Sailing at the brink – The no-limits of near-/now-term-technology solar sails and SEP spacecraft in (multiple) NEO rendezvous

Near-Earth object (NEO) in-situ exploration can provide invaluable information for science, possible future deflection actions and resource utilisation. This is only possible with space missions which approach the asteroid from its vicinity, i.e. rendezvous. This paper explores the use of solar sailing as means of propulsion for NEO rendezvous missions. Given the current state of sail technology, we search for multiple rendezvous missions of up to ten years and characteristic acceleration of up to 0.10 mm/s². Using a tree-search technique and subsequent trajectory optimisation, we find numerous options of up to three NEO encounters in the launch window 2019–2027. In addition, we explore steerable and throttleable low-thrust (e.g. solar-electric) rendezvous to a particular group of NEOs, the Taurid swarm. We show that an acceleration of 0.23 mm/s² would suffice for a rendezvous in approximately 2000 days, while shorter transfers are available as the acceleration increases. Finally, we show low-thrust options (0.3 mm/s²) to the fictitious asteroid 2019 PDC, as part of an asteroid deflection exercise.     

Small satellite survey mission to the second Earth moon

This paper presents an innovative space mission devoted to the survey of the small Earth companion asteroid by means of nano platforms. Also known as the second Earth moon, Cruithne, is the target identified for the mission. Both the trajectory to reach the target and a preliminary spacecraft budget are here detailed. The idea is to exploit high efficient ion thrusters to reduce the propellant mass fraction in such a high total impulse mission (of the order of 1e6 Ns). This approach allows for a 100 kg class spacecraft with a very small Earth escape energy (5 km²/s²) to reach the destination in about 320 days. The 31% propellant mass fraction allows for a payload mass fraction of the order of 8% and this is sufficient to embark on such a small spacecraft a couple of nano-satellites deployed once at the target to carry out a complete survey of the asteroid. Two 2U Cubesats are here considered as representative payload, but also other scientific payloads or different platforms might be considered according with the specific mission needs. The small spacecraft used to transfer these to the target guarantees the manoeuvre capabilities during the interplanetary journey, the protection against radiations along the path and the telecommunication relay functions for the data transmission with Earth stations. The approach outlined in the paper offers reliable solutions to the main issues associated with a deep space nano-satellite mission thus allowing the exploitation of distant targets by means of these tiny spacecraft. The study presents an innovative general strategy for the NEO observation and Cruithne is chosen as test bench. This target, however, mainly for its relevant inclination, requires a relatively large propellant mass fraction that can be reduced if low inclination asteroids are of interest. This might increase the payload mass fraction (e.g. additional Cubesats and/or additional scientific payloads on the main bus) for the same 100 kg class mission.     

A small mission for in situ exploration of a primitive binary near-Earth asteroid

We present a concept for a challenging in situ science mission to a primitive, binary near-Earth asteroid. A sub-400-kg spacecraft would use solar electric propulsion to rendezvous with the C-class binary asteroid (175706) 1996 FG3. A campaign of remote observations of both worlds would be followed by landing on the ∼1 km diameter primary to perform in situ measurements. The total available payload mass would be around 34 kg, allowing a wide range of measurement objectives to be addressed. This mission arose during 2004 from the activities of the ad-hoc Small Bodies Group of the DLR-led Planetary Lander Initiative. Although the particular mission scenario proposed here was not studied further per se, the experience was carried over to subsequent European asteroid mission studies, including first LEONARD and now the Marco Polo near-Earth asteroid sample return proposal for ESA’s Cosmic Vision programme. This paper may thus be of interest as much for insight into the life cycle of mission proposals as for the concept itself.     

