Target Selection and Architecture Design for a Manned Near Earth Object Mission

Target Selection and Architecture Design for a Manned Near Earth Object Mission Ralf Boden 1),2), Michihiro Matsumoto 2), Junichiro Kawaguchi 3),4) 1) Technical University Munich (TUM), Germany 2) The University of Tokyo, Japan 3) JAXA Space Exploration Center (JSpEC), Japan Aerospace Exploration Agency (JAXA) 4) Institute of Space and Astronautical Science (ISAS), Japan Aerospace Exploration Agency (JAXA) Abstract: In this paper, a first concept for a manned NEA mission is developed, including estimations of required efforts and mass of the spacecraft as well as a target selection. A first estimation of the in-leo masses required to successfully conduct a manned NEO mission can be calculated by examining the v requirements for the trajectories to the currently known NEO population. Using this data, a list of possible mission targets and opportunities, as well as example missions, based on different objectives and properties are presented. 1 Introduction 1.1 Interest in NEO missions Interest in the human exploration of asteroids within close proximity to Earth, the so-called Near Earth Objects, continues to increase due to the following factors. First, NEOs can serve as a stepping stone toward developing a lot of the technologies, necessary to reach Mars, as a main goal of most current human space exploration programs. While both missions have similar requirements with respect to deep space travel, asteroids are more accessible when v and launch availability are considered. These similarities provide an advantage of NEOs as targets, when compared to the Moon and have caused a shift in the priorities of human exploration missions towards these asteroids [1]. Second, NEOs are attractive targets for exploration on their own account, as samples from them can increase our understanding of the creation and early development of the Solar System. These samples can also provide detailed knowledge of the objects themselves. While much of this science could be achieved with robotic missions, a human presence can help maximize the benefits of sample return, by choosing the most promising sample locations and the best samples. Humans therefore add overall flexibility that is currently not possible with robotic missions [2]. A third point of interest in these asteroids, is the fact that some NEOs are potential Earth-impactors due to their orbit geometry. While the chance of one of these potentially hazardous asteroids (PHAs) colliding with Earth is small, sending a manned crew to such a threatening object is a first step in developing methods to deal with a future calamity and would provide a chance to test possible mitigation strategies. 1.2 Initial Restrictions Y Axis [AU] 6 4 2 2 4 6 6 4 2 2 4 6 X Axis [AU] Figure 1: Position of analyzed NEO population, as of Oct. 21. (heliocentric; in plane) Current limitations in technology and experience with manned deep space travel restrict potential NEO missions and can eliminate some of the possible targets from the prospective list of 6572 NEOs used as a basis for this study [3] (see Figure 1). 1

While all asteroids on this list are classified as NEOs, not all are suitable targets. This first step in narrowing down the list is by restricting the required mission v and duration, calculated through each asteroid s orbital elements. A mission duration limit of one year was chosen, potentially longer than a one-way transfer to Mars. The use of chemical propulsion was assumed to minimize transfer times and also adds a certain level of safety due to the broad experience we have with these systems. The actual amount of v for propulsive maneuvers is limited to 8. km/s, a value identical to the requirements of a manned mission to the lunar surface. A complete listing of the initial restrictions is shown in Table 1. Additional parameters such as NEO magnitude and amount of time spent at the target are kept relatively unrestricted as they can vary depending on the exact mission objectives. The time frame for these missions is expected to take place between 215 and 235. Table 1: Initial restrictions, used in this NEO mission study. Value Description Unit Limits mission duration dt mission [days] 365 time at NEO dt target [days] 5 max. velocity v total [km/s] 8. NEO size (mag.) H [ ] 28. 