1, Fig. 1]. In another words, very little effort to non-reusable rockets, and an opportunity for NASA’s Lunar has been made into creating long-term business in space. If Orbital Platform-Gateway (LOP-G) for synergy with services, the goal is to become a spacefaring civilization, an economy science, and technologies. In an effort to maximize the number in space with investment from the private sector is essential. of viable missions, the study focused on choosing a refueling In Wall Street, Morgan Stanley predicts the space industry orbit near LOP-G; thus, the Earth-Moon Lagrange points L1 and L2 were selected as possible choices for this paper. The will have tripled in size to USD 1.1 trillion by 2040. All resulting simulations of a Cislunar infrastructure orbiting in this comes in place due to expectations that, with the latest Lagrange points highlight differences in orbital options. Indeed, advancements of reusability and new space launch providers, the optimal option is balanced between a Near Rectilinear Halo space will be within reach for investors from many other Orbit (NRHO) in L1 and a NRHO in L2. The decision for industries to expand their base of operations. optimal positioning of any Gateway Station is dependent on the type of mission allocated to the Cislunar station. However, both options seem promising, not only for asteroid extraction and mining but also for crewed and cargo missions. In a worst case scenario, an operation of 3 decades starting in 2030 retrieving 33 asteroids, or 1.1 per year, comprising up to 2,581 tons in minerals. The work culminated in developing a data mining online tool that searches the entire Near Earth Asteroids (NEA) close approach database from JPL and the small body database from NASA. Using a ﬂeet of 10 BFRs over 30 years may provide a higher long-term success for a business case than any other investment on Earth.

TABLEOF CONTENTS

1. INTRODUCTIONAND MOTIVATION ...... 1 2. LAGRANGE POINTSIN CISLUNARSYSTEM ...... 2 3. ORBITAL OPTIONS ...... 3 4. LONGTERM ORBITAL ANALYSIS ...... 4 5. HARDWARE ...... 4 6. REFUELINGAND SERVICING ...... 6 7. METHODOLOGYAND MARM TOOL OVERVIEW ...7 8. METHODOF ASTEROID SELECTION ...... 8 Figure 1. The global space economy in context 9. A CASEOF STUDY ...... 10 10. CONCLUSIONS ...... 11 The United Nations Environment Program (UNEP) international panel declared that, in 2050, human beings would APPENDICES ...... 12 consume 140 billion tons of minerals, ores, fossil fuels and A.MARMCANDIDATES ...... 12 biomass per year – three times today’s standard. Projections for human population increase are of 50 percent in the next 3 REFERENCES ...... 13 decades. The world concurs that this unbalanced symbiosis is BIOGRAPHY ...... 13 the result of our way of life and scale, and that action must be 1 taken. The most promising answer may very well lie in space This spacecraft would then be reused in space for multiple exploration. asteroid retrieval missions. The reusability of the spacecraft for multiple missions will then reduce the total investment In fact, it is not only in the U.S. that we find evidence of necessary for a multiple asteroid retrieval mission (MARM). this tendency. Asteroids are being probed for composition elsewhere as well. According to astronomers at the Lincoln Assuming LOP-G will be the base of operations and how Near-Earth Asteroid Research Lincoln Lab, the ”162173 the refueling will operate, initial conditions will substantially Ryugu” asteroid is speculated to be worth USD 82 billions influence the amount of ARM available for a MARM project. in minerals, composed mostly of nickel, iron, cobalt, water, In the following sections, this paper will present an analysis nitrogen, hydrogen and ammonia. Several space organi- of these parameters: zations have planned missions to ascertain both worth and composition of these profitable gold mines. Hayabusa 2, the 1. Location for LOP-G. JAXA mission, successfully arrived at the asteroid on 27 June 2. Hardware. 2018, with a planned material return to Earth by the end of 3. Fuel logistics and costs. 2020. Depending on the valuation of Ryugu, we may see a private worldwide race towards asteroid mining. With the definition of these parameters, the method is applied selecting suitable ARMs within the time span and thus, creat- Today, the top 5 mining companies are, in revenue order, ing a MARM project. The outline is built on a tripod of recent Glencore, BHP Billiton, Rio Tinto SA, China Shenhua En- developments: (1) current update from close approaches from ergy and Vale SA. In 2015, this group generate USD 396.9 Center for Near Earth Object Studies (CNEOS); (2) Spacex’s billions in revenue. The question is: what percentage and Big Falcon Rocket and Spaceship; and (3) NASA’s Lunar amount of this revenue is invested in prospecting new sites? Orbital Platform-Gateway (LOP-G). These elements are used Analyzing Rio Tinto’s recent Annual Reports, we find that to perform feasibility analysis of multiple asteroid retrieval the investment on surveying new sites for minerals hit USD missions (MARM) and the return of NEOs for the purposes 0.942 billion in the last couple of years, an average of USD of space based refining and construction. Moreover, the tool 0.471 billion per year. Another piece of information is that in can be used dynamically to predict different MARM projects the last half a decade, the average profit from this company based on different parameters from upcoming updates from alone was USD 3.897 billion a year [2]. If we extrapolate SpaceX, NASA, JPL and the science community. the revenue percentage from this company’s data to a group perspective revenue percentage we get USD 44.5 billion in profit each decade, and USD 5.38 billions spent on surveying 2. LAGRANGE POINTSIN CISLUNARSYSTEM the Earth for new sites to generate minerals for the Earth’s demands. If this group enters space mining surveying with 0.5% of their profit and shifts 5% of their current budget from surveying the Earth to space surveying, a total of USD 14.56 billion would be available for a 30-year project, or around USD 500 millions per year of operation. That would represent an increase of 50% in the current private sector space industry. With the advancement of rocket reusability, new operational possibilities for fuel logistics have arisen. Shifting focus from optimizing rocketry for one launch to designing the hardware to include multiple launches will reduce the cost per launch substantially. It is hard to believe that the cost per launch may drop from millions to thousands of dollars. However, in the airline business the cost of fuel is 40% of the total cost of each flight. Then, analogously, a Falcon 9 launch would cost around USD 500,000, because current fuel cost is estimated at USD 200,000 per flight [3]. This percentage is due to the fact Figure 2. Lagrange points in the Earth-Moon system that airplanes are reused many times during their life span. Since 2015, booster landing has become a reality. If we get The three-body problem the launch manifest from SpaceX for 2018, more than half of their launches are reused boosters and as they own almost In 1722, Leonard Euler proved the existence of the collinear a quarter of the Earth launches’ market share, rocket launch libration point. In 1765, Lagrange found the triangular reusability is estimated to be 50% in 2018 for SpaceX. Should libration point L4 and L5 [4, Fig. 2]. In 1899, Henri Poincare´ the market trend continue, rocket launch reusability will be proved that the restricted problem was unsolvable and ana- present in half of Earth launches in 2021. lytic, differentiable function of both the initial conditions and the time. Solving this problem is infinitely complex because Overview and problem definition of its many nonlinear dynamical systems which still have no closed form solution [5]. With operations to start in 2030, this paper considers reusability to be ubiquitous and, thus, launch costs will make space This problem is defined as two objects with mass [5], called exploration very accessible. Taking advantage of this trend, primaries, which are executing simple circular orbits around this paper proposes to optimize escape velocity to the mini- their composed center of mass. And the third body assumes mum Delta V (∆V ), maximizing the amount of Asteroid Re- the form of infinitesimal relative mass, compared to the trieval Missions (ARM). That will require multiples launches primaries. The problem of motion of the third body is called to transfer the fuel needed to a much higher orbit, where the the circular, restricted, three-body problem, or the CR3BP. mission spacecraft will be refilled, launched, and returned to. However, when we restrict the motion of the third body so 2 that it is in the same orbital plane as the other two bodies, the Mission scenario using STK modeling software problem is called the planner circular restricted three-body A computational simulation was performed by using Satellite problem, or PCR3BP. Tool Kit (STK). The mission scenario includes a launch from Stability problem and stationkeeping the Kennedy Space Center (KSC), a translunar injection from Low Earth Orbit (400 km altitude) and, when a Lagrange Poincare´ proved that this restricted problem did not have a point is reached, sequential maneuvers to perform the sta- closed solution; thus, some techniques should be applied to tionkeeping in NRHO during one sidereal month (27.3 days). the restricted problem, such as periodic solutions. One may Because the stability and controllability of halo orbits near study this periodic solutions using the Floquet theory. There- Lagrange points are different, the purpose of this simulation is fore, a third body - e.g. hub space station - could be expected to discuss the feasibility of keeping the LOP-G station near a to be relatively stable in the vicinity of a Lagrange point Lagrange point and review both scenarios (L1 and L2) against within a periodic orbit. Moreover, if the point is unstable, several criteria to show which solution is more optimum, or or not perfectly stable - like L1 and L2 -, a stationkeeping profitable, for the LOP-G station. ∆V could be applied to the system to sustain the small disturbances even if other perturbations occurred. A ”halo orbit” is one of these periodic orbits. This orbit is largely studied and a candidate for a hub station. One great example of success of the use of a halo orbit within an L1 Lagrangian point is Solar Heliospheric Observatory (SOHO) which was placed successfully at orbit and has been operating with a stationkeeping ∆V since 1996 [6].