Economic assessment of high-thrust and solar-sail propulsion for near-earth asteroid mining

Asteroid mining has the potential to greatly reduce the cost of in-space manufacturing, production of propellant for space transportation and consumables for crewed spacecraft, compared to launching the required resources from the Earth’s deep gravity well. This paper discusses the top-level mission architecture and trajectory design for these resource-return missions, comparing high-thrust trajectories with continuous low-thrust solar-sail trajectories. The paper focuses on maximizing the economic Net Present Value, which takes the time-cost of finance into account and therefore balances the returned resource mass and mission duration. The different propulsion methods are compared in terms of maximum economic return and sets of attainable target asteroids. Results for transporting resources to geostationary orbit show that the orbital parameter hyperspace of suitable target asteroids is considerably larger for solar sails, allowing for more flexibility in selecting potential target asteroids. Also, results show that the Net Present Value that can be realized is larger when employing solar sailing instead of chemical propulsion. In addition, it is demonstrated that a higher Net Present Value can be realized when transporting volatiles to the Lunar Gateway instead of geostationary orbit. The paper provides one more step towards making commercial asteroid mining an economically viable reality by integrating trajectory design, propulsion technology and economic modelling.     

A multiple-rendezvous,sample-return mission to two near-Earth asteroids

We propose a dual-rendezvous mission, targeting near-Earth asteroids, including sample-return. The mission, Asteroid Sampling Mission (ASM), consists of two parts: (i) flyby and remote sensing of a Q-type asteroid, and (ii) sampling of a V-type asteroid. The targeted undifferentiated Q-type are found mainly in the near-Earth space, and to this date have not been the target of a space mission. We have chosen, for our sampling target, an asteroid from the basaltic class (V-type), as asteroids in this class exhibit spectral signatures that resemble those of the well-studied Howardite–Eucrite–Diogenite (HED) meteorite suite. With this mission, we expect to answer specific questions about the links between differentiated meteorites and asteroids, as well as gain further insight into the broader issues of early Solar System (SS) evolution and the formation of terrestrial planets. To achieve the mission, we designed a spacecraft with a dry mass of less than 3 tonnes that uses electric propulsion with a solar-electric power supply of 15 kW at 1 Astronomical Unit (AU). The mission includes a series of remote sensing instruments, envisages landing of the whole spacecraft on the sampling target, and employs an innovative sampling mechanism. Launch is foreseen to occur in 2018, as the designed timetable, and the mission would last about 10 years, bringing back a 150 g subsurface sample within a small re-entry capsule. This paper is a work presented at the 2008 Summer School Alpbach,“Sample return from the Moon, asteroids and comets” organized by the Aeronautics and Space Agency of the Austrian Research Promotion Agency. It is co-sponsored by ESA and the national space authorities of its Member and Co-operating States, with the support of the International Space Science Institute and Austrospace.     

Rapid design and optimization of low-thrust rendezvous/interception trajectory for asteroid deflection missions

Asteroid deflection techniques are essential in order to protect the Earth from catastrophic impacts by hazardous asteroids. Rapid design and optimization of low-thrust rendezvous/interception trajectories is considered as one of the key technologies to successfully deflect potentially hazardous asteroids. In this paper, we address a general framework for the rapid design and optimization of low-thrust rendezvous/interception trajectories for future asteroid deflection missions. The design and optimization process includes three closely associated steps. Firstly, shape-based approaches and genetic algorithm (GA) are adopted to perform preliminary design, which provides a reasonable initial guess for subsequent accurate optimization. Secondly, Radau pseudospectral method is utilized to transcribe the low-thrust trajectory optimization problem into a discrete nonlinear programming (NLP) problem. Finally, sequential quadratic programming (SQP) is used to efficiently solve the nonlinear programming problem and obtain the optimal low-thrust rendezvous/interception trajectories. The rapid design and optimization algorithms developed in this paper are validated by three simulation cases with different performance indexes and boundary constraints.     

Design of a formation of solar pumped lasers for asteroid deflection

This paper presents the design of a multi-spacecraft system for the deflection of asteroids. Each spacecraft is equipped with a fibre laser and a solar concentrator. The laser induces the sublimation of a portion of the surface of the asteroid, and the resultant jet of gas and debris thrusts the asteroid off its natural course. The main idea is to have a formation of spacecraft flying in the proximity of the asteroid with all the spacecraft beaming to the same location to achieve the required deflection thrust. The paper presents the design of the formation orbits and the multi-objective optimisation of the formation in order to minimise the total mass in space and maximise the deflection of the asteroid. The paper demonstrates how significant deflections can be obtained with relatively small sized, easy-to-control spacecraft.     