2 Mission Design 2.1 Outline The NEO mission design presented here will use concepts and technologies that are proven to be reliable, as these missions are the first step into deep space and the first manned missions to leave LEO in some time. Manned missions inherently must have a higher safety margin. Unlike in the case of an unmanned spacecraft, mission failure can lead to the loss of lives, which can be detrimental for future manned space exploration efforts. Disasters could lead to a questioning of human spaceflight in general and set all future plans on hold for extended periods of time. Thus, the implementation of new technologies occurs gradually, with initial mission designs often relying on concepts that have already been successfully used on previous missions. With this in mind, the mission design will utilize a spacecraft setup similar to the Apollo missions: a single spacecraft for both crew and equipment traveling to the target asteroid. This eliminates the need for en-route re-supply and minimizes the need for complex maneuvers during transfer and at the NEO. The mission profile of this spacecraft will consist of the two ballistic transfer trajectories between Earth and the target asteroid and the time spent at the NEO itself (see Figure 2). This concept will be the basis for calculating the required v s in Section 3. Outbound Trajectory Earth Departure NEO departure NEO arrival Target Asteroid Earth Return Inbound Trajectory Earth Figure 2: Diagram showing the two-impulse transfer trajectories that will be used by the spacecraft to travel between the target asteroid and Earth Additional safety to this concept can be added by designing the outbound transfer trajectory to have an orbital period of close to 365 days. This allows the spacecraft to automatically re-encounter Earth within a one year period, without the need for additional propulsion maneuvers after Earth departure. This can be helpful in case of engine failure and is similar to the free-return trajectories used for the Apollo Mission, which also allowed the spacecraft to automatically return to Earth. However, this requires a spacecraft design capable of supporting the crew for an entire year, independent of the nominal mission duration. The basic trajectory analysis is done by using the patched conics method. This method connects the spacecraft departure trajectory, within Earth s sphere of influence (SoI), to the actual transfer trajectory toward the NEO. In this study we assume that the spacecraft departs from an initial parking orbit around Earth by completing a Earth Departure Injection (EDI) and directly transfers onto the NEO rendezvous trajectory. 2.2 Earth Reentry Because of the high re-entry speeds for this type of mission, a ballistic re-entry approach is used got the spacecraft s return to Earth. We will assume a re-entry velocity of no greater than 12 km/s at 2

the Entry Interface (EI), due to technological limitations of the thermal shielding. This value is in-between the current records for fastest re-entry speeds for manned (Apollo 1 v EI = 11.1 km/s) and unmanned (Stardust v EI = 12.4 km/s) vehicles. 2.3 NEO Proximity Operations Since asteroids can be considered as micro gravity environment due to their relatively small size, proximity operations at these objects will not have to deal with special orbit insertion and departure maneuvers. However, depending on the target properties (Section 5) and the detailed mission objectives, there can be a wide range of different proximity operations that need to be performed. This can range from simple flybys and orbiting around the target, to complex surface operations performed by the crew. The scale of these operations will influence the required crew-time at the target (consisting of the actual time spent at the target and the Crew size) and also the mass of the equipment that needs to be transported to the NEO. Depending on the size of required equipment, a separate storage module may be required. This module can be disposed of when not needed anymore, resulting in a reduction of spacecraft mass for the NEO departure and the fuel needed to achieve the necessary v. Due to the possible need of a separate module, we consider two different configurations for the mass model in Section 4 and the choice of best targets in Section 6. A configuration including a separate equipment module is preferred for a sample-return type of mission, where more extensive operations need to be performed at the asteroid. The crew size for such a mission will be assumed as three persons. On a limited fly-by mission such a module will not be necessary and the lack of EVA operations at the NEO allow reducing the crew size to two persons. 