Delayed gratification habitable zones: when deep outer solar system regions become balmy during post-main sequence stellar evolution

Like all low- and moderate-mass stars, the Sun will burn as a red giant during its later evolution, generating of solar luminosities for some tens of millions of years. During this post-main sequence phase, the habitable (i.e., liquid water) thermal zone of our Solar System will lie in the region where Triton, Pluto-Charon, and Kuiper Belt objects orbit. Compared with the 1 AU habitable zone where Earth resides, this "delayed gratification habitable zone" (DGHZ) will enjoy a far less biologically hazardous environment - with lower harmful radiation levels from the Sun, and a far less destructive collisional environment. Objects like Triton, Pluto-Charon, and Kuiper Belt objects, which are known to be rich in both water and organics, will then become possible sites for biochemical and perhaps even biological evolution. The Kuiper Belt, with >10(5) objects > or =50 km in radius and more than three times the combined surface area of the four terrestrial planets, provides numerous sites for possible evolution once the Sun's DGHZ reaches it. The Sun's DGHZ might be thought to only be of academic interest owing to its great separation from us in time. However, approximately 10(9) Milky Way stars burn as luminous red giants today. Thus, if icy-organic objects are common in the 20-50 AU zones of these stars, as they are in our Solar System (and as inferred in numerous main sequence stellar disk systems), then DGHZs may form a niche type of habitable zone that is likely to be numerically common in the Galaxy.     

Preliminary assessment of a small robotic rover for Titan exploration

Titan is a very interesting target in deep space exploration. With its solid surface on which a rover can easily travel and its methane lakes which can be sailed it is the ideal target for a deep space mission which includes a mobile platform. In the present paper the general layout of a rover for a mission to Titan is studied, dealing with the mobility, power generation and trajectory control issues. A four-wheels configuration with slip steering was chosen; to compare this solution with the more conventional strategy based on steering wheels, simulations were performed on some trajectories computed through the well known ‘potential’ method, using both slip steering and conventional steering control, for different vehicle speeds. The comparison between the simulated trajectories allows to state the adequateness of the proposed approach.The results here obtained apply not only to a future mission to Titan, but also to other missions designed for the exploration of the satellites of the outer planets having a size comparable with that of Titan and the largest Kuiper belt objects like Pluto and 136472 Makemake.     

MESSENGER's use of solar sailing for cost and risk reduction

The MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) mission used six planetary gravity assists in order to enable capture into orbit about Mercury. A key element of MESSENGER's successful trajectory was achieving the proper gravity assist from each planetary flyby. The criticality of the MESSENGER gravity assists levied tight accuracy requirements on the planetary-flyby targeting. Major errors could have precluded Mercury orbit insertion or required modifications to the trajectory that increased mission complexity, cost, and risk by requiring additional Mercury flybys and extending mission duration. Throughout the mission, MESSENGER modified its strategy for achieving accurate planetary flybys. By using solar sailing, the MESSENGER team was able to eliminate all of the flyby approach maneuvers without sacrificing flyby accuracy, thereby saving mission ΔV margin. The elimination of these approach maneuvers also markedly reduced mission risk, as these approach maneuvers were nominally planned during a time of heightened sensitivity to errors and precluded unique flyby science opportunities. The paradigm shift used by MESSENGER may be useful for other interplanetary missions, particularly if their trajectories require gravity assists in the inner solar system.     

Rosetta visits asteroid (21)Lutetia

The International Rosetta Mission, cornerstone of the European Space Agency Scientific Programme, was launched on 2nd March 2004 to its 10 years journey to comet Churyumov–Gerasimenko. Rosetta will reach the comet in summer 2014, orbit it for about 1.5 years down to distances of a few Kilometres and deliver the Lander Philae onto its surface. After its successful asteroid fly-by in September 2008, Rosetta came back to Earth, for the final gravity acceleration towards its longest heliocentric orbit, up to a distance of 5.3 AU. It is during this phase that Rosetta crossed for the second time the main asteroids belt and performed a close encounter with asteroid (21)Lutetia on the 10th of July 2010 at a distance of ca. 3160 km and a relative velocity of 15 km/s. The payload complement of the spacecraft was activated to perform highly valuable scientific observations. The approach phase to the celestial body required a careful and accurate optical navigation campaign that will prove to be useful also for the comet approach phase. The experience gained with first asteroid flyby in 2008 was fed back into the operations definition and preparation for this highly critical phase; this concerns in particular the operations of the navigation camera for the close-loop autonomous asteroid tracking and of the main scientific camera for high resolution imaging. It was shortly after the flyby that Rosetta became the solar-powered spacecraft to have flown furthest from the Sun (>2.72 AU). This paper presents the activities carried out and planned for the definition, preparation and implementation of the asteroid flyby mission operations, including the test campaign conducted to improve the performance of the spacecraft and payload compared to the previous flyby. The results of the flyby itself are presented, with the operations implemented, the achieved performance and the lessons learned.     

