Systems design of a hybrid sail pole-sitter

This paper presents the preliminary systems design of a pole-sitter. This is a spacecraft that hovers over an Earth pole, creating a platform for full hemispheric observation of the polar regions, as well as direct-link telecommunications. To provide the necessary thrust, a hybrid propulsion system combines a solar sail with a more mature solar electric propulsion (SEP) thruster. Previous work by the authors showed that the combination of the two allows lower propellant mass fractions, at the cost of increased system complexity. This paper compares the pure SEP spacecraft with the hybrid spacecraft in terms of the launch mass necessary to deliver a certain payload for a given mission duration. A mass budget is proposed, and the conditions investigated under which the hybrid sail saves on the initial spacecraft initial mass. It is found that the hybrid spacecraft with near- to mid-term sail technology has a lower initial mass than the SEP case if the mission duration is 7 years or more, with greater benefits for longer duration missions. The hybrid spacecraft with far-term sail technology outperforms the pure SEP case even for short missions.     

Solar sail science mission applications and advancement

Solar sailing has long been envisaged as an enabling or disruptive technology. The promise of open-ended missions allows consideration of radically new trajectories and the delivery of spacecraft to previously unreachable or unsustainable observation outposts. A mission catalogue is presented of an extensive range of potential solar sail applications, allowing identification of the key features of missions which are enabled, or significantly enhance, through solar sail propulsion. Through these considerations a solar sail application-pull technology development roadmap is established, using each mission as a technology stepping-stone to the next.     

Accurate approximation of in-ecliptic trajectories for E-sail with constant pitch angle

Propellantless continuous-thrust propulsion systems, such as electric solar wind sails, may be successfully used for new space missions, especially those requiring high-energy orbit transfers. When the mass-to-thrust ratio is sufficiently large, the spacecraft trajectory is characterized by long flight times with a number of revolutions around the Sun. The corresponding mission analysis, especially when addressed within an optimal context, requires a significant amount of simulation effort. Analytical trajectories are therefore useful aids in a preliminary phase of mission design, even though exact solution are very difficult to obtain. The aim of this paper is to present an accurate, analytical, approximation of the spacecraft trajectory generated by an electric solar wind sail with a constant pitch angle, using the latest mathematical model of the thrust vector. Assuming a heliocentric circular parking orbit and a two-dimensional scenario, the simulation results show that the proposed equations are able to accurately describe the actual spacecraft trajectory for a long time interval when the propulsive acceleration magnitude is sufficiently small.     

Solar sailing in realistic mode,based on a locally-optimal control laws

Ballistic design of solar sailing missions in the solar system is composed of defining the design parameters, the control programs, and the trajectories that provide performance goals of a flight. The use of a solar sail spacecraft imposes specific restrictions on mission parameters that include the degradation limit on the flight duration, the maximum temperature of solar sail's surface, the minimum distance from the Sun, the maximum angular velocity of the spacecraft's rotation and others.Many authors considered the impact of these restrictions on the design of the mission separately, but they used a sophisticated method of finding the exact optimal motion control or applied the most straightforward laws of motion control. This paper uses local-optimal control laws at the complete mathematical models of motion and functioning of solar sail spacecraft to describe a technique of designing interplanetary missions. The described method avoids the need to obtain an accurate optimal solution to the control problem and does not cause significant computational difficulties.     

The dynamics and control of the CubeSail mission: A solar sailing demonstration

The CubeSail mission is a low-cost demonstration of the UltraSail solar sailing concept (, ,  and ), using two near-identical CubeSat satellites to deploy a 260 m-long, 20 m² reflecting film. The two satellites are launched as a unit, detumbled, and separated, with the film unwinding symmetrically from motorized reels. The conformity to the CubeSat specification allows for reduction in launch costs as a secondary payload and utilization of the University of Illinois-developed spacecraft bus. The CubeSail demonstration is the first in a series of increasingly-complex missions aimed at validating several spacecraft subsystems, including attitude determination and control, the separation release unit, reel-based film deployment, as well as the dynamical behavior of the sail and on-orbit solar propulsion. The presented work describes dynamical behavior and control methods used during three main phases of the mission. The three phases include initial detumbling and stabilization using magnetic torque actuators, gravity-gradient-based deployment of the film, and steady-state film deformations in low Earth orbit in the presence of external forces of solar radiation pressure, aerodynamic drag, and gravity-gradient.     

