Target selection and mass estimation for manned NEO exploration using a baseline mission design

Acta Astronautica 111 (2015) 198–221

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Target selection and mass estimation for manned NEO exploration using a baseline mission design Ralf C. Boden a,n, Andreas M. Hein a, Junichiro Kawaguchi b a b

Technische Universität München (TUM), Germany Japan Aerospace Exploration Agency (JAXA), Japan

a r t i c l e in f o

abstract

Article history: Received 14 January 2014 Received in revised form 12 January 2015 Accepted 17 February 2015 Available online 25 February 2015

In recent years Near-Earth Objects (NEOs) have received an increased amount of interest as a target for human exploration. NEOs offer scientifically interesting targets, and at the same time function as a stepping stone for achieving future Mars missions. The aim of this research is to identify promising targets from the large number of known NEOs that qualify for a manned sample-return mission with a maximum duration of one year. By developing a baseline mission design and a mass estimation model, mission opportunities are evaluated based on on-orbit mass requirements, safety considerations, and the properties of the potential targets. A selection of promising NEOs is presented and the effects of mission requirements and restrictions are discussed. Regarding safety aspects, the use of free-return trajectories provides the lowest on-orbit mass, when compared to an alternative design that uses system redundancies to ensure return of the spacecraft to Earth. It is discovered that, although a number of targets are accessible within the analysed time frame, no NEO offers both easy access and high incentive for its exploration. Under the discussed aspects a first human exploration mission going beyond the vicinity of Earth will require a trade off between targets that provide easy access and those that are of scientific interest. This lack of optimal mission opportunities can be seen in the small number of only 4 NEOs that meet all requirements for a sample-return mission and remain below an on-orbit mass of 500 metric Tons (mT). All of them require a mass between 315 and 492 mT. Even less ideal, smaller asteroids that are better accessible require an on-orbit mass that exceeds the launch capability of future heavy lift vehicles (HLV) such as SLS by at least 30 mT. These mass requirements show that additional efforts are necessary to increase the number of available targets and reduce on-orbit mass requirements through advanced mission architectures. The need for on-orbit assembly also becomes apparent, as availability of a HLV alone does not provide sufficient payload capabilities for any manned mission targeting NEOs. & 2015 IAA. Published by Elsevier Ltd. All rights reserved.

Keywords: Asteroids Near Earth Objects Human space exploration Manned missions Mission design

1. Introduction 1.1. Role of NEOs for human exploration

n

Corresponding author. E-mail addresses: [email protected] (R.C. Boden), [email protected] (A.M. Hein), [email protected] (J. Kawaguchi). http://dx.doi.org/10.1016/j.actaastro.2015.02.018 0094-5765/& 2015 IAA. Published by Elsevier Ltd. All rights reserved.

While the main goal of human space exploration efforts remains directed towards Mars, recent years have shown a trend that has shifted the focus of earlier mission targets from the Moon to asteroids within close proximity to the

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1.1.1. Scientific interests A major factor hereby is the scientific interest in asteroids. Pristine samples from asteroids allow an insight into the creation and early stages of the solar system as well as its planets. This makes NEOs an attractive target for both robotic and human sample-return missions. The presence of a human crew can provide flexibility that is still not achievable with robotic missions, and increase both quantity and quality of the returned samples [3]. Examples of robotic NEO exploration sample-return missions are JAXA's Hayabusa-I, -II, NASA's OSIRIS-REx, and ESA's Marco-Polo-R missions [4–7]. Studies of manned missions that consider NEOs as targets for manned missions can be found in [8,3]. 1.1.2. Bridging the gap to Mars NEOs are defined as asteroids (and comets) with a perihelion distance q of less than 1.3 AU,1 and no general restriction to their semi-major axis a [9]. Their orbits span the entire region between Earth (1 AU) and Mars (1.52 AU) (Fig. 1). NEOs can be grouped into different classes based on their orbital [10] and spectral characteristics [11]. Because of their distribution within the solar system, NEOs provide a bridge between current human spaceflight, centred around Earth, and interplanetary missions towards Mars. A spacecraft travelling to and from a NEO is exposed to the same deep-space environment and transfer times that occur during Mars transit. At the same time, total mission duration and delta-v requirements remain below those of a complete Mars mission. The minimum delta-v for a round trip mission to Mars (Mars-orbit) is approximately 6.3 km/s, with a total mission duration of 914 days (375 days spent in transit) [12, p. 22]. For NEO round-trip missions, both a mission duration under one year and delta-v requirements below 4.00 km/s can be achieved [2,13]. Due to the microgravity environment around asteroids, no ascent/descent vehicles or surface habitats are needed to conduct proximity operations such as sample collection. NEOs therefore provide an opportunity for manned missions beyond Earth's sphere of influence (SOI) at lower mission cost, schedule and complexity. At the same time they allow development and testing of 1 Astronomical Unit, equal to the mean Sun-Earth distance; 1 AU¼ 149 597 871 km.

2 1.5 1 0.5 Y Axis [AU]

Earth – so called Near-Earth Objects (NEOs). In terms of delta-v requirements, NEOs have the potential to be better accessible than the Moon. Compared to the 5 km/s required for a round-trip mission to Lunar orbit (without landing) [1], a round-trip mission to a NEO can be achieved with as little as 4 km/s [2]. Furthermore, the large size of the NEO population, with 8049 objects discovered as of May 2011 indicates that many promising mission opportunities to different NEOs should exist. While easy access plays a major part in the increased interest of NEO missions, it does not directly qualify NEOs to be considered as targets for future exploration. Especially for manned missions, the immense costs associated with these makes it necessary for a target to provide numerous incentives for its exploration.

199

0 −0.5 −1 −1.5 −2 −2

−1

0

1

2

X Axis [AU] Fig. 1. Heliocentric view of the location of all known NEOs within 2 AU around the Sun (October 2014).

technologies that are needed for future transfers to and from Mars. 1.2. Research goals and approach The aim of this paper is to identify promising NEO candidates for human exploration missions in the nearfuture. To achieve this, a baseline mission design is defined to analyse mission opportunities departing from a Low-Earth Orbit (LEO) in the 20 year time-frame between 2015 and 2035. The 2015 start allows checking if any promising opportunities are missed until the first mission can be realized; the year 2035 coincides with the possible start of human Mars exploration efforts [12]. The effect of target properties on the availability of mission opportunities is investigated. Specifically NEO size and rotation are investigated, as these two properties determine whether sample-return missions are feasible for specific NEOs or not (Section 2). Transfer times and delta-v requirements for individual mission profiles between Earth and the known NEO population are calculated using established methods [2,13]. A mass estimation model is developed, and used to determine the on-orbit masses associated with these NEO missions. The mass estimation model is based on existing technologies to provide a realistic mass estimate for missions taking place in the near future. In addition to providing an overview of available mission opportunities and target candidates, the baseline mission design can be used as a basis for evaluating more advanced mission concepts which are still being developed. 2. NEO population and asteroid properties Before discussing the baseline mission design and requirements, an overview of the NEO population used in this

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study is presented. NEO properties relevant to human exploration missions are also discussed.

2.1. Overview of the NEO population The entire NEO population is expected to contain as much as 20 50073000 asteroids with a diameter 4100 m [14]. As of 25 October 2014, a total of 11 604 NEOs have been discovered, including objects with diameters of only a few meters [15]. As the basis for this study, the list of 8049 NEOs from May 2011 will be used to determine if any of these objects provide good access for future human exploration missions. To keep the initial population, and therefore the number of potential targets as large as possible, no restrictions based on orbit uncertainty [16] are hereby imposed on this study.

2.2. Asteroid rotation and size Proximity operations of a manned spacecraft in the vicinity of an NEO make it necessary to select asteroids with a low spin rate to ensure crew safety. This applies in particular to sample-return type missions. In this study, a rotation period slower than 1–2 h is assumed as adequate to safely conduct proximity and surface operations. Only limited data on the rotation rates of asteroids exists [17]. A closer look at the size and composition of NEOs is therefore necessary to see if an upper limit for the rotation rate can be determined. NEO size is usually provided as absolute magnitude H [18]. H can be used to determine the diameter D of an NEO, based on the objects albedo α (avg. for NEOs: E0.154) [19]. Asteroids above a certain size are mainly composed of a “rubble pile” structure, with no tensile strength acting on its particles [20]. This effect can be seen in the distribution of known NEO size and rotation rates (Fig. 2) [17]. Rubble pile asteroids can only exist if their rotation rate T rot does not surpass the threshold where rotational forces exceed the gravitational forces holding the asteroid together. For the available light curve data shown in Fig. 2, asteroids with a H r21:5 (D4 200 m) have a 98% chance of a rotation period below 2 h and 99% to rotate slower than 1 h [17]. Based on the available lightcurve data, objects

with H 4 21:5 have only a 32% chance to rotate slower than 1 h (23% for 2 h). Besides being an indicator of an asteroid's rotation rate, size also determines whether an object should be considered as a target for human exploration missions. Smaller objects generally offer less incentive for exploration and may not provide enough reason to justify the high costs of sending a manned spacecraft to an asteroid if it is considered too small. While large objects are generally preferred, some missions may actually require smaller targets. One example for this case is asteroid redirect missions, which require an increased amount of resources (fuel) to successfully redirect larger/ heavier objects [21]. 3. A baseline mission design To provide a first basic mission architecture a single spacecraft will be used to transport crew, supplies, and equipment to the target asteroid from LEO. Similarities with the Apollo Lunar missions exist, which have so far been the only experience with manned missions beyond LEO. These provide a good starting point for developing a mission design that can serve as baseline for human spaceflight that reaches out beyond current operations around Earth. Like on Apollo, a 3 person crew is chosen for the baseline NEO mission. This provides sufficient manpower during the mission and allows 2 person EVA at the NEO. During proximity and surface operations at the NEO 100 kg of samples are expected to be collected, based on the amount of samples collected during later Apollo missions (Apollo 17: Msample ¼ 111 kg) [22]. Total mission duration from Earth departure to return is limited to one year to keep the mission within a reasonable time frame in regard to radiation exposure limits and general lack of experience with crewed missions of such an extended duration. The use of high thrust propulsion systems is expected to keep mission duration as short as possible. The baseline NEO mission will hereby rely on chemical propulsion systems, as no other existing technologies currently provide a sufficient amount of thrust. In addition to fast transfer times chemical propulsion systems offer high reliability, having seen extensive use on both manned and unmanned space missions. This results in two-impulse transfer trajectories between Earth and the targeted NEO. 3.1. Mission profile

Rotation Period T [h]

0.01

21.5

0.1 1

2h

10 100 1000 10,000 32

30

28

26

24

22

20

18

16

14

12

10

8

absolute Magnitude H

Fig. 2. Available sizes and rotation rates of NEOs. The size and spin rate thresholds at 21.5 magnitudes ( E170 m), and 2 h are clearly visible.