In the next section, this paper explores L1 and L2 options for LOP-G with the use of stationkeeping ∆V . This will serve as a hub for services and refueling operations for each ARM.

3. ORBITAL OPTIONS As presented in the overview, the starting point for each ARM will be the location of LOP-G. In this section, the paper investigates the optimal orbit for LOP-G. Two types of orbits are considered in this paper, namely:

1. Earth-Moon L1 Near Rectilinear halo Orbits (L1-NRHO) 2. Earth-Moon L2 Near Rectilinear halo Orbits (L2-NRHO)

Figure 4. Simulation of a launch from KSC and a translunar injection from LEO (400 km) to the Lagrange point L1 and a stationkeeping in NRHO during one sidereal month in the Earth inertial reference (top) and the Moon inertial reference (bottom) using STK

∆V rkm{ss L1-NRHO L2-NRHO

Translunar Injection 3.79 4.13

Stationkeeping 0.01866 0.00263 Table 1. ∆V required for L1 and L2 mission including 1 translunar injection and stationkeeping for 1 sidereal month Figure 3. L1-NRHO (left) and L2-NRHO (right) performed in the Earth-Moon 2-body system From the results obtained from the simulation (see Fig. 4 and Table 1), the transfer from LEO to the Lagrange point L1 The two potential locations are based on two requirements: 1) is intuitively less expensive in terms of ∆V than the transfer the ∆V required to perform stationkeeping and 2) the Earth’s from LEO (same altitude) to the Lagrange point L2, since it is accessibility of the station in terms of ∆V , launch frequency at a larger distance from Earth. However, the stationkeeping and time of ﬂight, TOF. in NRHO requires more correction in the L1 case than in the L2 case. When a stationkeeping in L2 requires a correction Some additional parameters may also be examined to select a of 2.63 m{s ∆V after one sidereal month, a stationkeeping station location, including: the thermal environment, the line in L1 for the same time period requires almost ten times this of sight with Earth for communication purposes, the attitude correction. This difference is due to the increasing instability control requirements and the eclipse durations of the station. in the Lagrange point L1 area [7]. 3 In a scenario of one single launch for each option from LEO, where: at 400 km, followed by an open-ended stationkeeping in a

Lagrange point, after 21.21 sidereal months („ 580 days) in ‚ Ve1Li is the velocity required to escape the Moon’s gravity NRHO, L2 becomes a better choice over L1, in terms of ∆V . d ˇ For the construction of LOP-G multiple launches it would be 2µ ˇ K ˇ Ve1L1 “ 0.5055 km{s necessary to add different modules, so further calculation is Ve1Li “ , ˇ (2) r Ve1L2 “ 0.4522 km{s needed. Moreover, LOP-G is a dynamic platform, operated Li´K by all space agencies, so transfers might occur for many operations other than the MARM project. Thus, a ∆V ‚ Ve2Li is the velocity required to escape the Earth’s gravity analysis for a long term mission is required to select the d ˇ optimal orbital options. ˇ 2µC ˇ Ve2L “ 1.5178 km{s V “ , 1 (3) e2Li ˇ V “ 1.3579 km{s rLiĆ e2L2 ONGTERM RBITAL NALYSIS 4. L O A ‚ V0 is the orbital speed of the Moon (having escaped the In this section we will further discuss L1 and L2 options Moon’s gravity) but focusing on the MARM project. Acknowledging that a c private consortium would base its operation on a life span µC V0 “ “ 1.0183 km{s (4) between 30 and 50 years, this paper arbitrarily selected a rK minimum thirty year period for this analysis. This is a good $ 3 2 assumption due to the fact that NASA’s current plan for the ’ µC “ 3.9859e5 km {s future of the the International Space Station (ISS) past 2024 ’ 3 2 ’ µK “ 4.9028e3 km {s is to transition responsibility for its operation — in whole or & rL1´K “ 3.8376e4 km in part — to a commercial entity or entities [8]. With: r “ 4.7951e4 km ’ L2´K ’ rL1Ć “ 3.4602e5 km The ISS had an average of 4.45 crewed missions per year ’ % rL2Ć “ 4.3235e5 km in the last 20 years. Considering planning crewed missions rK “ 3.8440e5 km to LOP-G at the same constant rate throughout the period of thirty years, we obtain the results presented in Table 2. VteL1 “ 0.7107 km{s, VteL2 “ 0.5655 km{s (5)