2.4 Level 1 Mission Design With all parts of the mission defined, we can establish a design for the NEO mission and the required spacecraft. This design can then be translated into a mass model to estimate the in- LEO mass requirements and associated cost. Both the concept and the spacecraft s general setup are shown in Figure 3. This basic design allows calculating the necessary trajectory data (Section 3) and creating the mass estimation model (Section 4). Since the design in this study is a first concept, we will refer to it as a Level 1 design. Future, more complex designs can base their analysis on this general setup but include more advanced technologies and ideas for the mission architecture. 3 Trajectory Calculation In this section, we use the Level 1 architecture, established in the previous section, to calculate the necessary v requirements for the proposed mission profiles. 3.1 Lambert Problem The first step involves calculating the v values, needed to transfer the spacecraft from the Earth or NEO onto the trajectories between them. v 2 v traj@target vtarget Target Orbit Origin Orbit Transfer Trajectory v traj@origin v 1 Figure 4: Velocities needed to calculate the v for entering and exiting a transfer trajectory. Figure 4 shows the required velocities needed to calculate the v to enter and exit a transfer trajectory. For a round-trip mission, the resulting v s are numbered according to the order in which they occur: v 1, = v traj1@earth v Earth at Earth Departure (t departure ), v 2, = v NEO v traj1@neo at NEO Arrival (t departure + dt 1 ), v 3, = v traj2@neo v NEO at NEO Departure (t departure + dt 1 + dt target ), v 4, = v Earth v traj2@earth at Earth Return (t departure + dt 1 + dt target + dt 2 ). As the velocities of the spacecraft at both the Earth and NEO are identical with respect to that object, these velocities can be calculated from the ephemeris data in Reference [3] and the calculations explained in References [4] and [5]. The starting and ending velocities of the trajectory s path need to be computed by solving the Lambert Problem for each trajectory: v origin s = f( r start, r end, dt). (1) The Lambert Problem states that a path s, connecting two points in space, is defined by its starting and ending position vectors ( r start, r end ) and the 3

Δv 1 Δv 2 1) Launch into LEO (and Assembly) 8) Spacecraft disposal upon Earth return 7) Inbound Trajectory Δv 3 6) EM Disposal and NEO Departure + 9) RM Ballistic Re-entry 5) NEO Arrival and Proximity Operations NEO 2) Departure from Parking Orbit 4) Outbound Trajectory Return Module RM Mission Payload EM Habitat Module HM Propulsion Stage PM 3) EDS Disposal Earth Departure Stage EDS Figure 3: Overview of the Level 1 NEO mission concept. time spent traveling between these points. The solution to this problem provides the velocity vectors at the beginning and end of the trajectory ( v start, v end ) [5],[4]. As the positional vectors are identical to those of the departure/arrival body, we can calculate all paths by circling through the range of possible transfer times dt. Doing this for a complete mission profile, we can compute all trajectories by using the algorithm shown in Figure 5. The loops the calculation program needs to cycle through, in order to obtain every possible profile, are the time dependent parameters in Table 2 and the list of NEOs, considered as possible mission targets. Table 2: Time variables defining each individual mission profile, including their range and the size of a step between values. variable range step size t departure 1 Jan. 215 1 Jan. 235 1 days dt 1 dt 1 + dt target + dt 2 365 d 1 days dt target dt 1 + dt target + dt 2 365 d 5 days dt 2 dt 1 + dt target + dt 2 365 d 1 days 3.2 Adaptation of Values Since asteroids can be considered as a microgravity environment, no conversion is necessary for the v values for arrival and departure at the NEO ( v 2/3 ). Earth s strong gravitational influence within its SoI however, requires additional calculations to obtain the actual values for the required propulsive maneuvers.[6] As the initial Earth departure will be using a EDI from an Earth parking orbit (h park = 5 km), we can use the Patched Conics method to calculate the required v 1 from the v 1, using with v 1park = v 2 1 + 2v2 park v park, (2) µ v park = (R + h park ). (3) In reverse, we can calculate the maximum v 4,, from the reentry speed limit of 12 km/s at the Entry Interface (EI), as: v 4,,max = vei 2 2 µ (R + h EI ). (4) With a height of h EI = 11 km for the EI, the maximum speed results to v 4,,max 4.6km/s. (5) 3.3 Resulting Data The data resulting from these calculations contains all possible mission profiles that fall within the 4