Cosmic ray modulation in the outer heliosphere: Predictions for cosmic ray intensities up to the heliopause along Voyager 1 and 2 trajectories

Time dependent cosmic ray modulation in the outer heliosphere is calculated and results are compared to Voyager 1 and 2 observations using a two-dimensional time-dependent cosmic ray transport model. We predict possible future 133–242 MeV proton observations along the Voyager 1 and 2 spacecraft trajectories. Recent theoretical advances in cosmic ray transport parameters are introduced in order to provide a time-dependence for the assumed transport parameters used in the model. This leads to results that are in general compatible with the spacecraft observations in the inner and outer heliosphere over multiple solar cycles. However, for the outer heliosphere, we find that the Voyager 1 and 2 spacecraft observations cannot be fitted with an identical set of parameters along both trajectories. This indicates a possible asymmetric heliosphere or a symmetric heliosphere but with different diffusion parameters in the northern and southern hemispheres, respectively. Furthermore, results indicate that Voyager 2 observations are still under the influence of solar cycle related changes because of the large modulation volume between the heliopause and spacecraft location in contrast to Voyager 1 which shows a steady increase in cosmic ray intensities.     

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.     

深远空连续推进动力与施图林格解的解析

针对太阳系远深距离的探测将是人类下一阶段深空探测活动的主要目标。这一目标的实现依赖于探测器连续推进动力技术的突破。从描述连续常值推力下太空飞行的施图林格解出发,对其中反映的深远空飞行任务有效载荷比、任务时间、飞行距离等关键参数与发动机性能之间的关系进行了深入分析。给出了在特定任务时长、特定飞行距离要求下发动机比冲、功率需要满足的条件及其对有效载荷比、最终飞行速度等指标的影响。此外,基于二体轨道动力学对太阳系行星探测的大椭圆转移轨道和转移能量进行了推导,并对连续推力的太阳帆任务方案涉及的关键技术指标做了理论性的计算。这些结论是对深空探测连续推力方案基础理论的归纳,可以为我国未来开展深远空探测活动提供重要的启发和指导。     

Improving magnetosphere in situ observations using solar sails

Past and current magnetosphere missions employ conventional spacecraft formations for in situ observations of the geomagnetic tail. Conventional spacecraft flying in inertially fixed Keplerian orbits are only aligned with the geomagnetic tail once per year, since the geomagnetic tail is always aligned with the Earth-Sun line, and therefore, rotates annually. Solar sails are able to artificially create sun-synchronous orbits such that the orbit apse line remains aligned with the geomagnetic tail line throughout the entire year. This continuous presence in the geomagnetic tail can significantly increase the science phase for magnetosphere missions. In this paper, the problem of solar sail formation design is explored using nonlinear programming to design optimal two-craft, triangle, and tetrahedron solar sail formations, in terms of formation quality and formation stability. The designed formations are directly compared to the formations used in NASA’s Magnetospheric Multi-Scale mission.     

Paths not taken – The Gossamer roadmap’s other options

Highly efficient low-thrust propulsion is increasingly applied beyond commercial use, also in mainstream and flagship science missions, in combination with gravity assist propulsion. Another recent development is the growth of small spacecraft solutions, not in size but in numbers and individual capabilities.Just over ten years ago, the DLR-ESTEC Gossamer Roadmap to Solar Sailing was set up to guide technology developments towards a propellant-less and highly efficient class of spacecraft for solar system exploration and applications missions: small spacecraft solar sails designed for carefree handling and equipped with carried application modules.Soon, in three dedicated Gossamer Roadmap Science Working Groups it initiated studies of missions uniquely feasible with solar sails such as Displaced L₁ (DL1) space weather advance warning and monitoring, Solar Polar Orbiter (SPO) delivery to very high inclination heliocentric orbit, and multiple Near-Earth Asteroid (NEA) rendezvous (MNR). Together, they demonstrate the capability of near-term solar sails to achieve at least in the inner solar system almost any kind of heliocentric orbit within 10 years, from the Earth-co-orbital to the extremely inclined, eccentric and even retrograde. Noted as part of the MNR study, sail-propelled head-on retrograde kinetic impactors (RKI) go to this extreme to achieve the highest possible specific kinetic energy for the deflection of hazardous asteroids.At DLR, the experience gained in the development of deployable membrane structures leading up to the successful ground deployment test of a (20 m)², i.e., 20 m by 20 m square solar sail at DLR Cologne in 1999 was revitalized and directed towards a 3-step small spacecraft development line from as-soon-as-possible sail deployment demonstration (Gossamer-1) via in-flight evaluation of sail attitude control actuators (Gossamer-2) to an envisaged proving-the-principle flight in the Earth-Moon system (Gossamer-3). First, it turned the concept of solar sail deployment on its head by introducing four separable Boom Sail Deployment Units (BSDU) to be discarded after deployment, enabling lightweight 3-axis stabilized sailcraft. By 2015, this effort culminated in the ground-qualified technology of the DLR Gossamer-1 deployment demonstrator Engineering Qualification Model (EQM). For mission types using separable payloads, such as SPO, MNR and RKI, design concepts can be derived from the BSDU characteristic of DLR Gossamer solar sail technology which share elements with the separation systems of asteroid nanolanders like MASCOT. These nano-spacecraft are an ideal match for solar sails in micro-spacecraft format whose launch configurations are compatible with ESPA and ASAP secondary payload platforms.Like any roadmap, this one contained much more than the planned route from departure to destination and the much shorter distance actually travelled. It is full of lanes, narrow and wide, detours and shortcuts, options and decision branches. Some became the path taken on which we previously reported. More were explored along the originally planned path or as new sidings in search of better options when circumstance changed and the project had to take another turn. But none were dead ends, they just faced the inevitable changes when roadmaps face realities and they were no longer part of the road ahead. To us, they were valuable lessons learned or options up our sleeves. But for future sailors they may be on their road ahead.     