∆V rkm{ss L1-NRHO L2-NRHO Then the ﬁnal Operational ∆V for 30 years will lead to selecting the L2 point for the station as we see on Table 3. LOP-G Orbit Insertion 3.79 4.13

Stationkeeping 7.50 1.06 ∆V rkm{ss L1-NRHO L2-NRHO

Crewed missions (133.5) 505.97 551.36 LOP-G Orbit Insertion 3.79 4.13

Total 517.62 556.55 Stationkeeping 7.50 1.06

Table 2. ∆V comparison in a 30-year analysis with 133.5 Crewed missions (133.5) 505.97 551.36 crewed missions using LOP-G in L1 and L2 Asteroid retrieval missions (90) 63.96 50.90

Interplanetary transfers (60) 42.64 33.93 In this configuration, L1 point would be the leading choice for the LOP-G station. However, the number of crewed missions Total 623.86 641.56 is not the only variable for the study. For a MARM project, interplanetary missions ∆V are inversely proportional in the Table 3. ∆V comparison in a 30-year analysis with 133.5 analysis. To this end, the Earth’s escape velocity required for crewed missions using LOP-G in L1 and L2 each Lagrange point is significant. The following calculations and assumptions are then considered: Analyzing Table 3, we can notice that L2 point becomes a bet- Assumptions: ter option as the configuration of operations becomes mainly interplanetary transfers and asteroid retrieval missions. This ‚ 4.45 crewed missions per year (133.5 in 30 years) result indicates a crucial challenge for NASA to decide where ‚ 3 asteroid retrieval missions per year (90 in 30 years) to operate LOP-G, depending on the administration’s strategy ‚ 2 interplanetary transfers per year (60 in 30 years) in operations. Here we assume a constant rate of missions during the 30 years of analysis but we can expect an increase in the number 5. HARDWARE of companies performing more asteroid retrieval missions year after year as well as more interplanetary transfers. Lunar Orbital Platform-Gateway LOP-G Formerly known as the Deep Space Gateway, currently The velocity required to escape both the Earth’s and the known as the Lunar Orbital Platform-Gateway LOP-G, the Moon’s gravitational influence is given by: conceptual plan remains the same: to enhance human pres- a ence in space by creating an infra-structured hub for fu- 2 2 VteLi “ pVe2Li ´ V0q ` pVe1Li q (1) ture Cislunar operations, Lunar Surface and Missions to 4 Mars. According to [9], 14 space agencies participating in technological breakthroughs, including developing the first NASA’s International Space Exploration Coordination Group private spacecraft to dock with the International Space Sta- (ISECG) have reached consensus regarding the importance tion, and successfully re-flying an orbital class rocket. The of such an infrastructure. Furthermore, like the International reusability of orbital rockets is extremely important to bring Space Station, this gateway will include functions such as the cost of spaceflight down; to re-use Musk’s own example, habitation, logistics resupply, airlock and robotics. commercial aviation would be prohibitively expensive if the plane could only be used once. The reflight of the Falcon 9 LOP-G will not only extend humanity’s footprint in space, but was an important milestone in developing the next level of also become an outpost which will enable many commercial space technology, namely the Big Falcon Rocket, or BFR. and scientific projects. By taking the step further and moving human activity from Low Earth Orbit to a Cislunar space, In September 2016, Elon Musk introduced the ”Interplan- LOP-G may become the Plymouth Rock for private sector in etary Transport System” (ITS) at the 67th annual meeting space. of the International Astronautical Congress (IAC). Musk outlined the initial technical designs of the ITS launch ve- Using the capabilities of Orbital ATK Signus module, Lock- hicle, which would allow for a deeper exploration of space. heed Martin’s Orion capsule and Boeing/NASA new rocket The launch system consisted of three vehicles; a booster, a Space Launch System, NASA’s plan is to start building LOP- spaceship, and a tanker. The booster was the first stage, and G by the end of 2022, when the first launch will put the Power would launch the spaceship and tanker into Earth orbit. The and Propulsion Element (PPE) module in place possibly on spaceship was designed for long-duration spaceflight, and the L1 or L2 Lagrangian points. The final module, an airlock, tanker was designed to refuel the spaceship and other vehicles would be in place no later than 2026. In this 4-year period, once in Earth orbit. some other modules will be incorporated; however, for the scope of this paper, the logistic operations will need a habitat, robotic arm, and airlock, which are all included in the 4-year building process. LOP-G crewed missions consist of 4 astronauts on station for 1 to 3-month periods. In order to maintain a scheduled rotation, LOP-G will be accessible by NASA’s Space Launch System as well as international and commercial ships, such as SpaceX’s Falcon Heavy, ESA’s Ariane 6 and SpaceX’s Big Falcon Spaceship (the spaceship version for the Big Falcon Rocket). The former, SpaceX’s Falcon Heavy, is the only one that is already tested and operational. Space Launch System SLS Figure 6. BFR Tanker/spaceship refueling scheme The Space Launch System, SLS, represents a great increment in the capability in metric tons for Cislunar, thus creating opportunities for mission profiles and a more reliable access Elon Musk unveiled the BFR in September 2017 with a for materials [10]. smaller architecture than that of the ITS, shrinking from a 12-meter diameter to a 9-meter diameter design. SpaceX For each configuration