parameter limits for vtotal (= v1 + v2 + v3 ), dtmission and reentry speed ( v4 4.6 km/s). The data reveals that out of the initial NEO population, only 68 targets remain from the initial NEO population. The minimum v for mission opportunities to these targets is shown in Figures 6 and 7. min. v [km/s] 8 35 total mission duration [days] 7.5 3 7 28 EA 25 9 6.5 Apophis (99942) 2 6 2 SG 344 15 5.5 29 OS 5 1 5 28 EA 1999 AO 9 1 5 4.5 215/1/1 216/12/31 219/1/1 22/12/31 223/1/1 224/12/31 227/1/1 228/12/31 231/1/1 232/12/31 235/1/1 departure date Figure 6: Minimum vtotal for free-return profiles to all remaining 68 NEOs. 8 29 OS 7 27 XB min dv total [km/s] 6 5 23 1999 AO 5 1 21 GP 4 2 28 EA 9 minimum v required for Earth Escape 3 2 SG 2 28 ST 344 Apophis 1 5 1 15 2 mission duration [days] 25 3 35 Figure 7: Minimum vtotal over mission duration for each target. Promising options are highlighted. While this data is a first indicator for promising targets, additional considerations are necessary to choose the best options. Sections 4 and 5 will address these concerns and allow listing the best targets for NEO missions, based on the required inleo mass. 4 Figure 5: Flowchart for the calculation of the mission profiles, showing all necessary loops. The saved data includes v1 4 and the time variables from Table 2. Spacecraft Mass Model We can establish a model for the spacecraft that allows estimating the required mass for each mission profile using the data acquired in the last section. Since these parameters influence each of the modules established in Section 2 in a different way, we describe each model separately. 4.1 Equipment Module EM Parameters of the equipment module can be estimated by the amount of equipment required at the NEO. As this depends on the exact mission objectives, we can choose the module s mass based on the two types of configurations defined in Section 5 4

2. For the fly-by mission, we assume a mass of kg for the EM, as neither sampling equipment nor additional EVA supplies and gear will be needed. For the sample-return type mission, we assume an equipment payload mass m pl of 5 kg, referencing the Apollo Lunar Module s equipment payload. When we account for the mass of the structure, the total mass of the EM is calculated as: M EM = 5 kg (1 + r structure ) = 5 kg (1 +.15) = 5 75 kg. (6) Total Habitable Volume per Crewmember [m 3 ] 25 2 15 1 5 fourth order polynomial approximation Optimal Performance Limit Tolerable Limit 1 2 3 4 5 6 7 8 9 1 11 12 Mission Duration [months] 4.2 Return Module RM Another module that is independent of both mission duration and v requirements is the reentry module, used to safely return the crew and any possible samples to Earth at the end of the mission. Using the mass of the Soyuz return capsule (m 3 kg) as basis for this module, we can calculate its total mass M RM as the sum of the capsule and the heat-shield, required for the ballistic re-entry. Since the heat-shield mass is dependent on the mass of the re-entry vehicle, we can calculate the total mass as M RM = 3 kg (1 + r heatshield ) = 3 kg (1 +.2) = 3 6 kg. (7) Because this type of mission results in high reentry speeds, a relatively high mass ratio for the heat-shield is chosen.[7][8] 4.3 Main Spacecraft HM Since the main spacecraft will be used to control all of the spacecrafts systems, as well as provide the habitation space for the crew during the entire mission, this model is the most complex one. In order to estimate its mass, the major mass driving subsystems will be analyzed separately, to determine their mass, and therefore get an estimate for the entire spacecraft. The main factor that influences this mass is the mission duration and the crew size. These two parameters directly affect both the required habitation space, as seen in Figure 8, as well as the masses of the life-support system (ECLSS), which is one of the main mass driving systems, next to thermal control (TCS) and the power supply (PCS). Since we have a fixed duration for which the spacecraft needs to be designed, due to the freereturn setup, all of these systems only depend on the crew size N crew. If we also consider a fixed crew size of either 2 or 3 persons, we can calculate two Figure 8: Per-person habitable volume requirement over the mission duration, taken from Reference [7]. masses for the HM, depending on the type of mission. Using the constant optimal habitational volume V hab = 19 m 3 /person for a 365 day mission, we can calculate the required pressurized volume V press = 3 V hab = 57 [m 3 /person], (8) according to Reference [8], resulting in a structural mass of M struct = V press 1 kg/m 3 [kg]. (9) The value of 1 kg/m 3 is a typical mass to volume ratio for a very light rigid structure module, such as the ones on the ISS.