The use of a single spacecraft and the two-impulse transfer trajectories results in a very straightforward mission profile for the baseline design (Fig. 3). An NEO mission profile consists of an outbound and inbound transfer trajectory, and a section on the same orbit as the target asteroid. During this time proximity and surface operations can be performed. These three legs each have their respective durations (dt1, dt t , dt2), summing up to the total mission duration dt mission . Each mission profile is defined by the unique combination of dt1, dt t , and dt2; the departure date T departure ; and the targeted NEO. A propulsive manoeuvre is required to enter and exit the transfer trajectories between Earth and the NEO. Because a ballistic re-entry during the Earth return is assumed, no

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Fig. 3. Mission Profile consisting of the outbound and inbound transfer trajectories, as well as the time spent in the vicinity of the NEO.

manoeuvre is expected to take place during that section. Therefore, three delta-v manoeuvres are performed during the mission: Δv1 , Δv2 and Δv3 . The Earth departure manoeuvre (Δv1 ) takes place on the initial low-Earth parking orbit at a height of 400 km. This places the spacecraft at a similar orbit height as the ISS (416 km; 2014-10-25). The orbit height of the Apollo missions prior to trans-Lunar injection was lower, at E190 km. The total mass of E140 mT at this point consisted of the Apollo Spacecraft (Crew, Service, and Lunar Module) and the partially used S-IVB stage used for the trans-Lunar injection [22]. Due to the likely need for on-orbit assembly of the NEO spacecraft a higher orbit is chosen to reduce orbit degradation caused by increased atmospheric drag.2

3.2. Requirements and restrictions Radiation exposure and limited experience with longterm space missions limit mission duration dt mission to r370 days. 5 days are hereby expected as the minimum available time at the NEO dt t to carry out scientific experiments and collect samples. This is similar to the Apollo surface stays, which were limited to 2–4 days [22]. Longer staying times are favourable in lieu of the expected length of the mission, however, the effects on on-orbit mass requirements need to be taken into account (Section 8.2). Total delta-v (Δv1 þ Δv2 þ Δv3 ) is limited to 8.00 km/s, remaining below the 8.8 km/s required by the Apollo missions [24]. The velocity vEI at the entry interface (EI) during the ballistic Earth re-entry is restricted to r 12:00 km=s. The EI is defined as the point at which a spacecraft begins encountering the effects of Earth's atmosphere, at a height of approximately 130 km. The 12.00 km/s limit is recommended by other studies related to future human space 2 Orbit lifetime at 400 km: 77–1025 days; at 200 km: 1–6 days [23, p. 985].

exploration [12]. This value can also be expected on missions returning from an NEO (Hayabusa vEI : 12:10 km=s [25]). Table 1 provides an overview of all baseline mission restrictions to both the dt and the Δv parameters. 3.3. Mission profile calculation The method used to calculate the NEO mission profiles for this study is described by [2]. Similar calculation methods are provided by other sources like the Near-Earth Object Human Space Flight Accessible Targets Study (NHATS) [13]. The required delta-v in the heliocentric inertial frame, Δv11 –Δv41 , are calculated by solving Lambert's problem for the inbound and outbound transfer trajectories, based on the positions of Earth and the NEO, and the transfer time between these positions [26, p. 263]. Due to the low gravity environment, the required manoeuvres at the NEO (Δv2 , Δv3 ) are identical to Δv21 and Δv31 . For Earth departure and arrival, the patched conics method [27, p. 210] is used to connect the heliocentric transfer trajectories to those inside the Earth's sphere of influence (SOI).3 The delta-v for Earth departure from LEO (Δv1 ) and the re-entry velocity vEI at 130 km are calculated via the Oberth Manoeuvre from Δv11 and Δv41 respectively [27, p. 167]: qffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi Δv1 ¼ Δv211  2v2orbit 400  vorbit 400 ; qffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi vEI ¼ Δv241  2v2orbit 130 ð1Þ vorbit 400 is the orbital velocity on the initial parking orbit (7.67 km/s); vorbit 130 is the orbital velocity at the EI height (7.83 km/s). Using the described mission profile and parameters, Fig. 4 shows an overview of the algorithm used to calculate all NEO mission profiles. The time-frame, durations, and step-sizes used to compute the round-trip mission profiles in this study are listed 3

Radius of Earth's SOI rSOI ¼ 0:932  106 km [27, p. 212].

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Table 1 Overview of requirements and restrictions for the baseline mission. Total mission duration Time spent at NEO Total delta-v Velocity at EI

dtmission ( ¼ dt1 þdttarget þ dt2) dttarget Δvtotal ð ¼ Δv1 þ Δv2 þ Δv3 Þ vEI

r 370 days Z 5 days r 8:00 km=s r 12:00 km=s

Fig. 4. Flowchart showing the computation of the individual mission profiles for all objects of the NEO population.

in Table 2. The 370 day maximum mission duration is chosen to provide a common multiple of both the 10 and 5 day stepsize used in the computation. The 10 day steps keep computation costs low, while a smaller resolution of 5 days is used for the stay at the target. 3.4. Baseline safety considerations Safety considerations are made for the baseline mission design to ensure that the spacecraft is able to return to Earth after departing from its parking orbit. This issue is related to the successful execution of the Δv2 and Δv3 manoeuvres, both of which are needed to place the spacecraft onto its return trajectory. Failure to perform either of these delta-v manoeuvres is likely to result in mission failure and loss of crew. The reason for such a failure would be related to the spacecraft's propulsion system (Section 5). To avoid such a scenario, two options that can ensure a safe return to Earth are considered. These are discussed in Sections 3.4.1 and 3.4.2.

Other contingency scenarios related to the failure of spacecraft subsystems and components are considered in the spacecraft design (Section 4). These are addressed by providing sufficient system reliability and allowing for enroute servicing (tools, spare parts). 3.4.1. Free-return trajectories The use of a free-return trajectory allows the spacecraft to automatically return to Earth without the need for any additional propulsive manoeuvres4 after Δv1 . This is achieved by having the spacecraft depart on an outbound trajectory, which has an orbital period T orb of close to 365 days (71 day accuracy). After one full orbit on this trajectory, the spacecraft is automatically re-encounters Earth if no propulsive manoeuvres are performed. 4 Minor adjustments may be required but do not require large delta-v or high thrust.

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While the use of a free-return trajectory is possible without adding any additional systems to the spacecraft, this option does require designing the spacecraft for a mission duration of one year, independent of nominal mission duration. For short mission durations, this can lead to a significant increase in spacecraft mass due to increased spacecraft volume (structural mass) and large supply masses. When the nominal mission duration is already close to one year changes in spacecraft mass caused by the free-return option become insignificant. To remain consistent with the chosen parameters in Table 1 the free-return mission duration is also set to 370 days when determining spacecraft masses (Section 4). 3.4.2. Redundant propulsion modules Another option that ensures safe return in case of propulsion system failure is to add redundancies. For the baseline mission design, this is achieved by adding a second propulsion system to the spacecraft, capable of performing both the Δv2 and Δv3 . While this option allows retaining a nominal mission duration, the mass of the redundant propulsion system can become large due to high delta-v requirements. The original stage that performs the Δv2 and Δv3 manoeuvres also needs to be resized, to account for the added mass of the redundant propulsion system. This option therefore favours mission profiles where only a small amount of delta-v is needed for the manoeuvres at the NEO. 3.5. Overview of the baseline spacecraft Fig. 5 shows a schematic overview of the baseline spacecraft. This configuration does not necessarily represent the actual layout. It does however show all required modules

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considered in the baseline design. The spacecraft is divided into individual modules, which in combination provide the necessary functionality to successfully conduct an NEO sample-return mission. The separation into modules serves two purposes: individual components can be exchanged and rearranged based on how they are transported to LEO (and possibly assembled). This set-up also allows separating spacecraft components based on how they are affected by mission duration dt and Δv requirements of individual mission profiles. The baseline spacecraft's modules are hereby grouped into two distinct sections: The first section consists of the propulsion stages (EDS, PM). Delta-v manoeuvres are divided between these two modules (Section 5). The EDS is used to perform the Δv1 manoeuvre for the initial departure from the low-Earth parking orbit. The propulsive manoeuvres for NEO arrival and departure are both performed by the PM. If no freereturn trajectory is used a redundant PM will be provided as well (Section 3.4). The second section is referred to as the main spacecraft. This section consists of the three pressurized modules: CM, HM, and LM, as well as the unpressurized EM. The pressurized modules contain all systems, supplies and accommodations needed during the 3 person NEO mission transfer sections. Module masses are influenced by nominal mission duration dt mission ; the exception is the use of free-return trajectories, in which case spacecraft mass will be fixed for a 370 day mission duration. CM and HM provide the pressurized volume and accommodations for the crew, while the LM is used to store consumables such as water, air and other supplies. EM mass is generally constant, as it contains only the equipment and supplies needed to perform in situ experiments and collect samples at the NEO.