of SLS, the core stage uses four is aiming for an initial Mars launch in the 2022 Mars con- Aerojet Rocketdyne RS-25, used by the Space Shuttle and junction window, well before the 2030 timeline of this paper. capable of producing 1,859 kN of thrust with an Isp of 452 The BFR still contains the three-vehicle design, with separate sec in vacuum and 366 sec at sea level. Block 1B will be able crew/cargo varieties of the spacecraft. One key feature of the to send more than 26 metric tons to Cislunar while Block 2 new design is the back to back refueling, as shown in figure 6. will lift more than 45 metric tons [11]. The idea is to use in-track impulses to help with transferring the fuel from tanker to spacecraft. As the BFR is still in design, the technical specifications are subject to change over the next decade. This paper assumes a BFR spaceship will be fully-refueled in orbit and available for asteroid mining. It will have a cargo volume of 825 m3, capable of carrying at least 200 metric tons. The spaceship will be able to produce 12.7 MN in thrust, providing around 9.1 km{s ∆V empty to 5.6 m{s ∆V full. In 2018’s IAC, Musk unveils more details of the BFR system as SpaceX progresses into development and manufacturing. Total Mass for Rocket and Tanker System before launch is estimated at 4,400,000 kg while the BFS, 2nd stage and Figure 5. SLS Evolution (credit:NASA) spacecraft have 1,335,000 kg total mass, 150,000 kg payload, 85,000 kg empty mass and 1,100,000 kg of propellant. A SpaceX BFR Tanker version of the spacecraft that has no payload is also presented for the refueling process. All stages of the systems SpaceX was founded in 2002 by Elon Musk to revolutionize use the raptor engine, already developed by SpaceX. The space systems and technology, with the end goal of enabling specific impulse for this engine is 330 sec at sea level and Mars colonization. The company has made history for several 375 sec for vacuum. 5 6. REFUELINGAND SERVICING be responsible for lifting fuel to a 400 km altitude orbit for t Space In this section we provide some preview of calculations the first refueling orbit using 1 launches for a fleet of 2 Tankers. It is important to note that the BFR system is fully for a logistic operation between Earth and Cislunar LOP-G reusable, i.e., the first and second stages land back on Earth for the case study of a Multiple Asteroid Retrieval Mission operation. Since the BFR system is still in development, platforms and, thus, they are required to have the extra fuel although advanced and in manufacturing, some assumptions for reentry and landing maneuvers. The study assumes that a 1,000 m{s for the first stage and a 3,000 m{s for the second were made. stage will be enough for those maneuvers. The paper assumes a total mass of 1,640,000 kg for the Tanker Elon Musk’s presentations suggest 4 or 5 fuel resupply Spacecraft version. Although this version lacks in- launches for a full transfer to a BFR spaceship, or BFS. formation, this assumption is made based on the fact that However, since the resuply will be for a BFR tanker of total the volume for 150,000 kg payload would be substituted kg for a bigger tanker with higher density for fuel. Another mass of 1,640,000 , the paper explores this assumption. Considering a total fuel of 1,550,000 kg, a t1 “ 4 launches assumption is that the empty mass for the tanker will be would imply a need for 387,500 kg of fuel per launch transfer, 90,000 kg. while 310,000 kg for t1 “ 5 launches. The former profile NASA LOP-G On-Orbit Servicing generates around 9.5 km{s ∆V while the latter, around 11 km{s (Tables 4 and 5). The mainstay for LOP-G is NASA’s budget. And that is interconnected to the Presidency and thus, society. LOP-G will be more expensive than ISS and the risks for the crew 1st Transfer Total Mass Empty Mass ∆V will be increased as there will be no Van Allen belts to protect people from solar radiation. t1 “ 4 rkgs rkgs rkm{ss Now, consider the possibility of a private company paying LEO 1st Stage 4, 505, 000 2, 200, 000 2.637 for some of these costs. This paper does not assume that Land 1st Stage 2, 200, 000 1, 790, 000 0.759 the private company would invest in the LOP-G construction costs, but rather maintaining and servicing costs. It is rather LEO 2nd Stage 1, 640, 000 550, 000 4.019 logical to think that the experts in on-orbit servicing for these operations would be NASA, and as occurs in airports with Transfer 550, 000 162, 500 N/A airline companies, NASA could benefit from providing these operations. These benefits would not only be in form of Supersonic Retro 162, 500 100, 000 1.435 revenues but also in form of a public/private partnership plus for LOP-G. This is completely aligned with its purposes [9] Final Land 110, 000 90, 000 0.738 and would ensure long-term benefits for public and private Total N/A N/A 9.588 exploration of space. Table 4. Earth Surface - LEO 400 km with t “ 4 Logistics and Refueling method 1 Considering previous result from section 4, L2-NRHO would be the choice for LOP-G location and, thus, MARM service and launch operations. That implies that the MARM Tanker 1st Transfer Total Mass Empty Mass ∆V BFR version must be fully refueled at this orbit in order to maximize the opportunities for retrievals. We can also infer t1 “ 5 rkgs rkgs rkm{ss that by the end of each mission the tank will be fully empty as the worst case scenario. Therefore, a logistic operation for LEO 1st Stage 4, 505, 000 2, 200, 000 2.637 bringing fuel from the Earth’s surface must be implemented to keep the MARM operational. Land 1st Stage 2, 200, 000 1, 790, 000 0.759