[7] The other systems are modeled using the formulas in Reference [9], with adaptations to account for changes in mission duration and crew size. Table 3 shows the resulting system masses, including the required supplies, for both the 2 and 3 person setup. 4.4 Propulsion Systems EDS and PM As both the Propulsion Module (PM) and the Earth Departure Stage (EDS) provide the necessary Thrust to conduct the v maneuvers, they can both be calculated using the same methods. The only differences are their relative sizes, and the type of fuel they operate with. Since the EDS is used at the start of the mission, it allows the utilization of cryogenic fuels, such as LOX/LH, without the concerns of off-gassing. This allows the EDS to have a better I sp value then the PM, which has to rely on storable fuels, similar to the Aerozine5/NTO combination, also used on the Apollo Service Module. The resulting I sp values for both modules, as well as their structural ratios are shown in Table 4. The difference in r struct comes from the size difference of both modules, which is caused by fuel density and 6

Table 3: Masses for each part of the spacecraft model in [kg] Name 2 Person 3 Person Structure 11 4 17 1 ECLSS Air 1 675 2 425 ECLSS Water 2 736 4 19 TCS 1 839 2 252 PCS 215 36 ACS 276 276 Onetime Supplies 387 581 Air Supply 849 1 244 Water Supply 2 22 2 979 Other Daily Supplies 1 666 2 498 Total Mass 23 65 33 68 the fact that the EDI is larger, due to the higher mass it has to transport. Table 4: Parameter Values for both EDS and PM Parameter Name EDS PM Specific Impulse I sp 45 s 35 s Structure Ratio ɛ struct.8.1 Using these parameters, the exact masses of the modules can be calculated as the sum of their main components. These are the fuel mass, M fuel, the tank structure, M struct, needed to hold the fuel, and the mass of the engine, M engine. Because M fuel affects both the structure and engine masses, and these in turn, need to be added to the spacecrafts burnout mass M bo in the Tsiolkovsky Equation: ( ( ) ) v M fuel = M bo exp 1, (1) I sp g these three masses need to be calculated iteratively, to get the total mass of both the PM and EDI.[7][6][8] 5 Target Selection Now that we have established the models to evaluate the calculated mission data, one last concern that affects the choice of targets is addressed, before the final results are presented in Section 6. While restrictions towards target size have been kept very loose until now, only excluding targets with magnitudes over 28., a decent target size can be a main concern for mission objectives related to scientific exploration. As scientific missions are most promising when the target asteroid has a decent size, smaller objects become less interesting as they offer less incentive for exploration. In case of sample return missions, our choice of targets is also limited to asteroids that allow a crew to safely land and operate on its surface. Allowing this will require some restriction regarding the rotation speed of the NEOs in question. The limit we will be using is a rotation period T of at least 1 2 hours. While direct landing might only be a problem for sample return missions, fast rotators pose additional safety risks for any type of proximity operation, and should therefore be avoided as much as possible. While the actual rotation rate of most NEOs is unknown, its coupling with their size can be used as a first indicator. Because larger asteroids are usually made up of a rubble pile structure instead of being one single solid piece, they can only exist if they are held together by a large enough gravitational force. This force acts on every particle that the asteroid is composed of and needs to be larger than the acceleration caused by the object s own rotation. Objects for which this is not the case