Asteroid interception and disruption for terminal planetary defense

Modern techniques for planetary defense from comets and asteroids involve the deflection of the bolide via kinetic, gravitational, ablative, or radiative means. While potentially effective, none of these methods are capable of operating in a terminal interdiction mode wherethe threat is discovered with little time prior to impact. We present a practical and effective method for planetary defense which enables extremely short interdiction time scales, but can also operate within longer time scales and can be effective for extremely large threats. Called PI (“Pulverize It”), the method makes use of an array of hypervelocity penetrators which uses the kinetic energy of the asteroid or comet to disrupt it. In the terminal interdiction mode, the fragments of maximum $～10$ m diameter disperse laterally as they continue towards the Earth, and then enter the Earth’s atmosphere where they burn up as a series of airburst events which spatially and temporally de-correlate the energy of the original parent bolide for any arbitrary observer on the ground in the form of acoustical shockwaves and optical pulses. We show that terminal interdiction modes ranging from 2 minutes prior to impact for 20-meter class bolides (such as the Chelyabinsk asteroid), 1 day prior to impact for 100 m-class asteroids, 10 days prior to impact for Apophis-class asteroids ($～370$ m), and even 60 days prior to impact for 1 km-class threats are all possible, though longer warning times are always preferred. Using only technologies readily available today, the PI method allows for a cost-effective and practical roadmap towards robust planetary defense capability.     

Biological assessment of Ariane 5 fairing

United Nations Space Treaties [10 and 11] require the preservation of planets and of Earth from contamination. All nations part to these Treaties shall take measures to prevent forward and backward contamination during missions exploring our solar system. As observer for the United Nations Committee on Peaceful Uses of Outer Space, the COSPAR (Committee of Space Research) defines and handles the applicable policy and proposes recommendations to Space Agencies [COSPAR Planetary Protection Panel, Planetary Protection Policy accepted by the COSPAR Council and Bureau, 20 October 2002, amended 24 March 2005. http://www.cosparhq.org/scistr/PPPolicy.htm.]. The goal is to protect celestial bodies from terrestrial biological contamination as well as to protect the Earth environment from an eventual biohazard which may be carried by extraterrestrial samples or by space systems returning to Earth. According to the applicable specifications, including in our case the French requirements [CNES, System Safety. Planetary Protection Requirements. Normative referential CNES RNC-CNES-R-14, CNES Toulouse, ed. 4, 04 October 2002.], the prevention of forward contamination is accomplished by reducing the bioburden on space hardware to acceptable, prescribed levels, including in some instances system sterilization, assembling and integrating the appropriate spacecraft systems in cleanrooms of appropriate biological cleanliness, avoiding or controlling any recontamination risk, and limiting the probability impact of space systems. In order to prepare for future exploration missions [Debus, A., Planetary protection: organization requirements and needs for future planetary exploration missions, ESA conference publication SP-543, pp 103–114, 2003.], and in particular for missions to Mars requiring to control the spacecraft bioburden, a test program has been developed to evaluate the biological contamination under the fairing of the Ariane 5 launcher.     

A sample collector for robotic sample return missions I: Temperature effect on collected mass

Sample return is playing an increasingly important role in solar system exploration. Among the possible mission on the horizon, are sample return from asteroids, comets, the Moon and Mars. A collector initially intended for near-Earth asteroids is the touch-and-go-impregnable-pad (TGIP). Here we explore the effect of temperature on its collection capabilities. Temperatures expected on near-Earth asteroid mission targets range from −43 to 36 °C. Experiments were conducted at −75, −50, −25, 23, 65, and 105 °C. It was found that the mass of sample collected by the TGIP increased almost linearly to 23 °C and then leveled off at higher temperatures. We also found that the collector did not lose its ability to collect samples after being subjected to −75 °C temperatures (essentially frozen) and then thawed. These experiments have shown that the TGIP can operate effectively at temperatures expected on near-Earth asteroids, especially if collection is performed on the sunward side of the asteroid.     