Assuming a fleet of 10 BFRs, the paper proposes a 2-step LEO 2nd Stage 1, 640, 000 550, 000 4.019 logistic transfer. The first step is the Earth Surface - LEO 400 Transfer 550, 000 162, 500 N/A km transfer, namely transfer t1, and the second step is the 400 km - L2-NRHO transfer, namely transfer t2. At L2-NRHO, Supersonic Retro 240, 000 130, 000 2.255 docked on LOP-G, the ARM fleet will be waiting the fuel to start each ARM (Fig. 7). Final Land 130, 000 90, 000 1.353 Total N/A N/A 11.023

Table 5. Earth Surface - LEO 400 km with t1 “ 5

Therefore, t1 “ 5 would be enough to fully refuel a single Space Tanker at LEO 400 km. Now that the Space Tanker is fully refueled and orbiting Earth at 400 km, it can then proceed for a lunar injection to the Figure 7. BFR Tanker/spaceship refueling scheme L2-NRHO transfer. The total ∆V for the tanker is around 10,600 m{s and a Hohmann transfer to L2 is 3,190 m{s. From the Earth’s surface, a fleet of 3 Launcher Tankers would Considering that the tanker needs to return for LEO 400 km 6 orbit after delivering the extra fuel, the leftover ∆V is around within NASA’s Framework. On one hand, JPL’s NEO 4,220 m{s. database [13] contains relative motion parameters in reference to Earth as the asteroid makes its close approaches. On Then selecting t2 “ 3, a 517,000 kg of extra fuel should be the other hand, NASA’s Small-Body Database [14] provides transferred in each step while if t2 “ 4, a 387,500 kg transfer the classical orbital elements for each asteroid. Although they is needed. However, we can see that t2 “ 3 is not viable, so both contain a unified object id string, each database has its the study shows that t2 “ 4 would be the choice for this step own separate format for data retrieval. (Tables 6 and 7). This paper uses data from both databases and thus, in order to perform a comprehensive mission calculation for each aster- 2nd Transfer Total Mass Empty Mass ∆V oid, a method for integrating both databases was developed to rearrange data into a usable format. This method is described t2 “ 3 rkgs rkgs rm{ss in the following three steps. to L2-NRHO 1, 640, 000 685, 000 3.212 1. Database Sources Transfer 685, 000 168, 333 5.163 First, concatenation is achieved by establishing an online connection with the near-Earth objects database us- Return to 400 km 168, 333 90, 000 2.303 ing the API CAD available from JPL/NASA website [13]. This service provides an interface to machine-readable data Table 6. LEO 400 km - L2-NRHO with t2 “ 3 (JSON-format) related to SSD (Solar System Dynamics) and CNEOS. NASA’s Small-Body Database requires an ex- port to an Excel file which then can be converted into an MSSQL(Microsoft SQL Server) database. Both of these 2nd Transfer Total Mass Empty Mass ∆V databases are then consolidated together using the object id as the MSSQL Primary Key. The result is a complete data t2 “ 4 rkgs rkgs rm{ss set with both relative motion values/distances and orbital elements providing access to all needed data from a single to L2-NRHO 1, 640, 000 685, 000 3.212 web point that can now be queried into the tool. Transfer 685, 000 297, 500 3.068 2. Tool of integrating information Return to 400km 297, 500 90, 000 4.398 Second, a Google doc spreadsheet is used as the computational environment to which database queries are passed. This Table 7. LEO 400 km - L2-NRHO with t2 “ 4 communication is done via GOOGLESCRIPT, derived from JAVASCRIPT, enabling the tool to work with on-line information from all JPL’s SSD and CNEOS, as shown in Figure 10. This structured data allows us to do two things. One, Therefore, each ARM would need 20 launches, t1 ˆ t2, from since all systems are connected on-line, the orbital parameters the Earth to fully refuel the spacecraft at L2-NRHO. At a for each selected asteroid allow us to build/propagate an glance, this is an overwhelming number of launches for a Asteroid Retrieval Mission (ARM) using real time JPL Data single mission. However, considering the reusable capability with a single button. In another words, orbital parameters are of SpaceX rocket BFR and the long term goal of achieving undated from the JPL database, calculations are recalibrated a Multiple Asteroid Retrieval Mission, 20 launches actually each time JPL’s SSD and CNEOS add or update the infor- makes sense. mation. And two, the spreadsheet is home to several other parameters required to compute a complete mission profile, Consider reusability again. Although each BFR is estimated e.g. rocket performance parameters, method of predicting to cost around USD 50 millions, this study estimates a fuel- asteroid diameter, et cetera. Each input parameter can be cost for each launch around USD 500,000. This may seem a adjusted or fixed based on business cases and mission needs small cost, but according to SpaceX for a Falcon 9 launch, the in order to find a suitable Multiple Asteroid Retrieval Mission fuel-cost is estimated in USD 200,000. Preliminary results (MARM). indicate 2.7 missions per year, and thus the total annual cost for refueling launches would be only around USD 10 3. STK Integration millions. At the end, enriching the scale of BFR would benefit not only this private company, but the space industry costs as Last, the Multiple Asteroid Retrieval Mission and all its data a whole. from each selected asteroid are ready to be pushed into AGI’s STK (Satellite Tool Kit). Using the provided Python API, all data can be projected for further analysis and refining mission 7. METHODOLOGYAND MARM TOOL planning of each retrieval. OVERVIEW Final output Given Lagrange point analysis as a starting criterion for All of this data culminates into a single stand-alone tool mission development, we utilized Near Earth Asteroid Or- capable of calculating all mission parameters for all ap- bital Data gathered from NASA’s JPL Center for Near Earth proaches in CNEO’s, currently over 750,000 possible Near Objects Studies (CNEOS) including database [12] which Earth Asteroid missions from start to finish. Perhaps best of provides detailed data for every Near Earth Object (NEO) all, the Google Sheets environment is an incredibly robust and to include its orbital elements, close approaches predictions, free platform that a large section of the population has some interactive orbit viewer, and other ancillary data such as experience using. This tool takes 3 parameter sets as inputs: discovery circumstances. However, relevant data needed to calculate a complete mission is stored in 2 separate databases 7 Figure 8. Method and MARM Tool Framework Architecture

1. Asteroid orbital parameters Asteroid Mass Estimation 2. The Lagrangian point choice as base of operations 3. Spacecraft of choice parameters After receiving, via API, the information on the candidates, estimations for the mass of the asteroid are the first step. Ac- cording [E. Bowell][15] and [A. Harris][16], the expression Set 1 is pulled from NASA JPL online databases without for diameter d in km as a function of absolute magnitude H human-in-the-loop. Set 2 is a definition from LOP-G posi- and geometric albedo a is given by the following equation. tion, which would be the base of operations for the MARM case. And Set 3 is the spacecraft of choice for this operation, d “ 100.5p6.259´log a´0.4Hq (6) where this paper used BFR for its refueling capabilities. By using the above expression, we assume a spherical object The resulting output is a list of all possible asteroids that with a uniform surface (no albedo variation). The albedo used are within range of being retrieved using near future rocket in our calculation is assumed to be 0.25. technology, along with several need to know parameters for each potential mission, e.g., total ∆V , Time of Flight (TOF), The density value returned by spacecraft measurements is date and time of departure, amount of propellant and so on. bulk density, which is the mass of an object divided by its Using an Object Oriented approach. The Google Sheet is volume (including the volume of its pore spaces) [17]. The separated into several tabs with each performing calculations ratio between grain and bulk density is the porosity, the for a specific segment of a mission. All calculations are then percentage of the bulk volume of a rock that is occupied by evaluated to return a True or False value to indicated if that empty space. Porosity can be a major component of asteroid particular asteroid can be returned given the specified sets of volume, and some porosity is found in most meteorites. In parameters above. our case, the average density of asteroids is 1.5 g{cm3 with a porosity coefficient of 30% of the asteroid’s volume.