would not be able to exist, as they are torn apart by their own rotational speed. A first estimation for analyzing the sum of these accelerations can be made by using a sphere as a simplified model for the asteroid s shape. If we assume a constant density ρ, the asteroid mass M = V ρ = 4 3 πr3 ρ (11) is only dependent on the asteroid s radius r. From this we can calculate the gravitational acceleration, acting on the particles on the asteroid s surface, as a grav = GM r 2 = G 4 3 πr3 ρ r 2 = 4 Gπrρ. (12) 3 While this force keeps the rubble pile together, the acceleration caused due to the object s rotation a rot = v2 r = (ωr)2 = ω 2 r, (13) r acts in opposing direction. Since both ρ and G are constant, the sum of these two forces is only dependent on the angular velocity ω and the object s size, defined by its radius r. Figure 9 shows the sum of these accelerations for a number of different ω, plotted over the target size. The conversion from radius r to magnitude H is done via ( ) p H = 5 log 2r, (14) 1329 7

with a value of.15 chosen for the asteroid s albedo p. This represents a typical rocky asteroid. acceleration [mm/s 2 ] 1 1 2 3 4 3h 2h 1h 45min 1h 1h 3min 16 18 2 22 24 26 28 absolute magnitude H Figure 9: Sum of gravitational (12) and rotational (13) acceleration for a simplified NEO model. Positive y-axis values indicate that the gravitational outweighs the rotational acceleration on the NEO surface. We can see that the rotational acceleration exceeds the gravitational pull for periods under 2 hours. This means that only slower rotating asteroids are able to exist as rubble piles without being torn apart by their own rotational speed. Figure 1 validates this simple estimation, showing that targets above a size of 21.5 magnitudes have periods above 2 hours. Smaller targets do not show this behavior, as they are not necessarily rubble piles. While this does not mean that smaller targets cannot have these slow rotational periods too, we can only be certain for larger objects. Because of this, the main focus for sample-return missions is on NEOs with magnitudes H 21.5.[1] rotation period [h],1,1,1 21.5 2h be considered for other mission types, such as flybys and might be a choice for very early NEO missions, or because of general spacecraft mass limitations. Asteroids above 21.5 magnitudes may also be a better target for testing mitigation strategies, as too large targets would not be significantly influenced by efforts to change their trajectory. number of available targets 1. 1. 1 1 total NEO population all available mission targets free return option targets 1 8 9 1 11 12 13 14 15 16 17 18 19 2 21 22 23 24 25 26 27 28 absolute magnitude H Figure 11: Comparison of initial NEO population with remaining results over the magnitude range. This means that deciding which asteroids will become targets for NEO missions depends on the mission objectives. While scientific research is always the preferred choice, existing mass limitations and limited budgets may shift the focus towards targets that are less ideal, but more affordable to reach. With the two possible setups, established in Section 2, we can define the sample-return setup to be the preferred choice for larger targets (H 21.5). And all other asteroids as options for flyby missions, since it is uncertain if they allow surface operations. 6 Results With the mass estimation model completed and all trajectory profile data available, the final results of this study, in the form of in-leo mass requirements, can be calculated and used to determine the best targets and mission opportunities. Since the main focus is on missions that allow scientific research, the presented data will focus on the heavier sample-return setup for the NEO spacecraft. 1 3 28 26 24 22 2 18 16 14 12 absolute magnitude H Figure 1: Distribution of rotational period and magnitude for all known asteroids, where data is available. While this means that smaller targets are not really suited for sample-return missions due to their tendency to be fast rotators (Figure 1), the extremely small number of available large NEOs (Figure 11) requires looking into these smaller asteroids as possible mission targets too. These objects can 1 8 6 4 2 total mission duration [days] 36 35 34 33 32 31 3 Apophis (99942) 224/1/1 224/7/1 224/12/31 225/7/2 225/12/31 226/7/2 227/1/1 227/7/2 228/1/1 228/7/1 departure date min. LEO Mass [tons] Figure 12: Minimum M LEO for ideal sample-return targets.(h 21.5) 5 45 4 35 3 8