Durability characterization of mechanical interfaces in solar sail membrane structures

The construction of a solar sail from commercially available metallized film presents several challenges. The solar sail membrane is made by seaming together precut lengths of ultrathin metallized polymer film into the required geometry. This assembled sail membrane is then folded into a small stowage volume prior to launch. The sail membranes must have additional features for connecting to rigid structural elements (e.g., sail booms) and must be electrically grounded to the spacecraft bus to prevent charge build up. Space durability of the material and mechanical interfaces of the sail membrane assemblies will be critical for the success of any solar sail mission. In this study, interfaces of polymer/metal joints in a representative solar sail membrane assembly were tested to ensure that the adhesive interfaces and the fastening grommets could withstand the temperature range and expected loads required for mission success. Various adhesion methods, such as surface treatment, commercial adhesives, and fastening systems, were experimentally tested in order to determine the most suitable method of construction.     

Solar ring mission: Building a panorama of the Sun and inner-heliosphere

Solar Ring (SOR) is a proposed space science mission to monitor and study the Sun and inner heliosphere from a full 360° perspective in the ecliptic plane. It will deploy-three 120°-separated spacecraft on the 1-AU orbit. The first spacecraft, S1, locates 30° upstream of the Earth, the second, S2, 90° downstream, and the third, S3, completes the configuration. This design with necessary science instruments, e.g., the Doppler-velocity and vector magnetic field imager, wide-angle coronagraph, and in-situ instruments, will allow us to establish many unprecedented capabilities: (1) provide simultaneous Doppler-velocity observations of the whole solar surface to understand the deep interior, (2) provide vector magnetograms of the whole photosphere — the inner boundary of the solar atmosphere and heliosphere, (3) provide the information of the whole lifetime evolution of solar featured structures, and (4) provide the whole view of solar transients and space weather in the inner heliosphere. With these capabilities, Solar Ring mission aims to address outstanding questions about the origin of solar cycle, the origin of solar eruptions and the origin of extreme space weather events. The successful accomplishment of the mission will construct a panorama of the Sun and inner-heliosphere, and therefore advance our understanding of the star and the space environment that holds our life.     

姿态角幅值约束下的太阳帆Lissajous轨道保持控制

太阳帆航天器以两姿态角作为轨道控制输入时, 其轨道动力学方程具有非仿射非线性特性. 通过人工平动点处线性化获得的线性系统可完成太阳帆航天器轨道保持控制器的分析与设计. 由于线性近似模型为有误差模型, 存在近似有效范围约束, 表现为轨道高度约束和姿态角幅值约束. 本文研究了姿态角幅值约束对线性近似模型有效性的影响, 通过计算给出满足近似误差要求的姿态角幅值约束. 当控制输入存在幅值约束时, 控制器轨道修正能力受到束缚. 通过研究姿态角幅值约束下的最大允许入轨误差, 设计了最大允许入轨误差下线性二次型调节器(LQR)用于轨道保持控制, 并将控制器应用于太阳帆日地三体系统非线性模型中, 实现了日地人工L₁点Lissajous轨道最大允许入轨误差的控制收敛和良好精度下的轨道保持控制.      

Effects of optical parameter measurement uncertainties and solar irradiance fluctuations on solar sailing

The heliocentric orbital dynamics of a spacecraft propelled by a solar sail is affected by some uncertainty sources, including possible inaccuracies in the measurement of the sail film optical properties. Moreover, the solar radiation pressure, which is responsible for the solar sail propulsive acceleration generation, is not time-constant and is subject to fluctuations that are basically unpredictable and superimposed to the well-known 11-year solar activity cycle. In this context, this work aims at investigating the effects of such uncertainties on the actual heliocentric trajectory of a solar sail by means of stochastic simulations performed with a generalized polynomial chaos procedure. The numerical results give an estimation of their impact on the actual heliocentric trajectory and identify whether some of the uncertainty sources are more relevant than others. This is a fundamental information for directing more accurate theoretical and experimental efforts toward the most important parameters, in order to obtain an accurate knowledge of the solar sail thrust vector characteristics and, eventually, of the spacecraft heliocentric position.     