Table 2 Targets, dates and step-sizes used in the trajectory computation algorithm (Fig. 4) for the baseline study. Earth departure time-frame Step-size of departure date Tdeparture Transfer trajectory duration step-size dt1 and dt2 Step-size of time spent at target dttarget

Earth-Departure Stage (EDS)

2015-01-01–2035-01-01 10 days 10 days 5 days

Command Module (CM) & Heat-shield

Propulsion Module (PM) Propulsion Systems

Logistics Module (LM)

Habitation Module (HM)

Equipment Module (EM)

Main Spacecraft

Fig. 5. Schematic overview of the modular spacecraft configuration. While the exact layout may vary, all components considered in the baseline NEO design are included.

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4. Main spacecraft model and configuration This section discusses the systems and components of the main spacecraft, and how these are divided between the three pressurized modules CM, HM, and LM, as well as the unpressurized EM. The model used to determine the masses of each module and its components is explained in detail. Systems and components directly linked to the presence of a human crew are given special attention. These are influenced the most by variations in mission duration and display strong interactions, which will be discussed in detail. Details about the developed mass estimation model can be found in [28], on the basis of [29]. For other systems present in both manned and unmanned spacecraft, mass estimates available in general literature are taken and adapted for the requirements of a human NEO mission [30,23]. 4.1. Main spacecraft systems and components A schematic overview of the dependencies between the different systems and components of the main spacecraft is shown in Fig. 6. The individual systems and components are explained in detail in the following sections, including how their masses are estimated. 4.1.1. Life-support systems – ECLSS The environmental control and life-support system (ECLSS) is a key component in manned spacecraft. As such it influences a number of spacecraft components such as crew accommodations (Section 4.1.5) and supplies (Section 4.1.6). Power supply and thermal control systems are hereby also affected. The ECLSS mass depends on the chosen technology, which can

range from open-loop to fully regenerative systems; the latter usually having increased system masses and power requirements. For the baseline NEO mission design an open loopsystem is chosen [29]. This system offers low mass and power requirements as its task is limited to supplying air and water to the pressurized spacecraft volume, from their respective storage locations. The open-loop ECLSS offers high reliability, as no complex recuperation systems are used to recycle water and/or air. The supply mass for the open-loop system is therefore based on established daily requirements (Section 4.1.6). The ECLSS can be divided into two components, the environmental control system (ECS) and the life-support system (LSS). The ECS monitors and controls the environment (temperature, humidity) inside the pressurized sections of the spacecraft; components mainly consist of valves, fans, sensors and filters. The LSS controls and monitors the supply of air and water; components include ducts, valves, fans and sensors. ECS and LSS mass and power requirements are listed in Table 3. Spares for likely parts are included as well, based on [29].

4.1.2. Electrical power system – EPS The electrical power system (EPS) consists of components for the control, storage, and generation of electrical power. For the NEO mission exclusive use of photovoltaic cells for power generation is assumed. Power control/conversion system and battery masses are assumed constant (Table 4) [28, p. 20]. The system provides sufficient storage capacity for operation around Earth and the Moon, based on the source design [29]; no eclipses are expected during the NEO mission after LEO departure.

Fig. 6. Schematic of the main spacecraft systems and components, considered in the mass estimation model. Dependencies between individual systems are shown.

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Table 3 Mass and power requirements of the ECLSS including a 5% margin. Individual component masses provided in [28, p. 107]. Mass (kg) Environmental control (ECS) Life support (LSS) Spares 5% Margin Total (M ECLSS )

Power (W)

306 809 35 58

510 1675 – 110

1207

2295

Units

Mass (kg)

40.0 30.0 40.0 9.0 8.0 4.5 11.3 –

5 3 6 1 1 3 6 –

200.0 90.0 240.0 9.0 8.0 13.5 67.8 31.4

P SC Q_  η

659

ð2Þ

A typical blanket mass mblanket for solar arrays (including cover glass) is 2.05 kg/m2 [30, p. 650]. Additional structure accounts for 45% of the total solar array mass MSA : M SA ¼

ASP  mblanket 0:55

ð3Þ

A drive-train provides Sun-tracking for the solar arrays. Drive-train mass M DT is calculated according to [30, p. 650] as M DT ¼ ð  0:014M SA þ 20:6ÞM SA =100

ð4Þ

Total EPS mass M EPS is calculated as the sum of the system's components: M EPS ¼ M EPSconst þM SA þM DT

ð6Þ

The mass per surface area for a two sided radiator mrad can be estimated as 8.5 kg/m2 [30, p. 523]. The heat rejection rate Q_ rad for a radiator is assumed as 200 W/m2, based on the deep-space environment that the NEO spacecraft primarily operates in [30, p. 519]. The selected Q_ rad also considers LEO operation, although at reduced power requirements and/or slightly increased radiator temperature. Additional TCS components with constant mass and power requirements, are listed in Table 5 (including spares) [29]. TCS mass MTCS is the sum of the variable radiator mass and the constant component masses: M TCS ¼ M TCSconst þ Mradiator :

Solar panel size and mass is calculated based on the electrical power requirements of the NEO spacecraft P SC . Considering cell efficiency, packing factor, degradation and environmental/system losses [30, p. 651], a total efficiency η of 13% is expected. Based on the efficiency η and the solar power flux density Q_ of 592 W/m2 at the maximum expected distance of 1.52 AU5 from the Sun, the required solar panel surface ASP is calculated for P SC : ASP ¼

mrad Q_ rad

Mass (kg/unit)

Sum of constant EPS mass M EPSconst

waste heat P heat is assumed to be identical to the Power generated by the EPS P SC (Section 4.1.2). The TCS includes heaters, piping, valves, filters, sensors, pumps, heat exchangers and an external radiator. It can be divided into an internal loop and an external loop. The radiator mass Mrad is a function of P SC : M rad ¼ P heat 

Table 4 EPS storage and control systems. Main and auxiliary components are available.

Power Interface Unit (PIU) Power Control Unit (PCU) Lithium Battery (BAT) Auxiliary Battery (A-BAT) Auxiliary PIU (A-PIU) Power Interface Box (PIB) Power Conversion Unit (PCU) 5% Margin

205

ð5Þ

4.1.3. Thermal control system – TCS The thermal control system (TCS) is tasked with regulating the temperature of the spacecraft. Its main task is the removal of excess heat by radiating it into space. The 5 The NEO mission is expected to always remain within Sun–Mars distance of 1.52 AU.

ð7Þ

4.1.4. Fixed mass systems Additional systems present on the manned spacecraft are the data handling and command systems (DHCS), and the attitude determination and control systems (ADCS). Both systems have constant mass and power requirements, as they are not influenced by any mission parameters. Table 6 provides a breakdown of all three systems. The systems of the Mars reconnaissance orbiter (MRO) are used as reference for the DHCS of the NEO spacecraft [31]. This ensures that stable communications are provided for all NEO missions considered in the baseline design. ADCS components are based on common components used in attitude determination systems [30, p. 634]. Attitude control is provided by a reaction control system (RCS) identical to those found on existing spacecraft such as the ISS supply vehicles (HTV & ATV), and the Apollo CSM. A 28 thruster set-up is hereby assumed; a mass of 20 kg is allocated for each thruster (includes structure, piping, valves). Similar RCS masses are found in the HTV [29]. Fuel and tank structure is not included, as the amount of fuel for attitude adjustments is considered to be negligible. The PM (Section 5) is considered to provide sufficient margin to account for any attitude adjustments made by the RCS. 4.1.5. Crew accommodations Accommodations are designed to provide the crew with all necessities for the expected mission duration of up to one year. Table 7 list all accommodations, based on examples for long-duration missions in microgravity [30, p. 582]. 25% additional stowage mass/volume is added to account for fixations and racks. Total accommodation power consumption is 3000 W, based on the assumption that not all appliances are operated at the same time, which would require an additional 2100 W. Additional mass, power and volume is added for every crew member (Table 8) [30, p. 582]. This brings the total

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Table 5 TCS component mass and power requirements. The radiator mass is not included as it is a function P heat . Mass (kg)

Power (W)

Internal TCS loop External TCS loop (without radiator) Spares 5% Margin

379 262 44 34

247 565 – 41

Total constant TCS mass (M TCSconst )

720

853

Mass (kg)

Table 6 Breakdown of the ADCS and DHCS systems, including a 5% margin. Mass (kg)

Power (W)

ADCS: Inertial measurement units Sun sensors Star scanners/mappers Horizon scanners RCS (28  20 kg)

28.0 4.0 14.0 10.0 560.0

210 6 40 20 –

DHCS: X-bad components Ka-band components Misc. components Antennas Gimbals and drive motors

11.3 2.3 4.9 22.6 45.0

188 81 – – 14

35.1

28

5% Margin Total (M ADCS þ M DHCS )

737

Mass (kg) Volume (m3) Power (W)

Total MAccfixed

2883

1250 100 850 – 1000 400 1500 –

13.95

5100

accommodation mass MAcc to M Acc ¼ M Accfixed þ Ncrew  M Acccrew

Power (W)

2 2 20 59

0.014 0.005 0.002 0.850

– – – 70

90

1.00

–

25% Stowage

43

0.468

–

216

2.34

70

Table 9 List of daily air, water, food, and other supplies for a single person. Air and water is stored outside the pressurized spacecraft volume and does therefore not have any volume requirements.

Table 7 Mass, volume, and power requirements of the NEO spacecraft accommodations used for the baseline mission. Maximum power requirements are lower than the 5100 W total since not everything needs to operate at the same time.

1.12 2.19 0.37 0.54 1.00 0.50 5.44 2.79

Volume (m3)

Cooking/eating supplies Personal hygiene kit Operational supplies Personal accommodations and Stowage EVA equipment

Total M Acccrew

587

Food preparation 175 Hygiene and waste collection 53 Housekeeping 163 Operational equipment 100 Maintenance equipment 300 Photo/Video equipment 120 Health-Care/Medical 1395 25% Stowage 577

Table 8 Additional accommodation mass, volume, and power added for every crew member.