1.8.6.4.2.2.4.6.8 1 229 3 8 229 1 27 229 2 6 229 4 7 229 4 27 229 1 7 228 5 12 228 12 8 228 6 11 228 11 8 228 7 11 228 1 9 228 8 1 1.8.6.4.2.2.4.6.8 1 X Axis [AU] 228 9 9 1.8.6.4.2.2.4.6.8 1 229 3 8 229 3 28 228 4 2 229 2 6 228 5 2 228 6 1 229 1 7 228 7 1 228 12 8 228 7 31 228 11 8 228 8 1 1.8.6.4.2.2.4.6.8 1 X Axis [AU] 228 1 9 228 9 9 6.1 Ideal Sized Target Options If we also limit our targets to the sizes necessary for this preferred mission type, the remaining targets and required in-leo masses result to the ones seen in Figure 12. Individual targets are shown in Figure 13. One of the few remaining options in this category is Apophis (99942), with a minimum mass requirement of 291 tons, for the sample-return setup. Spacecraft in LEO Mass [kg] for 3 person sample return mission 5. 45. 4. 35. 3. 25. 2. 15. 1. 5. 27 XB 23 29 OS 5 1999 AO 1 28 EA 9 2 SG 344 21 GP 2 28 ST Apophis 5 1 15 2 25 3 35 mission duration [days] Figure 13: Minimum spacecraft mass for individual targets over mission duration. As an example for such a sample return mission, the minimum mass mission to Apophis is presented in Figure 14, with the mission profile data from table 7 Line 2. An overview of the spacecraft module masses is listed in Table 5. Table 5: Module Masses for the Lowest-Mass Mission to Apophis(99942) the same time offer wider launch windows, shown in Figure 15. They are also plotted in Figure 13. Depending on the type of mission, it is possible to lower the mass for a flyby of asteroid 2 SG 344 to 87 tons. total mission duration [days] 35 3 25 2 15 1 5 2 SG 344 28 EA 9 28 EA 9 29 GP 2 28 ST min. LEO Mass [tons] 215/1/1 216/12/31 219/1/1 22/12/31 223/1/1 224/12/31 227/1/1 228/12/31 231/1/1 232/12/31 235/1/1 departure date Figure 15: Minimum M LEO for all available targets up to H = 28.. As with Apophis, an example of a minimum mass mission to 2 SG 344 is shown in Figure 16. The mission profile s parameters are listed in Table 7 Line 12, and the resulting module masses can be seen in Table 6. Table 6: Module Masses for the Lowest-Mass Mission to 2 SG 344 Module EM RM HM PM EDS Total Mass [tons] 3.6 23.2 1.2 5.1 86.8 6 55 5 45 4 35 3 25 2 15 Module EM RM HM PM EDS Total Mass [tons] 5.75 3.6 33.8 66.1 182. 291.3 Y Axis [AU] to Sun Y Axis [AU] 228 9 9 228 1 9 228 11 8 Earth 99942 Spacecraft Trajectory Free Return Trajectory Moon Orbit 228 12 8 229 1 7 229 2 6 229 3 8 228 4 2 229 3 28 228 5 2 228 8 1 228 11 8 228 6 1 228 7 11 228 12 8 229 1 7 228 1 9 228 9 9 228 8 1 228 7 31 228 7 1 Earth 2 SG 344 Spacecraft Trajectory Free Return Trajectory Moon Orbit to Sun 228 5 12 228 6 11 229 4 27 229 2 6 229 1 27 Figure 16: Trajectory plot for a minimum mass mission to 2 SG 344 229 4 7 229 3 8 Figure 14: Trajectory plot for a minimum mass mission to Apophis 6.2 Smaller Target Options If we look at the entire range of asteroid sizes, we get a number of additional options that also have significantly lower mass requirements and at 6.3 Conclusion In conclusion, if we look at the entire NEO population and a limit for the total LEO mass of 37 tons, we have around 4 asteroids remaining as possible mission targets. Out of these, only 2 have a magnitude H 21.5. A selection of the best available options is shown in Table 7, sorted by the asteroid s size. Both the flyby and sample-return setup 9

Table 7: Minimum mass mission profiles for a selection of targets with low M LEO, sorted by absolute magnitude H. Both the flyby mission mass M flyby and the sample return mission mass M sr are listed. Line Target H M flyby M sr Departure dt mission dt 1 dt target dt 2 v total Name [ ] [kg] [kg] Date [days] [days] [days] [days] [km/s] 1 27945 19.29 278 33 414 522 18.5.227 355 18 15 16 7.53 2 99942 19.7 192 36 291 256 12.5.228 35 26 1 8 6.31 3 28 EV 5 19.97 236 833 352 849 12.6.224 365 22 5 14 6.89 4 27 SQ 6 21.9 23 818 33 769 6.1.223 365 12 35 21 6.89 5 27 UY 1 22.88 153 863 234 35 1.3.221 365 21 5 15 5.75 6 26 FH 36 22.92 178 542 269 989 19.8.234 355 27 5 8 6.13 7 29 OS 5 23.57 227 992 342 126 24.2.22 175 1 5 7 6.62 8 1999 AO 1 23.86 225 976 34 933 8.5.225 35 25 5 5 6.59 9 29 HC 24.77 126 12 192 729 5.1.225 365 1 5 26 5.9 1 2 SG 344 24.79 86 984 135 117 2.4.228 36 12 1 23 3.92 11 21 GP 2 26.88 9 472 141 268 27.1.219 345 28 5 6 4.4 12 27 XB 23 27.7 118 355 184 143 15.12.223 365 32 5 4 4.97 13 28 EA 9 27.74 118 536 18 791 18.8.219 355 15 5 2 4.82 mass requirements are included, making the additional costs between both mission types visible. We