8. METHODOF ASTEROID SELECTION Capture and Return total ∆V Missions are classified as viable if the total mission propellant Combining the information gathered from JPL6 database via required to achieve sufficient ∆V for intercept, rendezvous, API and mass, we need to gather the classical orbital elements and return with the asteroid mass is less than the total amount from each asteroid to begin the calculations for a possible of propellant available in the spacecraft system. capture estimate. In order to combine JPL6 Close Approach to JPL SmallBody. The proposed method downloaded the For capturing the asteroid, assumptions are that the engineer- latter into a CSV file and uploaded it into a private SQL server ing part of capturing the asteroid is already solved and that that we could use with a connection. This was necessary BFR is capable of grabbing, maneuvering and returning with because JPL does not provide, for now, an API for the COE’s the asteroid. Moreover, assumptions that the procedure for for Small Bodies. grabbing and starting the return will be instantaneous at the close approach were considered. According to the Bulk densities DENSITY 8 This paper divides this method into three steps: VA is the Asteroid’s velocity upon encounter given by

1. Ejection: Hohmann transfer from LOP-G VA “ VrA ` 29.7845km{s 2. Injection: Rendezvous combined with Inclination change 3. Propelant needed for retrieving the asteroid All velocities in this step are relative to the Sun. δi is the angle formed between the two vector velocities upon encounter. After step 3, if the spacecraft has propellant left, the asteriod Assuming CNEOS gives us the Earth’s relative δi , and that is selected as a candidate. This paper makes the assumption CA a cislunar plane, δiK “ 5.14 deg in the ecliptic plane angle, that the Earth and Asteroids are in circular orbits. the final δi is given by 1. Ejection δi “ δiCA ´ δiK (12) To calculate the total ∆VEjection the method proceeds by stacking: 1) the ∆V for escaping Earth’s gravitational influence plus the ∆V for reaching the asteroid given the 3. Propelant needed for retrieving the asteroid knowledge of its ecliptic radius, and 3) ∆V change for inclination change, given knowledge of ecliptic inclination In this step, this paper’s method derives an equation for the for the asteroid. mass of Mp needed, assuming that after a spacecraft-burn, the difference between total mass M0 and burnout mass Mb 1.1 Escaping the Earth’s gravitational influence is the propellant mass. Using the Tsiolkovsky rocket equation - mostly known as ideal rocket equation -, Given the assumption that LOP-G is placed at L2-NRHO, the total ∆V is Hohmann transfer from L2-NRHO to the Aster- M0 µ 3.98588e5km3 s2 ∆V“ueq ln oid in the ecliptic plane. With Earth “ { Mb and rL2 “ 456, 100km during lunar apogee, the required to escape the Earth-Moon system (L2-NRHO): or ´ ¯ c ´∆V M M 1 e ueq 2µC p “ 0 ´ Vesc “ “ 1.322km{s (7) rL2

Assuming current orbital velocity VL2 “ 0.934km{s we get assuming the payload mass and the structural mass does not change in the burn, discretizing the mission in a n-stage step- ∆Vesc “ 0.387km{s (8) problem. Thus, for each k step of the mission, it can add or subtract mass emulating the logistic steps for an ARM as follows 1.2 Asteroid relative distance rA «˜ ¸ ˆ ˙ff N k ÿ ÿ ÿ ´∆V k Each asteroid has an estimated relative close approach mea- 0 i k u M “ M ´ M ˘ δM 1 ´ e eq sured in astronomical units, AU. Since this measurement is p 0 p k“1 i“1 relative to Earth, we add 1 AU to get the radius from the (13) asteroid on the ecliptic plane. Assuming the Earth’s radius 0 where M0 is the initial total mass of the rocket of choice and to be 1 AU for every mission, it calculates the extra ∆V to k achieve the asteroid. For this calculation it is used ˘δM is the mass added to the system at step k, and ∆V is „c c  twice the sum of steps 1 and 2, given by 2µ@ 2µ@ µ@ km ∆Vtransfer “ 29.7845 ´ ´ ∆V “ 2p∆Vinjection ` ∆Vejectionq (14) rC rC ` rA rC s (9) where 29.7845 km/s is the velocity of the Earth on the ecliptic this equation calculates, for any rocket available how much orbit. propellant it will need for each asteroid mission. Thus, if the amount of propellant is within the limits of the rocket and 0 Then, the ∆VEjection for reaching the asteroid is given by the below M0 , then the candidate asteroid is selected. sum of 1.1 and 1.2: Set of candidate ARM ∆V “ ∆V ` ∆V (10) Ejection esc transfer Therefore, using a online spreadsheet tool that propagates throughout all closed approaches, the method construct a set 2. Injection of ARM, thus creating the MARM operation. Each mission has all previous information from CNEOS and NASA For this injection, the paper assumes that a combined incli- databases coupled, such as day of departure, time of flight nation correction with the relative velocity, namely VrA, of and so on. the asteroid from the Earth, given by the observations from CNEOS. Using the cosine law we have Dividing the amount of ARM by the years, we get the total a missions per year that need to be operated. And considering 2 2 ∆Vinjection “ |VBFR| |VA| ´ 2|VBFR||VA| cospδiq that upon arrival, the BFR will be refueled and reused, the (11) paper calculates how many BFRs will be needed for the where VBFR is ARM BFR’s velocity upon encounter given MARM, by by c ÿ 2µ@ 2µ@ TOF VBFR “ 29.7845km{s ´ F leet “ ARM (15) rA rC ` rA lifespan 9 where TOF is the MARM average time of flight, in years, 1.1 Escaping the Earth’s gravitational influence using eq. (8): with each ARM TOF given by d ∆Vesc “ 0.387km{s a3 TOF “ 2π t years (16) µ 1.2 Asteroid relative distance rA using eq. (9): @ „c c  2µ@ 2µ@ µ@ km ∆Vtransfer “ 29.7845 ´ ´ 0 rC rC ` rA rC s Furthermore, assuming the capability for a BFR as: M0 “ 0 1650t, structural mass Ms “ 50t and Isp “ 375s a plot of where rA “ 1.041AU, then using canonical units: the mission capability in terms of ∆v and asteroid mass is achieved. (figure 9). ∆Vtransfer “ 0.2977km{s

Using eq. (10): km ∆V “ 0.387 ` 0.2977 “ 0.6847 Ejection s

2. Injection using eq. (11): a 2 2 ∆Vinjection “ |VBFR| |VA| ´ 2|VBFR||VA| cospδiq c km 2µ@ 2µ@ VBFR “ 29.7845 ´ s rA rC ` rA km V “ 28.8972 BFR s Figure 9. BFR-Tanker Capability km V “ V ` 29.7845 A rA s km 9. A CASEOF STUDY V “ 36.066 A s Using Asteroid ”2014 WE6” as our case of study, we can ˝ systematically determine the possibility for capture using our Considering δiK “ 5.14 method given the data in Table 8. δi “ 0.3391˝ ´ 5.14˝ “ ´4.8˝ a ∆V “ |V |2|V |2 ´ 2|V ||V | cospδiq 2014 QN266 Data injection BFR A BFR A km ∆V “ 3.995 Closest Approach Date/Time March, 30th of 2058 @03:42 injection s