ð8Þ

For the 3 person baseline NEO mission, accommodations have fixed mass, volume and power requirements (Section 4.3.1).

4.1.6. Supplies Supplies are required on a daily, per person basis. They can be divided into air, water, food and other consumables. Air and water supplies have no volume requirements as they are stored outside the pressurized spacecraft volume. They are supplied to the crew via accommodations (Section 4.1.5)

Mass (kg)

Volume (m3)

Potable water (drinking) Potable water (food prep.) Hygiene water Water storage tank Air Air storage tank Food (partially dehyd.) Food packaging Other: Waste collection Housekeeping/cleaning Hygiene supplies (wipes, etc.) Clothing

2.0 0.5 1.0 0.02 0.84 0.2 1.8 0.5 0.44 0.55 0.63 0.78

– – – – – – 0.006 0.002 0.001 0.004 0.006 0.013

Total (M supplies )

9.26

0.032

and the ECLSS (Section 4.1.1). Food and other consumables are stored inside the pressurized spacecraft volume. An overview of the mass and volume requirements of all supplies is listed in Table 9. Values for food and water are based on expected requirements [30, p. 582]. The distribution of water and food masses depends on the amount of dehydration; however, total mass remains identical, and the chosen values are recommended for longer mission durations [30, p. 585]. The baseline design does not provide a shower and instead relies on the use of wipes for hygiene purposes. A limited amount of hygiene water (1.0 kg) is assumed per person and day. No stowage mass/volume is added for these supplies, as tank masses and packaging are listed separately for these consumables. The disposal of used supplies is considered in the baseline design. Due to the dual use of water supplies for radiation shielding purposes, disposal of water supplies is only performed when water supplies exceed the amount required for radiation shielding (Section 4.1.8). 4.1.7. Pressurized spacecraft structure The mass of the main spacecraft's pressurized structures depends on the pressurized volume of the individual modules (CM, HM and LM). M struct ¼ r struct  V press

ð9Þ

R.C. Boden et al. / Acta Astronautica 111 (2015) 198–221

A volume-to-mass ratio r struct of 130 kg/m3 is used to convert between pressurized volume and structure mass, based on data from existing space structures [30, p. 357]. While lower volume-to-mass ratios can be achieved using inflatable structures (e.g. TransHab) [32], these systems are still awaiting validation, and are therefore not considered for in baseline mission design. The current mass-based radiation shielding model (Section 4.1.8) would also negate any structural mass savings, by requiring the same amount as dedicated shielding mass. Suggestions of the habitable space that should be provided for each crew member are shown in Fig. 7 [30, p. 149]. The different options range from a minimum tolerable limit, to what can be considered the optimal habitable volume. Between them lies the performance limit, which marks the limit below which crew performance is expected to decrease. The personal space available on a nuclear submarine is listed as an additional reference point [29]. For the baseline mission the optimum habitable volume in Fig. 7 is chosen as a conservative option for the long mission durations that are expected. This amount of space is still a factor E3 below what has been and is provided on long-term space missions (ISS, Mir, Skylab) [30, p. 404]. A constant volume of 15 m3 is hereby allocated to the CM, independent of mission duration (Section 4.2.1), while the remaining habitable space is provided by the HM (Section 4.2.2). This way, the CM provides sufficient volume to remain just above the tolerable limit for a 3 person long duration mission. Pressurized volume for mission supplies is provided separately by the LM (Section 4.2.3). Water and air are stored in external tanks and therefore do not add to the pressurized volume of the LM. The habitable volume is substantially smaller than the pressurized volume inside a spacecraft, since spacecraft systems and accommodations occupy part of it. When the occupied volume is not known, a linear scaling factor sf be used to convert habitable to pressurized volume: V press ¼ sf  V hab

ð10Þ

Normally sf is taken to be 3 [33], however, since accommodation volume is defined for the NEO spacecraft (Section 4.1.5), a reduced sf of 1.8 is used, derived from the internal volume of ISS modules, and the pressurized sections of the ATV and HTV [28, p. 119].

Total Habitable Volume per 3 Crewmember [m ]

25 20

Optimum

15 Submarine (fixed and variable)

10 Performance Limit Tolerable Limit (fixed and variable)

5 0

0

1

2

3

4

5

6

7

8

9

10

11

12

Mission Duration [months]

Fig. 7. Habitable volume per crew member depending on mission duration. The optimum value is used as a reference for the baseline NEO mission design [30, p. 149].

207

Additional structural masses are the docking adapters needed to connect the individual modules of the main spacecraft. A mass of 350 kg is assumed for each individual adapter,6 based on the international berthing docking mechanism (IBDM) [34]. To perform EVA during the NEO mission, an airlock (AL) is required. Its mass of 440 kg, based on the space shuttle's internal airlock [30, p. 709], can be added to the structural mass of the main spacecraft when radiation shielding mass requirements are considered (Section 4.1.8). 4.1.8. Radiation shielding To reduce crew exposure to deep-space radiation during the NEO mission, radiation shielding is implemented. In order to provide sufficient shielding to remain within yearly exposure limits, a shielding mass mrad of 20–25 g/cm2 for the external surface of the spacecraft is suggested [30, p. 114]. For the baseline NEO mission, a shielding mass of 20 g/cm3 is assumed. The surface area Arad in need of shielding is estimated from the pressurized volume of the CM and HM by using a simplified model to calculate the outer surface of the spacecraft (Fig. 8): Arad ¼ 2π 

d V press 4  ¼ ^  V press 2 π  d22 d

ð11Þ

2

A spacecraft diameter d of 4 m is used in this calculation. Similar diameters are commonly found in both spacecraft [22] and space structures [30, p. 356]. The end caps of the cylinder are not considered as part of the surface requiring radiation shielding. Other spacecraft modules will be located at these points, providing sufficient shielding mass. The volume used for the storage of supplies (LM) is excluded from radiation shielding requirements, since the crew is expected to spend most of the time inside the habitable space of the CM and HM. Both structural masses M struct and available water supplies M water are subtracted from the required radiation shielding mass: M rad ¼ mrad  Arad  M sturct  M water

ð12Þ

What remains is the dedicated radiation shielding mass M rad , which is added to the CM and HM. Water supplies are first distributed to the CM until it requires no dedicated shielding – then to the HM. This helps reducing CM mass during the return to Earth, and with it the mass of the heatshield (Section 4.1.9), as water can be easily disposed of prior to entering Earth's atmosphere, while shielding mass remains present during the re-entry phase. Fig. 9 shows the dedicated radiation shielding mass based on total mission duration dt mission . Initially, its progression is dominated by the increase in pressurized volume according to Fig. 7 and (10). Once the maximum volume is reached shielding mass decreases linearly based on the increase in water supplies; first for the CM, then for the HM. On missions with flight times of up to one year a shielding factor of 20 g/cm2 may not provide sufficient protection during a solar particle event (SPE) [30, p. 114]. 6 To form a connection between modules, two IBDMs are required; one on each module.

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CM

longer mission durations due to increased water supplies (Section 4.1.8). This decrease stops once the radiation shielding requirements of the CM are covered entirely by structure and water. The collection of samples at the target asteroid also affects heat-shield mass; for the 100 kg of samples considered in the baseline design the mass increase of the heat-shield is only 20 kg however. This becomes more critical when larger masses than those considered in the baseline mission are being returned. For the specific design of the CM (Section 4.2.1), MHS lies at E2600 kg.

HM

Cylinder as Approximation of entire Spacecraft Volume

dedicated shield spacecraft structure water supplies

4m

4m l

4.2. Main spacecraft configuration

Fig. 8. Simplified model for estimating the external HM and CM surface area.

dedicated rad. shielding mass [kg]

7000 CM − dedicated shielding HM − dedicated shielding

6000 5000 4000 3000 2000 1000 0 0

50

100

150

200

250

300

350

400

450

500

total mission duration dtmission [days]

Fig. 9. Mass progression of dedicated radiation shielding for both CM and HM. Its progression is influenced by the linear increase in water supplies and the increase in pressurized volume based on Fig. 7.

While no additional shielding concept is provided for the occurrence of an SPE in the current model, distribution of shielding mass on the spacecraft can be investigated to provide a higher shielded shelter for these types of events. If and how much additional shielding mass would be required for such a shelter, is one point for future investigation (Section 9). In general, this mass based model is a first attempt to include shielding requirements in the mass estimation model and use available mass such as water and spacecraft structure for shielding. One remaining problem with this approach is the different shielding properties of the materials, as currently only mass is taken into account. Detailed simulation of these effects, including the investigation of secondary cascades are one major point suggested for future studies in Section 9. 4.1.9. Thermal shielding Thermal shielding is needed for the atmospheric reentry of the NEO spacecraft during its return to Earth. Heat-shield mass M HS is hereby assumed to scale linearly with the mass of the spacecraft re-entering the atmosphere mass M return by a conservative factor of 0.2 [33,30, p. 397]: M HS ¼ 0:2  M return

ð13Þ

Heat-shield mass initially decreases with mission duration as less dedicated radiation shielding is required for