A sample collector for robotic sample return missions II: Radiation tests

Sample return from small solar system objects is playing an increasingly important part in solar system exploration. Critical to such missions is a robust, simple, and economic sample collector. We have developed a collector such as this for near-Earth asteroid sample return missions that we have termed the Touch-and-Go Impregnable Pad (TGIP). The collector utilizes a silicone substrate that is pushed into the dust and gravel surface layer of the asteroid. As part of a systematic evaluation of the TGIP, we have investigated the resilience of this substrate to ionizing radiations. Several miniature versions of the collector, containing typically ∼3 g of the collection substrate, were exposed to 0.564 MeV beta particles from a ⁹⁰Sr source and a 6 MeV electron beam in a linear accelerator to simulate the wide range of energies of solar and galactic ionizing radiation. Various radiation levels up to eight times greater than expected on a six-year asteroid mission (in the case of beta radiation) and 50 times greater than expected (in the case of the 6 MeV electron radiation) were administered to the substrate. After irradiation, the efficiency of the substrate in collecting samples of mock regolith was compared with that of collectors that had not been irradiated. No difference beyond experimental uncertainty was observed and we suggest that the operational TGIP will not be affected adversely by radiation doses expected during a typical six-year inner solar system mission.     

Trajectory optimization of multiple asteroids exploration with asteroid 2010TK7 as main target

In this study, the Earth’s Trojan asteroid 2010 TK₇ is selected as the rendezvous target. The multiple flyby sequence of asteroid exploration was proposed by optimizing the probe’s orbit. Impulsive maneuvers and low-thrust propulsion were used respectively to design the trajectories of the multiple asteroids exploration mission. Under impulsive maneuvers, gravity assist technique was adopted to reduce fuel consumption. First a reference orbit with only 2010 TK₇ as the rendezvous target was designed. Then five asteroids near the reference orbit were selected as candidates. Finally, we obtained a multiple asteroids exploration sequence of three asteroids based on gravity assist technique and genetic algorithm, and an additional velocity impulse of 0.4?km/s was required. In the subsequent section, a sixth-degree inverse polynomial shape-based method is applied to the low-thrust trajectory design of 2010 TK₇, and the exploration sequence under the action of low-thrust propulsion was provided.     

CubeSail: A low cost CubeSat based solar sail demonstration mission

CubeSail is a nano-solar sail mission based on the 3U CubeSat standard, which is currently being designed and built at the Surrey Space Centre, University of Surrey. CubeSail will have a total mass of around 3 kg and will deploy a 5 × 5 m sail in low Earth orbit. The primary aim of the mission is to demonstrate the concept of solar sailing and end-of-life de-orbiting using the sail membrane as a drag-sail. The spacecraft will have a compact 3-axis stabilised attitude control system, which uses three magnetic torquers aligned with the spacecraft principle axis as well as a novel two-dimensional translation stage separating the spacecraft bus from the sail. CubeSail’s deployment mechanism consists of four novel booms and four-quadrant sail membranes. The proposed booms are made from tape-spring blades and will deploy the sail membrane from a 2U CubeSat standard structure. This paper presents a systems level overview of the CubeSat mission, focusing on the mission orbit and de-orbiting, in addition to the deployment, attitude control and the satellite bus.     