Gamma-ray burst localization with the international solar polar mission

The European Space Agency's Solar Polar spacecraft is scheduled for launch in 1986. A solar X-ray and cosmic gamma ray burst detector will be aboard. Although the solar polar mission will not provide the long baselines originally planned, due to the cancellation of the NASA spacecraft, it is shown that arrival time analysis between the remaining ESA spacecraft and other missions will nevertheless achieve extremely precise localizations.     

The fundamental physics explorer: An ESA technology reference study

ESA technology reference studies are used as a process to identify key technologies and technical challenges of potential future missions not yet in the science programme. This paper reports on the study of the Fundamental Physics Explorer (FPE), a re-usable platform targeted to small missions testing fundamental laws of physics in space. The study addresses three specific areas of interest: special and general relativity tests based on atomic clocks, experiments on the Weak Equivalence Principle (WEP), and studies of Bose–Einstein condensates under microgravity conditions. Starting from preliminary science objectives and payload requirements, three reference missions in the small/medium class range are discussed, based on a re-adaptation of the LISA Pathfinder spacecraft. A 700/3600 km elliptic orbit has been selected to conduct clock tests of special and general relativity, a 700 km circular orbit to perform experiments on the Weak Equivalence Principle and to study Bose–Einstein condensates, each mission being based on a three-axis stabilised spacecraft. It was determined that adaptation of LISA Pathfinder would be required in order to meet the demands of the FPE missions. Moreover it was established that specific payload and spacecraft technology development would be required to realise such a programme.     

Energetic ion emission for active spacecraft control

Future space missions aiming at the accurate measurement of cold plasmas and DC to very low frequency electric fields will require that the potential of their conductive surfaces be actively controlled to be near the ambient plasma potential. In the near-Earth space these spacecraft are usually solar-cell powered; consequently, parts of their surface are most of the time exposed to solar photons. Outside the plasmasphere, a positive surface potential due the dominance of surface-emitted photoelectrons over ambient plasma electrons is to be expected. Photo- and ambient electrons largely determine the potential and positive values between a few Volts up to 100 V have been observed. Active ion emission is the obvious solution of this problem. A liquid metal ion emitter and a saddle field ion emitter are nearing the stage of flight unit fabrication. We will attempt to clamp the spacecraft potential to values close to the plasma potential. We present first results from vacuum chamber tests and describe the emission behaviour and characteristics of emitters producing, respectively, In⁺ and N₂⁺ beams with an energy of ≥ 5 keV.     

Automated design architectures for co-orbiting spacecraft swarms for planetary moon mapping

This work describes the design and optimization of spacecraft swarm missions to meet spatial and temporal visual mapping requirements of missions to planetary moons, using resonant co-orbits. The algorithms described here are a part of Integrated Design Engineering and Automation of Swarms (IDEAS), a spacecraft swarm mission design software that automates the design trajectories, swarm, and spacecraft behaviors in the mission. In the current work, we focus on the swarm design and optimization features of IDEAS, while showing the interaction between the different design modules. In the design segment, we consider the coverage requirements of two general planetary moon mapping missions: global surface mapping and region of interest observation. The configuration of the swarm co-orbits for the two missions is described, where the participating spacecraft have resonant encounters with the moon on their orbital apoapsis. We relate the swarm design to trajectory design through the orbit insertion maneuver performed on the interplanetary trajectory using aero-braking. We then present algorithms to model visual coverage, and collision avoidance in the swarm. To demonstrate the interaction between different design modules, we relate the trajectory and swarm to spacecraft design through fuel mass, and mission cost estimations using preliminary models. In the optimization segment, we formulate the trajectory and swarm design optimizations for the two missions as Mixed Integer Nonlinear Programming (MINLP) problems. In the current work, we use Genetic Algorithm as the primary optimization solver. However, we also use the Particle Swarm Optimizer to compare the optimizer performance. Finally, the algorithms described here are demonstrated through numerical case studies, where the two visual mapping missions are designed to explore the Martian moon Deimos.     

Spinning solar sail orbit steering via spin rate control

The orbit of a solar sail can be controlled by changing the attitude of the spacecraft. In this study, we consider the spinning solar power sail IKAROS (Interplanetary Kite-craft Accelerated by Radiation Of the Sun), which is managed by Japan Aerospace Exploration Agency (JAXA). The IKAROS attitude, i.e., the direction of its spin-axis, is nominally controlled by the rhumb-line control method. By utilizing the solar radiation torque, however, we are able to change the direction of the spin-axis by only controlling its spin rate. With this spin rate control, we can also control indirectly the solar sail’s trajectory. The main objective of this study is to construct the orbit control strategy of the solar sail via the spin-rate control method. We evaluate this strategy in terms of its propellant consumption compared to the rhumb-line control method. Finally, we present the actual flight attitude data of IKAROS and the change of its trajectory.