The main spacecraft consists of four separate modules. With exception of the EM (Section 4.1.1), each of these modules provides some functionality based on the previously discussed systems and components. 4.2.1. Command Module – CM The Command Module (CM) is considered the central module of the main NEO spacecraft. It hold all of the spacecraft systems and accommodations (Sections 4.1.1– 4.1.5), and provides a radiation shielded habitable volume of 15 m3. For a three person crew, this volume represents the fixed tolerable limit of 5 m3 shown in Fig. 7. The CM is equipped with a single IBDM (NIBDM ¼ 1) to allow docking of other modules. For the baseline mission, the CM is expected to be equipped with a heat-shield for Earth reentry (Section 4.1.9). This makes it possible to reuse the CM for multiple missions, similar to other spacecraft designs such as the Orion Crew exploration Vehicle (CEV) and the SpaceX Dragon. The CM mass M CM is calculated as the sum of all its components: M CM ¼ M ECLSS þM EPS þM TCS þ ðM ADCS þM DHCS Þ þ M Acc þM structCM þ ðN IBDM  M IBDM Þ þM AL þ MradCM þ M HS ð14Þ More advanced mission designs may consider parking part of the spacecraft in the vicinity of Earth in between missions, and reuse these components over multiple missions. Such designs use additional manoeuvres and can require both advanced trajectory designs and additional equipment/fuel to place the spacecraft at its parking position near Earth. Therefore, these options are not considered in the baseline mission design, but are on option for more advanced mission concepts (Section 8.4.3). 4.2.2. Habitation Module – HM The HM provides additional habitable space to extend the 15 m3 of the CM to the chosen optimum (Fig. 7). Dividing the habitable volume between two modules allows keeping CM mass low for the Earth return phase, while still providing the required amount of space for the crew. Two IBDMs (NIBDM ¼ 2) on the HM allow connecting to the CM and having one additional docking point available. Like the CM, radiation shielding is required for the pressurized volume of the HM. HM mass M HM is calculated as M HM ¼ MHMsturct þ M HMrad þ ðN IBDM  M IBDM Þ

ð15Þ

R.C. Boden et al. / Acta Astronautica 111 (2015) 198–221

While crew accommodations are considered as part of the CM mass, relocating them to the HM can be an option similar to the use of the ISS rack system. 4.2.3. Logistics Module – LM The LM is a partially pressurized module used to store the supplies needed during the NEO mission. The pressurized volume of the LM is directly based on the volume requirements of the supplies (Table 9). The use of a dedicated LM allows separating the pressurized volume requiring shielding from the volume that can remain unshielded. As the LM only serves as storage space, no radiation shielding is considered for this module. While water is stored in the CM and HM structure during the mission for shielding purposes, the LM is considered as the initial storage location. By carrying all supplies like this, the LM can function as both a storage and resupply module. This provides increased flexibility when considering on-orbit assembly and possible resupply of the NEO spacecraft. LM mass M LM is the sum of the supplies listed in Table 9 and the dry mass M LMdry , consisting of the structural mass and two IBDMs (NIBDM ¼ 2) to allow attachment to the other main spacecraft modules: M LM ¼ M supplies  dt m Ncrew þM LMstruct þ ðNIBDM  MIBDM Þ

ð16Þ

4.2.4. Equipment Module – EM Equipment used at the NEO is transported in the EM. The transported equipment can range from nothing (flyby missions) to heavy equipment used for in situ resource utilization (ISRU), PHA7 mitigation, or asteroid mining. The exact payload mass will depend on the exact mission objectives. A EM mass of 2000 kg is assumed for the sample-return mission. It includes supplies and mobility aids for conducting EVA and equipment for proximity-surface experiments, including sample collection. The scientific equipment mass is assumed to be 500 kg, comparable with the equipment mass of the later Apollo missions ( E370 kg). These consisted of the lunar rover vehicle (LRV), Apollo Lunar surface experiment package (ALSEP), and the modularized equipment stowage assembly (MESA) [35–37]. Studies of future Lunar missions also suggest equipment masses of 500 kg [38]. Consumables include supplies for EVA suits as well as additional mobility units (manned manoeuvring units (MMU)) [30, pp. 709, 723]. Supply masses are based on a maximum of 10 2-person EVAs, expected to take place at the NEO. The EM itself is unpressurized, identical to the MESA of the Apollo Lunar Modules [37]. EVA supplies can be accessed externally or supplied through the docking ports, similar to the distribution system of the ISS via the common berthing mechanism (CBM) [39]. Table 10 provides a detailed breakdown of all EM masses. The EM's structural mass is estimated as 25% of the 7 Potentially hazardous asteroid: min. Earth orbit intersection distance r 0:05 AU; abs. magnitude H r 22:0.

209

Table 10 Breakdown of EM masses used for the baseline sample-return mission.

Science equipment MMU EVA supplies: LiOH O2 H2O MMU fuel EM structure (25%) docking port (IBDM)

Mass per unit (kg)

Units

500 150 2.9 0.63 4.5 12

1 2 (þ 1) 10  2 10  2 10  2 10  2

350

1

Total EM mass M EM

Mass (kg) 500 450 58 13 90 240 338 350 2039

equipment and supply mass. One IBDM is added to allow attaching the EM to the other modules of the main spacecraft. 4.3. Total mass of the main spacecraft The total mass of the main spacecraft and its components at the start of the mission is shown in Fig. 10, based on actual mission duration dt m . Fig. 10 shows that a number of systems and components have constant masses (ECLSS, Sci, ADCS, DHCS, TCS, EPS, Accommodations, CM structure, EM). Other components increase linearly with mission duration (supplies, LM structure). A linear decrease is observed for both the dedicated CM radiation shielding (Fig. 7) and heat-shielding (13). More complex dependencies are seen for the HM structure, which depends on the habitable volume requirements (Fig. 7), and its dedicated radiation shielding (Fig. 8). 4.3.1. Mass progression Four points during the mission are of interest, mainly for calculating propulsion system masses (Section 5), and also HS mass: 1. 2. 3. 4.

On-orbit; prior to Δv1 (dt ¼0). NEO arrival; prior to Δv2 (dt ¼dt1). NEO departure; prior to Δv3 (dt ¼dt1 þ dt t ). Earth return; when entering Earth atmosphere (dt¼ dt1 þ dt t þ dt2 ¼dt m ).

The main spacecraft mass M M at these points are onorbit: NEO arrival:

M Mð1Þ ¼ M CM þ M HM þM LM þ M EM M Mð2Þ ¼ M Mð1Þ  M supplies  dt 1  N crew

NEO departure:  dt t  Ncrew

ð17Þ ð18Þ

M Mð3Þ ¼ M Mð2Þ þ Msample  M EM  M supplies ð19Þ

Earth return: MMð4Þ ¼ M Mð3Þ M HM  MLMdry  Msupplies  dt 2  Ncrew ð20Þ

4.3.2. Fixed spacecraft mass for the free-return option When considering the use of free-return trajectories (Section 3.4.1), the mass of the main spacecraft is fixed to a mission duration of 370 days, independent of nominal

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engine burn in under Δt burn of 1000 s:

mission duration. The total mass of the main spacecraft for this case is listed in Table 11, including a breakdown of the CM, HM, and LM masses (EM available in Table 10).

F ¼ M0 

The propulsion system of the baseline NEO mission consists of two separate Modules. The Earth departure stage (EDS) is used for the Δv1 manoeuvre at the start of the mission. The propulsion module (PM) is used for the Δv2 and Δv3 manoeuvres at NEO arrival and departure. A delta-v margin of 0.1 km/s is added to both Δv2 and Δv3 to account for any orbit and attitude corrections during the mission. The required amount of fuel for any delta-v manoeuvre can be determined by Tsiolkovsky's rocket equation [27, p. 20]:   M fuel ¼ M 0  1  e  Δv=ðg0 ISP Þ ð21Þ

M EDS ¼ M EDSeng þ MEDSstruct þ Mfuel1

ð22Þ

Based on the expected size difference and different technologies (fuels), a 10% structure ratio is assumed for the EDS, and a 15% structure ratio for the PM. Engine mass M eng is calculated from the engine's thrust-to-weight ratio (TWR) [30, p. 788]. F ¼ 0:0006098F þ 13:44 M eng

ð25Þ

Table 11 Total mass of the main spacecraft for the free-return case. The total mass is rounded to 10 kg.

M0 is hereby the spacecraft mass prior to the manoeuvre. g0 is the standard earth acceleration (9.80665 m/s2) and I SP the specific impulse, defined by the type of propellant. The EDS is assumed to use LH/L02, which provides a high I SP of 450 s [23, p. 692]. A hypergolic combination of monomethyle-hydrazine (MMH) and nitrogen-tetroxide (NTO) is selected for the PM to avoid problems with long-term storage and engine re-ignition. The trade-off is a lower specific impulse of 340 s [23, p. 692]. The structural mass (tanks, plumbing) Mstruct can be estimated from the fuel mass, using a structure ratio r struct :

TWR ¼

ð24Þ

Since M fuel , M struct , and M eng are all part of the initial spacecraft mass M0, an iterative calculation of M0 needs to be performed (Fig. 11a and b). EDS and PM masses are calculated as the sum of their individual components:

5. Propulsion systems – EDS and PM

M struct ¼ r struct  M fuel

Δv Δt

ð23Þ

Symbol

Mass (kg)

Crew CM structure (Volume:  21 m3 ) Airlock Docking ports (  1) Spacecraft systems ECLSS EPS TCS DHCS ADCS Accommodations Dedicated CM shielding Thermal shielding Total CM

Mcrew MstructCM MAL 1  M IBDM MECLSS MEPS MTCS MDHCS MADCS MAcc MradCM MHS MCM

240 5796 440 350 1207 885 1015 90 647 3531 0 2731 16 932

HM structure (Volume:  86 m3 ) Docking ports (  2) Dedicated HM shielding Total HM

MstructHM 2  M IBDM MradHM MHM

10039 700 3828 14 567

LM structure (Volume:  36 m3 ) Docking ports (  2) Supplies (M supplies ): water Air Food ( þpackaging) Other Total LM

MstructLM 2  M IBDM Mwater Mair Mfood Mother MLM

4618 700 3885 1152 2553 2664 15 572

Total EM

MEM

Main Spacecraft (on-orbit)

MMð1Þ

2039 49 110

The required thrust F is chosen to allow completing the

sum of main spacecraft component masses [kg]

60,000 55,000 50,000

Crew Systems Accommodations Food & other supplies Air Water CM structure HM structure LM structure CM rad. shield HH rad. shield CM heat shield EM

45,000 40,000 35,000 30,000 25,000 20,000 15,000 10,000 5,000 0 0

50

100

150

200

250

300

350

400

450

500

mission duration dtm [days]

Fig. 10. Overview of the initial main spacecraft mass for the 3 person NEO sample-return mission, based on total mission duration dt m .