Multiobjective genetic optimization of Earth–Moon trajectories in the restricted four-body problem

In this paper, optimal trajectories of a spacecraft traveling from Earth to Moon using impulsive maneuvers (ΔV maneuvers) are investigated. The total flight time and the summation of impulsive maneuvers ΔV are the objective functions to be minimized. The main celestial bodies influencing the motion of the spacecraft in this journey are Sun, Earth and Moon. Therefore, a three-dimensional restricted four-body problem (R4BP) model is utilized to represent the motion of the spacecraft in the gravitational field of these celestial bodies. The total ΔV of the maneuvers is minimized by eliminating the ΔV required for capturing the spacecraft by Moon. In this regard, only a mid-course impulsive maneuver is utilized for Moon ballistic capture. To achieve such trajectories, the optimization problem is parameterized with respect to the orbital elements of the ballistic capture orbits around Moon, the arrival date and a mid-course maneuver time. The equations of motion are solved backward in time with three impulsive maneuvers up to a specified low Earth parking orbit. The results show high potential and capability of this type of parameterization in finding several Pareto-optimal trajectories. Using the non-dominated sorting genetic algorithm with crowding distance sorting (NSGA-II) for the resulting multiobjective optimization problem, several trajectories are discovered. The resulting trajectories of the presented scheme permit alternative trade-off studies by designers incorporating higher level information and mission priorities.     

Spacecraft formation flying for Earth-crossing object deflections using a power limited laser ablating

A formation flying strategy with an Earth-crossing object (ECO) is proposed to avoid the Earth collision. Assuming that a future conceptual spacecraft equipped with a powerful laser ablation tool already rendezvoused with a fictitious Earth collision object, the optimal required laser operating duration and direction histories are accurately derived to miss the Earth. Based on these results, the concept of formation flying between the object and the spacecraft is applied and analyzed as to establish the spacecraft’s orbital motion design strategy. A fictitious “Apophis”-like object is established to impact with the Earth and two major deflection scenarios are designed and analyzed. These scenarios include the cases for the both short and long laser operating duration to avoid the Earth impact. Also, requirement of onboard laser tool’s for both cases are discussed. As a result, the optimal initial conditions for the spacecraft to maintain its relative trajectory to the object are discovered. Additionally, the discovered optimal initial conditions also satisfied the optimal required laser operating conditions with no additional spacecraft’s own fuel expenditure to achieve the spacecraft formation flying with the ECO. The initial conditions founded in the current research can be used as a spacecraft’s initial rendezvous points with the ECO when designing the future deflection missions with laser ablation tools. The results with proposed strategy are expected to make more advances in the fields of the conceptual studies, especially for the future deflection missions using powerful laser ablation tools.     

Thermal protection system development,testing, and qualification for atmospheric probes and sample return missions: Examples for Saturn,Titan and Stardust-type sample return

The science community has continued to be interested in planetary entry probes, aerocapture, and sample return missions to improve our understanding of the Solar System. As in the case of the Galileo entry probe, such missions are critical to the understanding not only of the individual planets, but also to further knowledge regarding the formation of the Solar System. It is believed that Saturn probes to depths corresponding to 10 bars will be sufficient to provide the desired data on its atmospheric composition. An aerocapture mission would enable delivery of a satellite to provide insight into how gravitational forces cause dynamic changes in Saturn’s ring structure that are akin to the evolution of protoplanetary accretion disks. Heating rates for the “shallow” Saturn probes, Saturn aerocapture, and sample Earth return missions with higher re-entry speeds (13–15 km/s) from Mars, Venus, comets, and asteroids are in the range of 1–6 KW/cm². New, mid-density thermal protection system (TPS) materials for such probes can be mission enabling for mass efficiency and also for use on smaller vehicles enabled by advancements in scientific instrumentation. Past consideration of new Jovian multiprobe missions has been considered problematic without the Giant Planet arcjet facility that was used to qualify carbon phenolic for the Galileo probe. This paper describes emerging TPS technologies and the proposed use of an affordable, small 5 MW arcjet that can be used for TPS development, in test gases appropriate for future planetary probe and aerocapture applications. Emerging TPS technologies of interest include new versions of the Apollo Avcoat material and a densified variant of Phenolic Impregnated Carbon Ablator (PICA). Application of these and other TPS materials and the use of other facilities for development and qualification of TPS for Saturn, Titan, and Sample Return missions of the Stardust class with entry speeds from 6.0 to 28.6 km/s are discussed.     