Relative Distance [rA] 0.041 AU Adding both: Relative Velocity from Earth [V ] 3.281 km/s km rA ∆V “ 4.680 rendezvous s inclination(i) 0.3391 deg magnitude (H) 30.4 3. Propelant needed for retrieving the asteroid: Table 8. JPL’s CNEOS and NASA Small Object databases Using eq. (13): «˜ ¸ ˆ ˙ff N k ÿ ÿ ÿ ´∆V k 0 i k M “ M ´ M ˘ δM 1 ´ e ueq Using eq. (6), we can calculate the diameter of the asteroid: p 0 p k“1 i“1 0.5p6.259´log 0.25´0.4Hq d1 “ 10 km Mp “ 1, 516 metric tons of propellant for the whole mission, where δM “ Then, the average diameter for this asteroid is: 6.1912 metric tons (mass of the asteroid). Thus, by com- paring to the amount of propellant in Table 9, the asteroid d “ 2.24m becomes a candidate. Then using the linear extrapolation from the density using the The tool also calculates the TOF, using eq. (16): the volume of a sphere, we get mass m “ 6.1912 metric tons. d a3 The following steps to get the results are detailed in section TOF “ 2π t years 8. µ@ 1. Ejection: TOF “ 188.3days 10 Although this technique is generally accepted in the scientific ARM BFR Data community, this paper acknowledges that this is a theoretical ISP Vacuum 375 sec approach and there may be some inconsistency on smaller objects, which are the first choices for their lower ∆V ARM. Thrust 1900 kN The method chosen for rendezvous trajectory of each ARM is Empty Mass (Tanker) 90 metric tons consistent. This paper takes into account the relative motion catalogued by JPL CNEOs close approaches observations, Propellant 1550 metric tons specifically taking into account the relative velocity and the ecliptic inclination, which have an enormous effect on ∆V Total Mass 1640 metric tons calculations. However, the paper acknowledges that further Ueq 3678.75 studies might result in better choices for rendezvous trajectory, especially if taking into consideration an individual Table 9. ARM BFR data ARM approach. The method for calculating the propellant needed for each mission is interesting, because it enables the hardware provider and the private investor to discuss and further ana- 10. CONCLUSIONS lyze what would be the optimum size specs for a customized BFR depending on the MARM objective. Orbital positions for LOP-G, the Lagrangian points L1 and L2 were studied and this paper concludes that incrementing Then, this paper ran the tool for a query of close approaches a MARM operation into NASA’s LOP-G will result in L2 between 2030 and 2060 resulting in 33 ARM candidates, becoming more desirable in terms of ∆V . Furthermore, this listed in the appendix. The resulting MARM adds up to result is sensitive to the operations on the LOP-G. On one 2,581 metric tones of asteroids to mine. The complete list is hand, more interplanetary operations will choose L2 as a also presented in the appendix. Although the total amount of preferred launch point. On the other hand, more LEO/LOP-G metric tones in this MARM seems low compared to the total operations will elect L1 as the best option. For creating the amount of minerals explored on Earth by the top 5 companies, MARM, this paper assumes L2 for LOP-G position. having a start of 2,581 metric tones in space will lead to new methods and technology to explore further and increase the With this assumption, this paper progresses to the field of MARM throughout the next decades. In Fig. 10 this paper studying hardware capabilities chosen for a MARM mission. presents a plot that correlates total mission ∆V and relative With the refueling capability of SpaceX’s BFR, this paper velocity from each asteroid on the MARM. derives an operational refueling system comprising 10 BFRs for all 30-year life span of the MARM operation: 3 BFRs Having around 1 ARM each year and considering average deliverinig fuel to 2 BFRs at 400 km at LEO orbit, while these TOF to be 190 days for rendezvous and another 190 days for 2 BFRs would deliver the fuel to a fleet of 5 BFRs on LOP-G. returning, the total number of BFRs needed for ARMs drops While at a glance this operation might be disconnected from from 5 to 3. reality by today’s standards, this paper presents evidence that a 20 launches for 1 ARM from LOP-G could be easily done Finally, in this MARM project every ARM mission are below USD 10 million by 2030, where the MARM is set to designed to bring the whole asteroid back. This decision start. significantly reduced the number of minerals of the MARM available for the 30-year project. Retrieving partial mass An initial starting point to understand a cost per kg is to from a single asteroid in multiple ARMs would increase the considering only acquiring the hardware and refueling costs number of minerals in reach. The amount of ARMs available for the 30-year project. The amount of money needed would to each asteroid throughout a 30-year operation should prove be around USD 330 million for refueling and USD 2 billion important to determine a new number of minerals available. for the BFRs, assuming that a BFR would cost 200 million This provides a good starting point for discussion and further to be delivered on L2 orbit. Dividing this total of USD 2.330 research and, thus, may indicate a substantially reduction in billion for 2.581 tones of minerals would result in USD 902 the cost per kg from this study. thousands cost for each kg of minerals. It is important to note that there are many other significant costs in such operation The total amount of minerals in space enormously dwarfs and that this result gives only an idea from the cost. Further the amount of minerals on Earth, and by initiating a small research and study to develop a proper business plan for this operation in space, with a low relative investment, a mining operation is needed. company, a consortium or even a new company would not only have a great chance of establishing the new mainstream By creating a method and an on-line tool, this paper intro- mining for the market but also develop new technology for duces the opportunity for further analysis. JPL’s CNEOS data further exploring the Solar System. is growing and becoming more and more accurate and that might be an indicative of further growth on the set of ARM candidates. Furthermore, hardware providers with the private investors can use the tool to embrace enhancements on the current design in order to add the range needed for a best optimum scenario in space mining. It is important to note that as better observations are made by the community, the results might change. The method that selects the asteroids uses the albedo and absolute magnitude to estimate the diameter of the asteroid. 11 Figure 10. Possible Captures Diagram; Circle sizes proportional to the mass of the asteroid