R.C. Boden et al. / Acta Astronautica 111 (2015) 198–221

211

Fig. 11. Algorithms used to calculate EDS (a) and PM (b). Use of the PM during two delta-v manoeuvres with a change in spacecraft mass requires a more complex approach.

M PM ¼ M PMeng þ MPMstruct þ M fuel2 þ M fuel3

ð26Þ

The main spacecraft mass changes during the mission due to disposal of used up consumables and equipment (EM) and addition of collected samples. The calculation of M M depends on the mission profile, as it is based on the time that the propulsive manoeuvre is performed (Section 4, (17), (18), (20)).

available under the given restrictions. This percentage complies with the results of other surveys, such as [40]. The minimum delta-v (Δv1 þ Δv2 þ Δv3 ) required for a round-trip mission for all available 162 NEOs in the 2015– 2035 time-frame is shown in Fig. 12, based on mission duration dt m . Nprofile is shown in Fig. 13, both as the sum of all NEOs, and for each individual NEO. An especially large number of available mission profiles (4 20 000) are observed between April and June 2028. The largest number of available profiles in this time-frame (38%), as well as in total (10%) is associated with the NEO 2000 SG344. A number of other NEOs are also accessible in the same time-frame, although with a much smaller Nprofile . One of these is the asteroid (99942) Apophis. Fig. 14 shows a histogram of the number of available targets based on the number of available mission profiles (Nprofile ).

6. Available NEO mission opportunities and on-orbit mass requirements

6.1. Effects of free-return trajectories

Based on the restrictions in Section 3 (Table 1) the availability of NEOs and opportunities for a manned sample-return mission are discussed. The number of available individual mission profiles (Nprofile ) is hereby used as an indicator of availability and accessibility. This number can indicate the existence of multiple and/or long launch windows, as well as low transfer times and delta-v/mass requirements within these windows. The effect of target size restrictions (Section 2.2) and the implementation of safety features (Section 3.4) on the availability of mission opportunities are also investigated. Computation of all mission profiles according to Section 3.3 shows that 162 (E2%) of the initial 8049 NEOs remain as potential mission candidates without requiring a free-return trajectory. 2 630 575 mission profiles are

Including use of free-return trajectories (Section 3.4.1) reduces the available mission opportunities, as it restricts the outgoing transfer trajectories of the mission profiles. As a result, the number of accessible NEOs decreases by approximately 27%, from 162 to 118 (Fig. 15a). More significant than this reduction of targets, however, is the decrease of available mission profiles by over one order of magnitude (Fig. 15b). Only 132 234 (E5%) of the 2 630 575 mission profiles provide the option of a free-return. In addition to restricting the number of mission profiles, free-return profiles require an increased amount of delta-v in some cases (Fig. 16). The average increase for the 118 accessible NEOs is 0.29 km/s. For 9 NEOs this increase exceeds 1.00 km/s. 2000 SG344 still remains the NEO with the largest number of individual mission profiles (33%).

The initial spacecraft mass M0 at the beginning of each delta-v manoeuvre consists of the main spacecraft mass M M and the respective propulsion systems: at these times, M 0 is calculated as onorbit: M 0ð1Þ ¼ M Mð1Þ þ M PM þ M EDS

ð27Þ

NEO arrival: M0ð2Þ ¼ M Mð2Þ þM PM

ð28Þ

NEO departure: M 0ð3Þ ¼ MMð3Þ þ M PM  M fuel2

ð29Þ

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8 350 7.5

7

6.5

250

6 200 5.5 150

5

minimum Δv [km/s]

total mission duration dtmission [days]

300

4.5 100 4 50 3.5

0 2015 2016 2017 2018 2019 2020 2021 2022 2023 2024 2025 2026 2027 2028 2029 2030 2031 2032 2033 2035

3

departure date

Fig. 12. Minimum delta-v requirements to access any of the 162 available target candidates. 25,000

number of available mission profiles (count)

22,500 20,000 17,500 15,000 12,500 10,000 7,500 5,000 2,500 0 2015 2016 2017 2018 2019 2020 2021 2022 2023 2024 2025 2026 2027 2028 2029 2030 2031 2032 2033 2035 departure date

Fig. 13. N profile , representing the available number of individual mission profiles for all targets. The black line represents the sum over all individual NEOs.

6.2. Restrictions for sample-return based on NEO properties NEO properties are viewed under the aspects of allowing sample-return operations. As a first step, only spin rate and size restrictions are applied; the latter as a means to estimate a maximum rotation rate (Section 2). The largest accessible NEO (207945) hereby has a magnitude of 19.18; the smallest NEO has a magnitude of 32.04. Spin rates are known only for a very limited selection of NEOs [17]. From the 118 target candidates, spin rates are only

known for 13 of them. 7 of them are fast rotators with T rot o1 h. 2006 RH120 is one such asteroid, with a period of only E3 min. This leaves 155 NEOs as potential candidates (111 in case of free-return). The number of mission profiles to these targets is 2 314 057 (120 989 in case of free-return). This accounts for 88% (5% in case of free-return) of all available mission profiles, calculated according to Section 3.3. The size restriction of H r21:5 results in 5 available targets (3 in case of free-return). The rotation rate of two of these NEOs is actually known (99942 and 2000 EV5).

500000

100000

50000

213

10000

5000

1000

500

100

50

10

5

15 14 13 12 11 10 9 8 7 6 5 4 3 2 1 0

1

number of available target NEOs

R.C. Boden et al. / Acta Astronautica 111 (2015) 198–221

number of available mission profiles Fig. 14. Histogram showing the number of available NEOs based on their N profile . Due to the wide range of N profile , a logarithmic scale is used for the x-axis.

sum of available mission profiles [count]

10,000,000 1,000,000 100,000 10,000 1,000 100 10 all available mission profiles free−return mission profiles

1 18

19

20

21

22

23

24

25

26

27

28

29

30

31

32

absolute magnitude [H]

Fig. 15. Number of available NEOs (a) and mission profiles (b) with a target of magnitude r H. A comparison between non-restricted and free-return cases is shown.

delta−∆ v (all vs. free−return) [km/s]

1.75 1.5 1.25 1 0.75 0.5 0.25 0 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 absolute magnitude H [−−] Fig. 16. Difference in minimum delta-v requirements between free-return and non-restricted mission profiles. All available NEOs are plotted over their absolute magnitude H.

Based on both the size and rotation rate restriction, a combined number of 9 NEOs (6 in case of free-return) have either a known T rot of Z 1 h or an absolute magnitude

r21:5 with unknown T rot . These targets allow samplereturn operations based either on their known spin rate, or their absolute magnitude (H r21:5). Fig. 17 shows the

R.C. Boden et al. / Acta Astronautica 111 (2015) 198–221

number of available mission profiles (count)

214

250 240 220 200 2001 QC34 2002 TZ66 2007 HL4 2008 EV5 2009 SH2 2010 EX11 2011 CG2 207945 99942

180 160 140 120 100 80 60 40 20 0 2015 2016 2017 2018 2019 2020 2021 2022 2023 2024 2025 2026 2027 2028 2029 2030 2031 2032 2033 2035 departure date

Fig. 17. Number of available mission profiles, restricted to NEOs that allow sample-return. Both the free-return (□) and non-restricted (◯) case is shown.

number of available mission opportunities to these targets, for both the unrestricted and free-return case. 8 out of these 9 are known slow rotators with T rot r 1–2 h. 2001 QC34 has a 4 99% chance of rotating slow enough for proximity and surface operations according to Section 2.2. 7. Baseline mission on-orbit mass requirements Based on the results of the previous section, we now estimate the on-orbit mass (M LEO ) of spacecraft for conducting these NEOs sample-return mission is shown, based on the developed mass estimation model in Sections 4 and 5. An upper limit of 500 mT is set for the on-orbit mass. This provides an upper limit without imposing too heavy restrictions on spacecraft masses. It is also similar to the 534 mT of a manned Mars spacecraft capable of a roundtrip mission to Mars orbit (no surface components) [12]. Both safety options of either utilizing free-return trajectories or carrying a redundant PM are investigated; one of these options is required to allow the spacecraft to return to Earth. 7.1. NEO selection All available targets will be considered in the on-orbit mass estimation, as long as they are not excluded by high rotation rates. This includes NEOs with a magnitude above 21.5 as well. While these targets may not fulfil the requirement for sample-return operations, some of these objects have significantly lower delta-v requirements [2]. Compared to objects with slow rotation rates and/or larger sizes (5.45–7.39 km/s), smaller NEOs with unknown spin rates can require as little as 3.80 km/s (2000 SG344). 12 promising target candidates are listed in Table 12, including their properties. The top 3 objects are the accessible NEOs which fulfil the size requirement of H r 21:5. All three, 207945, (99942) Apophis, and 2008 EV5, have slow rotation rates. 2000 H4 and 2009 SH2 are two smaller objects, but have a known rotation period which lies within the 1–2 h threshold (Section 2.2).

As these 5 NEOs all have very small Nprofile due to high delta-v requirements above 6.02 km/s and short launch windows, an additional 7 asteroids have been chosen, based on a combination of low delta-v and large Nprofile relative to their smaller size. These objects provide a means to compare the on-orbit mass requirements of NEOs that allow samplereturn, to those where such operations may not be possible, as these small objects have a 68–77% chance of rotating too fast to allow proximity operations (Section 2.2). 7.2. On-orbit mass requirements 7.2.1. Free-return trajectory option The minimum on-orbit mass for all NEOs that allow free-return type mission profiles is shown in Fig. 18. The number of target candidates that do not exceed 500 mT is 76 (64 026 mission profiles). The available targets and mission profiles based on on-orbit mass are shown in the histograms in Fig. 19a and b. For individual targets the on-orbit mass based on mission duration is shown in Fig. 20a for all 76 objects. Fig. 20b provides an overview of the 15 selected targets in Table 12; additional details are provided in Table 13. The fixed mass of the main spacecraft of 49.35 mT results in the preferred mission profiles having a duration close to one year (370 days). 7.2.2. Redundant PM option For a mission that uses a redundant PM instead of freereturn trajectories as safety feature, a larger number of potential targets and mission profiles exist. However, only 49 NEOs can be accessed with an on-orbit mass of less than 500 mT (276 918 mission profiles) using this redundant PM option. Out of these 49 objects, 47 overlap with those of the free-return option. Only the two NEOs 2001 BA16 and 2007 PS9 are accessible with under 500 mT only using the redundant PM option. However, both NEOs have unknown rotation rates, magnitudes above 23.5 and require an on-orbit mass greater than 400 mT, and are therefore not desirable targets.