Dynamics and control of spacecraft hovering using the geomagnetic Lorentz force

To achieve hovering, a spacecraft thrusts continuously to induce an equilibrium state at a desired position. Due to the constraints on the quantity of propellant onboard, long-time hovering around low-Earth orbits (LEO) is hardly achievable using traditional chemical propulsion. The Lorentz force, acting on an electrostatically charged spacecraft as it moves through a planetary magnetic field, provides a new propellantless method for orbital maneuvers. This paper investigates the feasibility of using the induced Lorentz force as an auxiliary means of propulsion for spacecraft hovering. Assuming that the Earth’s magnetic field is a dipole that rotates with the Earth, a dynamical model that characterizes the relative motion of Lorentz spacecraft is derived to analyze the required open-loop control acceleration for hovering. Based on this dynamical model, we first present the hovering configurations that could achieve propellantless hovering and the corresponding required specific charge of a Lorentz spacecraft. For other configurations, optimal open-loop control laws that minimize the control energy consumption are designed. Likewise, the optimal trajectories of required specific charge and control acceleration are both presented. The effect of orbital inclination on the expenditure of control energy is also analyzed. Further, we also develop a closed-loop control approach for propellantless hovering. Numerical results prove the validity of proposed control methods for hovering and show that hovering around low-Earth orbits would be achievable if the required specific charge of a Lorentz spacecraft becomes feasible in the future. Typically, hovering radially several kilometers above a target in LEO requires specific charges on the order of 0.1 C/kg.     

Interplanetary spacecraft navigation using pulsars

We demonstrate how observations of pulsars can be used to help navigate a spacecraft travelling in the solar system. We make use of archival observations of millisecond pulsars from the Parkes radio telescope in order to demonstrate the effectiveness of the method and highlight issues, such as pulsar spin irregularities, which need to be accounted for. We show that observations of four millisecond pulsars every seven days using a realistic X-ray telescope on the spacecraft throughout a journey from Earth to Mars can lead to position determinations better than ∼20 km and velocity measurements with a precision of ∼0.1 ms⁻¹.     

The plasma sheet and boundary layers under northward IMF: A multi-point and multi-instrument perspective

During conditions of northward interplanetary magnetic field (IMF), the near-tail plasma sheet is known to become denser and cooler, and is described as the cold-dense plasma sheet (CDPS). While its source is likely the solar wind, the prominent penetration mechanisms are less clear. The two main candidates are solar wind direct capture via double high-latitude reconnection on the dayside and Kelvin–Helmholtz/diffusive processes at the flank magnetopause. This paper presents a case study on the formation of the CDPS utilizing a wide variety of space- and ground-based observations, but primarily from the Double Star and Polar spacecraft on December 5th, 2004. The pertinent observations can be summarized as follows: TC-1 observes quasi-periodic (∼2 min period) cold-dense boundary layer (compared to a hot-tenuous plasma sheet) signatures interspersed with magnetosheath plasma at the dusk flank magnetopause near the dawn-dusk terminator. Analysis of this region suggests the boundary to be Kelvin–Helmholtz unstable and that plasma transport is ongoing across the boundary. At the same time, IMAGE spacecraft and ground based SuperDARN measurements provide evidence of high-latitude reconnection in both hemispheres. The Polar spacecraft, located in the southern hemisphere afternoon sector, sunward of TC-1, observes a persistent boundary layer with no obvious signature of boundary waves. The plasma is of a similar appearance to that observed by TC-1 inside the boundary layer further down the dusk flank, and by TC-2 in the near-Earth magnetotail. We present comparisons of electron phase space distributions between the spacecraft. Although the dayside boundary layer at Polar is most likely formed via double high-altitude reconnection, and is somewhat comparable to the flank boundary layer at Double Star, some differences argue in favour of additional transport that augment solar wind plasma entry into the tail regions.