APPENDICES A.MARMCANDIDATES All selected asteroid for possible missions with their total ∆V . The amount of ∆V is the sum of going to each asteroid, ID DES LOP-G Departure ∆V adding the mass of the asteroid to the spacecraft and returning 30 2013 UX2 4/27/2058 6:16:51 PM 7.884 to the same point of departure. 31 2017 QB35 3/2/2052 2:21:22 PM 9.758 32 2017 JB2 5/13/2049 2:33:44 AM 7.803 ID DES LOP-G Departure ∆V 33 2011 CQ1 7/15/2046 10:46:00 PM 10.512 1 2018 AV2 10/24/2056 12:11:28 AM 6.384 2 2007 UN12 4/20/2049 6:10:48 AM 7.792 3 2012 WR10 4/11/2038 8:11:17 PM 7.726 4 2014 WE6 9/22/2057 5:23:21 PM 9.360 5 2009 BD 10/10/2045 11:44:42 AM 6.603 6 2015 PS228 10/14/2039 10:37:28 PM 7.282 7 2008 JL24 11/13/2044 12:18:16 AM 8.968 8 2010 UE51 6/23/2054 6:41:04 AM 6.356 9 2013 CY 5/16/2032 4:54:43 PM 7.532 10 2014 JR24 3/21/2047 12:52:11 AM 8.972 11 2015 KK57 1/30/2056 10:15:16 PM 6.011 12 2011 UD21 7/31/2042 6:14:39 PM 6.242 13 2014 WU200 7/27/2039 2:03:37 PM 6.093 14 2010 VQ98 10/24/2039 11:37:49 AM 5.688 15 2017 FT102 11/3/2053 1:24:53 AM 6.154 16 2017 BG30 12/20/2042 4:00:59 AM 9.629 17 2017 KJ32 11/19/2041 12:20:50 AM 7.343 18 2014 BA3 5/12/2044 6:39:14 AM 7.041 19 2010 JW34 10/28/2044 6:59:59 AM 5.369 20 2012 FS35 1/5/2050 2:40:24 PM 9.798 21 2015 VU64 8/14/2053 3:01:24 PM 8.915 22 2011 MD 12/9/2048 11:11:10 AM 5.347 23 2016 SX1 4/17/2054 9:02:57 AM 8.665 24 2000 LG6 12/21/2035 12:47:37 PM 6.609 25 2018 BC 7/21/2046 8:12:26 PM 5.490 26 2012 AQ 6/30/2053 5:30:15 AM 7.719 27 2012 LA 4/30/2047 6:55:45 PM 5.480 28 2016 EU84 9/8/2053 6:10:36 AM 6.116 29 2013 GH66 10/12/2039 11:09:07 PM 5.395

12 REFERENCES Engineering Ph.D student in Kevin T. Crofton Aerospace and Ocean Engineering Department at Virginia Tech and [1] F. A. A. (FAA), “The annual compendium of comercial a Graduate Research Assistant for The Hume Center for space transportation: 2018,” January 2018. National and Security Technology and for Center for Space [2] R. T. SA, “2017 annual report,” January 2018. Science and Engineering Research (Space@VT). His current [3] E. Musk. The falcon has landed! epic research activities include On-Orbit Servicing focused on views of spacex’s amazing rocket landing. Autonomous Unmanned Aerospace Systems. [Online]. Available: https://www.space.com/31444- spacex-falcon-rocket-landing-epic-photos.html David Alexandre received his B.S. de- [4] (2018, October) The lagrange point l1 between gree from Universite´ d’Angers (France) the earth and the moon. [Online]. Available: and his M.S. degree from Ecole Centrale http://www.ottisoft.com de Nantes (France) in 2015 both in Me- chanical Engineering. He is currently a [5] W. E. Wiesel, Space Flight Dynamics, 2010. Ph.D. research scientist in Space Engi- [6] NASA. Soho mission overview. [Online]. Available: neering in Kevin T. Crofton Aerospace https://www.nasa.gov/mission pages/soho/overview and Ocean Engineering Department at Virginia Tech. He worked at AIRBUS [7] D. Guzzetti, E. M. Zimovan, K. C. Howell, and D. C. (France, United Kingdom, Germany) Davis, “Stationkeeping analysis for spacecraft in lunar between 2012 and 2015 as a Quality Manager in the A320, near rectilinear halo orbits,” 2017. A330 and A380 aircraft programs. His current research [8] P. K. Martin. (2018, May) Examining the future of activities include Computational Modeling using Sonar Sys- the international space station. [Online]. Available: tems. https://oig.nasa.gov/docs/CT-18-001.pdf [9] NASA, “Gateway memorandum for the record,” May Marco Peterson received his B.S. and 2018. M.S. degrees in Computer Science from Virginia State University in 2015. He [10] J. C. Crusan et al., “Deep space gateway concept: is also a Graduate of the United States Extending human presence into cislunar space,” May Army’s Aviation Operations and Flight 2018. Schools and is currently serving as a [11] NASA. Space launch system lift Rated Rotary-Wing Pilot. He is cur- capabilities. [Online]. Available: rently pursuing a PhD in Aerospace En- https://www.nasa.gov/sites/default/files/atoms/files/sls gineering in Kevin T. Crofton Aerospace lift capabilities and configurations 508 08202018 0.pdf and Ocean Engineering Department at [12] N. J. P. Laboratory. Center for near earth object studies. Virginia Tech with a research focus in Unmanned Aerial [Online]. Available: https://cneos.jpl.nasa.gov platforms. [13] P. Chodas and S. Khudikyab, “Neo earth close approaches,” 2018. [Online]. Available: Kevin Schroeder Research Faculty in https://cneos.jpl.nasa.gov/ca/ the Kevin T. Crofton Aerospace and Ocean Engineering Department at Vir- [14] R. S. Park and A. B. Chamberlin, “Jpl small- ginia Tech. Kevin received his B.S. in body database,” May 2018. [Online]. Available: Engineering from Oral Roberts Univer- https://ssd.jpl.nasa.gov/sbdb query.cgi sity in 2014. He went on to receive his [15] E. Bowell, Asteroids II, 1989, pg. 524-556. Ph.D. in Mechanical Engineering from Virginia Tech in 2017. As a Ph.D. candi- [16] A. W. Harris, On the Revision of Radiometric Albedos date, Kevin studied Entry, Descent, and and Diameters of Asteroids, 1997, icarus 126:450-454. Landing (EDL) systems and became a [17] D. Britt, D. Yeomans, K. Housen, and NASA Innovative Advance Concepts Fellow for his invention G. Consolmagno, “Asteroid density, porosity, of TANDEM. Following graduation, Kevin was hired by and structure,” May 2018. [Online]. Available: Virginia Tech to work in the Center for Space Science and https://www.lpi.usra.edu/books/AsteroidsIII/pdf/3022.pdf Engineering Research (Space@VT). Currently, Kevin serves as the technical lead on multiple projects with a focus on Astrodynamics, Optimal Decision Processes, and Autonomy BIOGRAPHY[ of Dynamical Systems.

Gustavo Gargioni received his B.S. degree in Mechanical/Industrial Engineer- ing from Instituto Maua de Tecnologia (Brazil) in 2002 and a Business Spe- cialization from Fundacao Getulio Var- gas (Brazil) in 2005. Since college, he pursued a software and industrial entrepeneur career. While C.E.O. of Ecoplasticos Industria de Reciclagem Ltda, he received the 2011 Environmen- tal Award from Federacao das Industrias do Estado da Bahia (FIEB, Brazil). Since 2017, he has been an Aerospace