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215

Table 12 List of properties for the selected target candidates. Both redundant PM (RED) and free-return (FR) options are listed. The Δv and dt m values for 207945 are listed as reference (n) even though the mass limit of 500 mT is exceeded. Name/ID

H

T rot (h)

Min. Δv (km/s)

Min. dt m (d)

(RED)

(FR)

(RED)

(FR)

(RED)

(FR)

(RED)

(FR)

N profile

Max. dt t (d)

207945 99942 2008 EV5

19.18 19.20 20.00

3.15 30.40 3.73

7.42n – –

6.34n 6.40 6.90

340n – –

335n 340 370

– – –

– 50 10

– – –

– 43 1

2007 HL4 2009 SH2

24.23 24.90

1.48 1.26

– 5.45

6.93 6.02

– 355

360 345

– 65

10 25

– 37

3 16

2006 FH36 2009 OS5 1999 AO10 2000 SG344 2001 GP2 2011 BL45 1991 VG

22.92 23.57 23.85 24.79 26.88 27.16 28.39

– – – – – – –

5.91 – – 3.80 4.14 4.42 4.09

6.14 6.65 6.52 3.92 4.14 4.42 4.14

325 – – 135 230 95 175

310 165 270 145 250 125 155

25 – – 165 145 120 145

95 5 15 185 155 130 180

33 – – 84 090 13 186 16 431 27 264

311 4 45 9087 1687 10 610 5811

500 350

total mission duration [days]

400 250

350 200

300

150

250

100

200

50

150

0 2015 2016 2017 2018 2019 2020 2021 2022 2023 2024 2025 2026 2027 2028 2029 2030 2031 2032 2033 2035

minimum on−orbit Mass [mT]

450 300

100

departure date

Fig. 18. Minimum on-orbit mass for missions using a free-return trajectory as safety feature. REDO GRAPH.

10 number of available mission profiles (count)

7000

9

number of available NEOs

8 7 6 5 4 3 2 1 0

6000 5000 4000 3000 2000 1000

100 125 150 175 200 225 250 275 300 325 350 375 400 425 450 475 500 525 550

0 100 125 150 175 200 225 250 275 300 325 350 375 400 425 450 475 500 525 550

on−orbit mass [mt]

on−orbit mass [mt]

Fig. 19. Histograms showing (a) number of accessible NEOs and (b) mission profiles based on on-orbit mass for the free-return option.

Fig. 21 shows that launch windows are generally larger for this option, without the abrupt cut-off observed for the free-return option (Fig. 18). Also indicated are opportunities with shorter mission duration; a reduction in on-orbit mass compared to the free-return option is however not observed.

In fact, for targets that allow sample-return operations, only 2009 SH2 remains available and with 40 mT higher on-orbit mass than the free-return type mission. The available targets and mission profiles based on onorbit mass are shown in the histograms in Fig. 22a and b.

R.C. Boden et al. / Acta Astronautica 111 (2015) 198–221

500 475 450 425 400 375 350 325 300 275 250 225 200 175 150 125 100

minimum on−orbit mass [mT]

minimum on−orbit mass [mT]

216

0

25

50

500 475 450 425 400 375 350 325 300 275 250 225 200 175 150 125 100

75 100 125 150 175 200 225 250 275 300 325 350370

207945* 99942 2008 EV5 2007 HL4 2009 SH2 2006 FH36 2009 OS5 1999 AO10 2000 SG344 2001 GP2 2011 BL45 1991 VG 0

25

50

75 100 125 150 175 200 225 250 275 300 325 350 370

mission duration [days]

mission duration [days]

Fig. 20. Minimum on-orbit mass over mission duration for the free-return option. NEOs where 500 mT of on-orbit mass are exceeded are marked by an n in (b). (a) All 76 NEOs. (b) Promising candidates.

Table 13 Mission profiles with minimum on-orbit mass requirements using free-return trajectories. The asteroid 207945 hereby exceeds 500 mT. Name/ID

T departure

dt m

dt t

Δv1

Δv2

Δv3

MM

M PM

M PMr

M EDS

M LEO

207945 99942 2008 EV5

2027-05-28 2028-05-12 2024-06-12

355 355 370

5 5 10

3.76 3.78 3.79

2.42 1.59 1.63

1.24 1.03 1.48

49 49 49

178 88 121

– – –

393 240 300

620 377 470

2007 HL4 2009 SH2

2032-10-28 2016-10-02

360 360

10 10

3.72 4.20

1.62 0.77

1.59 1.05

49 49

133 52

– –

310 214

492 315

2006 FH36 2009 OS5 1999 AO10 2000 SG344 2001 GP2 2011 BL45 1991 VG

2034-08-19 2020-02-24 2025-04-28 2028-04-02 2019-10-27 2030-02-11 2017-05-30

355 175 315 355 340 360 365

5 5 5 5 10 10 5

3.80 3.23 3.21 3.21 3.31 3.27 3.23

1.39 2.11 2.07 0.34 0.61 0.62 0.45

0.95 1.31 1.24 0.37 0.22 0.53 0.46

49 49 49 49 49 49 49

73 155 136 19 21 30 23

– – – – – – –

217 275 247 93 100 111 100

339 479 432 160 170 190 172

Fig. 22b shows that the use of a redundant PM results in a fast increase of on-orbit mass based on the distribution of available profiles over on-orbit mass requirements. A detailed view of the minimum on-orbit masses over mission duration for all 49 NEOs and the promising candidate selection is shown in Fig. 23a and b. Details of the minimum mass mission profiles are available in Table 14. While 207945 still exceeds 500 mT of on-orbit mass, the mass is almost 80 mT lower than for the free-return option. 7.3. Options for human NEO missions Comparing the minimum on-orbit mass requirements for the free-return and redundant PM options shows that the additional supply masses affect total on-orbit mass less than the addition of a redundant propulsion module (Fig. 24). Cases where the difference in on-orbit mass is only a few metric tons occur when the delta-v manoeuvres for NEO arrival and departure are small. One example for this is 2000 SG344, where Δv1 and Δv2 are both only around 0.35 km/s – compared to the 1.1–1.6 km/s for (99942) Apophis. These results indicate that a free-return mission to (99942) Apophis is the most likely scenario when all restrictions are taken into account. Its known rotation period of over 30 h is significantly below the speed limit of 1–2 h, making this asteroid a good option for proximity

and sample-return operations. However, the on-orbit mass requirements for conducting a manned mission to (99942) Apophis are in the upper region of the 500 mT limit, at 381 mT for the free-return case. Access to (99942) Apophis is also restricted by a small launch window and a very low number of available mission profiles. 2000 SG344 provides a more accessible target with both lower on-orbit mass requirements and a much larger Nprofile . While the on-orbit mass of 161 mT is still large, it is less than half of what is required for a mission to (99942) Apophis. This NEO is also accessible using both the free-return and redundant PM options. The free-return option is still the better choice, as on-orbit mass is still 15 mT lower, and the time spent at the NEO is identical. The main problem with 2000 SG344 is the uncertainty regarding proximity and samplereturn operations, as its high magnitude of 24.8 and unknown rotation rate indicate a 68–77% chance of it being a fast rotator. 8. Conclusion 8.1. Availability of promising targets and mission opportunities The previous sections indicate that, although the initial NEO population suggests a high chance of finding numerous promising targets, restrictions and high on-orbit mass

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217

500 350 450 300

350 200 300 150

250

100

200

50

150

0 2015 2016 2017 2018 2019 2020 2021 2022 2023 2024 2025 2026 2027 2028 2029 2030 2031 2032 2033 2035

minimum on−orbit Mass [mT]

total mission duration [days]

400 250

100

departure date

Fig. 21. Minimum on-orbit mass of missions using a redundant PM instead of a free-return trajectory. REDO GRAPH.

10 number of available mission profiles (count)

45,000

9 number of available NEOs

8 7 6 5 4 3 2 1

40,000 35,000 30,000 25,000 20,000 15,000 10,000 5,000

0 100 125 150 175 200 225 250 275 300 325 350 375 400 425 450 475 500 525 550

0 100 125 150 175 200 225 250 275 300 325 350 375 400 425 450 475 500 525 550

on−orbit mass [mt]

on−orbit mass [mt]

500 475 450 425 400 375 350 325 300 275 250 225 200 175 150 125 100

minimum on−orbit mass [mT]

minimum on−orbit mass [mT]

Fig. 22. Histograms showing (a) number of accessible NEOs and (b) mission profiles based on on-orbit mass for the redundant PM option.