can see that options within a reasonable LEO mass are available for both mission setups. Since this study has been designed for an optimal habitation Volume, it is also possible to further lower the spacecraft s mass by reducing the living space of the crew to a level between the optimum and the performance limit. What we can also see is that the increase in mass requirements is not only caused by the larger crew size and additional equipment, but results from the less ideal trajectory profiles for larger asteroids in the analyzed NEO population. While this connection between smaller targets and lower v s is obvious for the data presented here, different time frames have the possibility of better accessible large targets. However, until after 235, our options for scientific missions are fairly restricted to these few targets. The launch windows for the most promising targets for this time-frame are shown in Table 8. Options of increasing the possibilities within this given time-frame are the discovery of new large NEOs or development of more advanced mission architectures that allow for lower in-leo mass requirements, than this Level 1 design. One of these options is described in Reference [11]. References [1] NASA Fiscal Year 211 Budget Estimates. NASA, 21. [2] Huntress, W., Stetson, D., Farquhar, R., Zimmerman, J., Clark, B., O Neil, W., Bourke, R., and Foing, B., The next steps in exploring deep space A cosmic study by the IAA, Acta Astronautica, Vol. 58, No. 6-7, 26, pp. 34 377. [3] JPL, N., Horizons Small-Body Database, 21. [4] Curtis, H., Orbital mechanics for engineering students, Butterworth-Heinemann, 25. [5] Noton, M., Spacecraft navigation and guidance, Advances in industrial control, Springer Verlag, 1998. [6] Walter, U., Astronautics, Wiley-VCH, Weinheim, 27. [7] Larson, W. and Pranke, L., editors, Human Spaceflight: Mission Analysis and Design (HS- MAD), McGraw-Hill, 1999. [8] Hofstetter, W., de Weck, O., and Crawley, E., Modular Building Blocks for Manned Spacecraft: A Case Study for Moon and Mars Landing Systems, Proceedings of the 15th Annual International Symposium, International Council on Systems Engineering, Rochester, New York, 25. [9] JAXA - International Manned Spacecraft Investigation Team, Manned Spacecraft System Investigation Report, Tech. rep., Japan 1

Table 8: Launch windows, showing the smallest and largest minimum in-leo mass per window (sample-return setup). Count represents the number of individual profiles, showing each window s flexibility. Target H Window Window Count MLEO,min MLEO,max Name [ ] Start End [kg] [kg] 99942 19.7 228-4-22 228-5-22 57 291 256 348 673 28 EV5 19.97 224-6-12 224-6-12 1 352 849 352 849 27 SQ6 21.9 223-1-6 223-1-6 6 33 769 33 769 27 UY1 22.88 221-3-1 221-3-1 32 234 35 234 35 26 FH36 22.92 22-8-12 22-9-1 32 33 867 347 115 26 FH36 22.92 234-3-12 234-3-22 81 284 554 35 831 26 FH36 22.92 234-8-9 234-9-8 7 269 989 343 72 29 OS5 23.57 22-2-24 22-2-24 14 342 126 342 126 1999 AO1 23.86 225-4-18 225-6-7 29 34 933 359 758 29 CV 24.25 215-7-3 215-8-19 246 241 13 248 743 29 CV 24.25 215-9-8 215-9-8 1 353 339 353 339 29 HC 24.77 225-9-25 225-1-5 111 192 729 3 59 29 HC 24.77 226-5-3 226-5-23 27 222 797 282 583 2 SG344 24.79 227-3-9 227-11-4 1836 15 262 291 54 2 SG344 24.79 227-11-24 227-12-14 424 276 854 283 541 2 SG344 24.79 228-1-23 228-4-2 6385 135 117 22 453 2 SG344 24.79 228-11-8 229-1-7 1434 143 611 329 357 2 SG344 24.79 229-8-25 229-9-4 2 295 41 39 829 21 GP2 26.88 219-3-11 219-3-21 147 218 886 274 535 21 GP2 26.88 219-9-27 219-12-16 176 141 268 271 35 21 GP2 26.88 22-1-31 22-11-1 68 28 851 292 626 27 XB23 27.7 223-12-15 224-1-14 28 184 143 334 236 27 XB23 27.7 224-6-2 224-6-12 183 26 83 279 961 27 XB23 27.7 224-12-9 224-12-9 48 252 523 252 523 28 ST 27.7 225-3-29 225-5-8 387 197 389 33 515 28 ST 27.7 225-11-14 225-11-24 95 24 846 355 132 28 EA9 27.74 218-1-12 219-5-3 16 22 736 336 237 28 EA9 27.74 219-6-19 219-8-18 3283 18 791 252 83 28 EA9 27.74 231-6-26 231-8-5 1362 182 251 288 819 Aerospace Exploration Agency (JAXA), March 21, Yuujin Uchuusen Shisutemu Kentou Houkokusho Kokusai Yuujin Uchuusen Kentou Chiimu. [1] Pravec, P., Harris, A., and Michalowski, T., Asteroids III, chap. Asteroid Rotations, University of Arizona Press, 22, pp. 113 122. [11] Farquhar, R., Dunham, D., Guo, Y., and McAdams, J., Utilization of libration points for human exploration in the Sun-Earth- Moon system and beyond, Acta Astronautica, Vol. 55, No. 3-9, 24, pp. 687 7. 11