0

25

50

75 100 125 150 175 200 225 250 275 300 325 350370 mission duration [days]

500 475 450 425 400 375 350 325 300 275 250 225 200 175 150 125 100

207945* 99942* 2008 EV5* 2007 HL4* 2009 SH2 2006 FH36 2009 OS5* 1999 AO10* 2000 SG344 2001 GP2 2011 BL45 1991 VG 0

25

50

75 100 125 150 175 200 225 250 275 300 325 350370 mission duration [days]

Fig. 23. Minimum on-orbit mass over mission duration for the redundant PM option. NEOs where 500 mT of on-orbit mass are exceeded are marked by an n in (b). (a) All 49 NEOs. (b) Promising candidates.

requirement limit the number of mission candidates in the near future. Under the given restrictions to delta-v and mission duration (Section 3.2), safety considerations (Section 3.4), target properties (Section 2.2), only 9 out of the 8049 analysed NEOs qualify as potential targets for samplereturn operations (Section 6.1). Only four of these are accessible with an on-orbit mass below 500 mT, calculated

using the developed mass estimation model (Section 4), with the lowest on-orbit masses still above 300 mT. One of these potential targets is (99942) Apophis; a well known object due to its upcoming close approach to Earth in 2029 [15]. The on-orbit mass for a manned mission to this NEO lies at a minimum of 380 mT. Despite this, (99942) Apophis is one of the most promising targets for a sample-return mission due to its slow rotation rate of 30.4 h and magnitude

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Table 14 Mission profiles with minimum on-orbit mass requirements using a redundant PM. On orbit masses are all above those for the free-return option (Table 13). An increased number of missions exceed the 500 mT limit. Name/ID

T departure

dt m

dt t

Δv1

Δv2

Δv3

MM

MPM

MPMr

M EDS

M LEO

207945 99942 2008 EV5

2027-05-18 2028-05-02 2024-06-22

365 360 370

55 10 60

4.16 3.81 4.37

0.94 1.46 1.00

1.24 1.11 1.00

49 49 49

68 83 59

59 73 52

364 363 364

540 568 525

2007 HL4 2009 SH2

2019-04-20 2015-09-28

365 365

5 5

3.77 3.98

1.84 0.54

1.10 0.93

49 49

106 39

94 33

431 233

680 354

2006 FH36 2009 OS5 1999 AO10 2000 SG344 2001 GP2 2011 BL45 1991 VG

2034-09-08 2019-08-08 2025-08-06 2028-05-02 2019-10-27 2030-02-11 2017-06-09

335 365 215 315 340 360 360

5 5 5 5 10 10 10

3.78 3.24 3.32 3.26 3.31 3.27 3.22

1.18 2.05 1.85 0.13 0.61 0.62 0.46

0.95 1.06 0.91 0.42 0.22 0.53 0.41

48 49 45 47 48 49 49

62 118 91 14 20 30 22

54 104 80 10 16 25 18

288 367 303 101 121 144 121

452 639 519 174 206 248 209

However, even less restricted designs based on a 20 mT Orion CEV suggests an on-orbit mass requirement of 523 mT for a manned mission to (99942) Apophis, at a reduced mission duration of 180 days and a total delta-v of 12.48 km/s [42]. 8.2. Available time at target asteroid

Fig. 24. Direct comparison of minimum on-orbit masses of the freereturn and redundant PM option for the 12 target candidates.

of 19.20 [17]. 2008 EV5, which is the target candidate of the Marco-Polo-R mission (Section 1.1.1) [7], is also identified as a target candidate for a manned sample-return mission. However, only one single mission profile with a mass below 500 mT (M LEO ¼ 496 mT) is available. In summary, the number of available mission profiles and launch window size for all four NEOs is extremely limited (Table 12). Better accessible targets exist, but their small size (H 4 21:5) gives a high probability (68%) that these objects are fast rotators (T rot 41 h), and may therefore not be suitable for sample-return operations. However, the number of available mission profiles Nprofile is a number of magnitudes higher for these objects. Still, even for the best of these asteroids, on-orbit masses do not drop below 160 mT. The implication is that human missions targeting an NEO are not possible with only a single launch, even with a heavy lift vehicle such as NASA's future Space Launch System (SLS). Multiple launch vehicles and onorbit assembly will be necessary for any future manned NEO sample-return mission. Other studies such as [41,3] also indicated a limited amount of available targets. However, mass requirements are much lower as the utilized spacecraft is based on the Orion CEV, and therefore much smaller than the baseline spacecraft used throughout this study. These studies also do not impose safety requirements such as the use of freereturn trajectories or redundant propulsion systems. This makes a trade-off to reduce mission duration at the cost of additional delta-v much more favourable, allowing a shorter mission duration of 180, 120, and even 90 days.

While a longer time spent at the target asteroid is always favourable – especially considering the long mission duration – results show that minimum on-orbit mass missions have rather short stays at the target asteroids of only 5–10 days. Increasing dt t is possible at the expense of an increased on-orbit mass (Fig. 25). While smaller objects such as 2000 SG344 provide some room for increasing dt t , larger NEOs such as (99942) Apophis quickly exceed the 500 mT limit due to the already high minimum on-orbit mass requirements. 8.3. Free-return vs. redundant PM When comparing the two methods established to ensure return of the spacecraft to Earth, the use of freereturn trajectories provides lower masses, despite the fixed mass of the main spacecraft and restriction to the number of mission profiles (Section 6). An explanation for this is the large influence of delta-v on the mass of the spacecraft. The reduction of mission duration and spacecraft mass (structures and supplies) is not able to compensate the increased delta-v associated with faster transfer trajectories. The need for a different safety feature when free-return trajectories are not in use, in the form of a redundant propulsion module, increase of on-orbit mass even further. Given the current knowledge of the NEO population, the use of free-return trajectories offers the best access to all available NEOs in terms of on-orbit mass requirements. Because targets that offer good access (low delta-v and on-orbit mass) and those that provide more incentive for human exploration (sample-return, size) do not overlap, a trade-off between selecting interesting targets and keeping on-orbit mass low enough is required, when selecting

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219

minimum on−orbit mass [mT]

900 H ≤ 21.5

800

all 700 600 500 400 300 200 100 0

0

10

20

30

40

50

60

70

80

90 100 110 120 130 140 150 160 170 180 190 200 210

time spent at target NEO (dt ) [days] t

Fig. 25. Minimum on-orbit mass based on time spent at target NEO dt t . Larger targets (H r 21:5) are shown separately.

targets for future human NEO exploration. Most likely, the masses required for the larger NEOs (H r21:5) are too high for initial NEO missions, and manned missions will have to gradually evolve, starting with targets that offer easier access, and then moving towards more demanding missions. These results suggest that NEO accessibility can only be achieved through a scalable program, not a reliance on heavy lift vehicles. Conducting missions to any NEO cannot be achieved with a single heavy lift vehicle, and these missions will rely on the utilization of technologies such as on-orbit assembly and resupply/refuelling, including long-term fuel storage. This becomes even more important when access to a large number of NEOs is to be provided, as on-orbit masses quickly exceed the payload capability of multiple heavy lift vehicles. This is especially true for larger sized NEOs which offer higher scientific incentive for exploration, and are easier to track remotely, prior to the start of any manned exploration effort.

8.4. Increasing available targets and mission opportunities In order to conduct NEO missions it is necessary to increase the number of available and accessible targets, and thus increase the chances of having an ideal target. This can be attempted through a number of different ways:

8.4.1. Discovery of new NEOs The ongoing discovery of new objects provides the chance of asteroids being found, that satisfy both the requirements regarding size and rotation rate, and are at the same time easy to access. Discovery of new asteroids with diameters in the range of kilometres occurs less than 10 times per year, NEOs with magnitudes below 20, however, have a reasonably high discovery rate (100 per year) [43]. While the discovery of more NEOs is in no way a guarantee to find better targets, efforts towards maintaining/increasing NEO discovery rates do increase the chance of finding potential target candidates. In addition, these efforts also help with the discovery of PHAs, providing additional value regarding space situational awareness.

8.4.2. Refined analysis of potential targets Other than waiting for the discovery of an ideal target, there might be more targets within the currently known NEO population that expected. One reason is that the actual rotation rates of most NEOs are unknown. This means that the chance of an asteroid rotating slow enough for sample-return operations also exists in the magnitude range above the 21.5 limit established in Section 2.2, and even within the currently known NEO population. 8.4.3. Development of advanced mission concepts Next to increasing the number of targets through new discoveries, the development of advanced mission concepts can provide a reduction in on-orbit mass requirements as well as increase the number of available target candidates. Based on the developed mass model, the use of regenerative life-support systems does not reduce spacecraft mass [44]. The main reason for this is that any water saved by recuperation will need to be replaced by shielding mass, next to requiring more complex hardware with higher mass and power consumption. Concepts that could possibly reduce on-orbit mass for a manned NEO mission are the use of a separate cargo vehicle for transporting equipment, supplies and possibly fuel ahead of the manned spacecraft. As a robotic vehicle is less restricted by mission duration, long-term transfers and use of solar-electric propulsion (SEP) are possible options. Concepts such as this require future analysis to determine whether targets allow such a mission scenario, and to evaluate exactly how much on-orbit mass can be saved this way. These concepts will also rely on the availability of technologies for on-orbit assembly and refuelling/resupply. 9. Points for future research Two points that require further research into the development of the baseline mission design and mass estimation model are safety requirements and radiation shielding. Safety requirements in particular will need to be defined in more detail for future missions that go beyond Earth's SOI. The long mission durations make it necessary to clearly

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determine the reliability of critical spacecraft components as well as provide backup plans during contingency scenarios, where necessary. While basic safety considerations to ensure a safe return of the crew to Earth have been implemented into the baseline mission design, additional study of safety requirements is needed, as these can have a great impact on target availability and on-orbit mass requirements. This impact has already been observed with both free-return trajectories and the use of redundant propulsion systems in this study. Further development of the mass estimation model is also required to improve its accuracy. Especially in regard to radiation-shielding requirements, implementation of an adaptive mass based on mission duration and the type of shielding material can both improve mass estimation accuracy and increase crew safety. Since radiation shielding has a critical influence on on-orbit mass, detailed study and simulation is an important next step. The use of simulation tools such as the GEANT4 toolkit, used in the simulation of high energy particles, is likely a starting point for this [45,46]. Research into this will allow evaluating the shielding properties of different materials not only based on mass, but also on their actual shielding properties and tendency to cause secondary cascades. Appendix A. Supplementary data Supplementary data associated with this article can be found in the online version at http://dx.doi.org/10.1016/j. actaastro.2015.02.018.

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