The Aurora project: Estimation of the optical sail parameters

The thrust model for computing a sailcraft trajectory contains thermo-optical parameters that are averages over the spectrum of the incident photons, namely, with respect to energy, intensity and polarization. These parameters are not observables and could change considerably from a sail to sail for a number of practical reasons. The mission analysis for a sailcraft is a progressive task from a simple trajectory propagator to the orbit determination. Aurora sailcraft mission analysis has advanced another step forward by processing experimental data related to aluminium. Once appropriate fitting functions have been selected, differential specular and diffused reflectance and differential absorptance have been averaged over incident solar spectrum, assumed a Planckian here. The result has been to get incident-angle-dependent optical parameters more reliable than mere literature values. The procedure has entailed the computation of grids of complicated definite integrals. They are particularly important during the trajectory optimization of the Aurora solar flyby, a very sensitive profile that would allow the sailcraft to achieve cruise speeds ranging from 12 to 20 AU/yr.     

Opportunity options for rendezvous,flyby and sample return mission to different spectral-type asteroids for the 2015–2025

The feasible rendezvous, flyby and sample return mission scenario to different spectral-type asteroids for the 2015–2025 are investigated. The emphasis is put on the potential target selection and the design of preliminary interplanetary transfer trajectory in this paper. First, according to different scientific motivations, some potential targets with different spectral-type and physical property are selected. Then, some optimal rendezvous and sample return opportunities for different spectral-type asteroids are presented by using pork-chop plots method and Sequential Quadratic-Programming (SQP) algorithm. In order to reduce the launch energy and total velocity increments for sample return mission, the Earth swingby strategy is used. In addition, the feasible trajectory profiles of flyby and rendezvous with two different spectral-type asteroids in one mission are discussed. A hybrid optimization method combing the Differential Evolution (DE) algorithm and SQP algorithm is introduced as a trajectory design method for the mission. Finally, some important parameters of transfer trajectory are analyzed, which would have a direct impact on the design of spacecraft subsystem, such as communication, power and thermal control subsystem.     

“新视野”探测器热控设计特点分析及启示

"新视野"探测器是第一个用于探测冥王星、冥王星卫星以及柯伊伯带其他天体的航天器。其在发射、在轨飞行和探测过程中经历了复杂多变的热环境,热控系统须保证探测器及器上仪器设备在适宜的温度范围内正常工作。文章介绍了该探测器热控系统的总体设计方案,重点分析了热控设计的依据、特点及其带来的启示。     

Cassini Saturn-escape trajectories to Jupiter,Uranus, and Neptune

Potential encore-mission scenarios have been considered for the Cassini mission. In this paper we discuss one of the end-of-life scenarios in which the Cassini spacecraft could perform a Saturn escape via gravity assists from Titan. It is shown that such satellite-aided escape requires a small deterministic maneuver (e.g., Δv<50 m/s), but provides enough energy for the Cassini spacecraft to reach a range of targets in our Solar System, as close to the Sun as the asteroid belt or as far as the Kuiper belt. The escape sequence could be initiated from an arbitrary point during the on-going Cassini mission. Example tours are presented in which the final Titan flyby places the spacecraft into ballistic trajectories that reach Jupiter, Uranus, and Neptune. After years of heliocentric flight, the spacecraft could impact on the target gas giant or perform a flyby to escape from the Solar System (if not to another destination). The concept can be generalized to a new kind of missions, including nested-grand tours, which may involve satellite-aided captures and escapes at more than one planet.     

Investigation of nuclear electric powered interstellar precursor missions

Nuclear Electric Propulsion (NEP) is a technology conceptually proposed since the 1940s by E. Stuhlinger in Germany. The JIMO mission originally planned by NASA in the early 2000s produced at least two designs of ion thrusters fed by a 20–30 kW nuclear powerplant. When compared to conventional (chemical) propulsion, the major advantage of NEP in the JIMO context was recognized to be the much higher I_sp (lab-tested at up to 15,000 s) and the capability for sustained power generation, up to 8–10 years when derated to I_sp about 8000 s.The goal of this paper is to show that current or near term NEP technology enables missions far beyond our immediate interplanetary backyard. In fact, by extending the semi-analytical approach used by Stuhlinger, with reasonable ratios α≡power/mass of the propulsion system (i.e., 0.1– 0.4 kW/kg), missions to the Kuiper Belt (40 AU and beyond) and even the so-called FOCAL mission (at 540 AU) become feasible with an attractive payload fraction and in times of order 10–15 years.Further results regarding missions to Sedna’s perihelion/aphelion, and to Oort’s cloud will also be presented, showing the constraints affecting their feasibility and mass budget.     

Very high delta-V missions to the edge of the solar system and beyond enabled by the dual-stage 4-grid ion thruster concept

A new and innovative type of gridded ion thruster, the “Dual-Stage 4-Grid” or DS4G concept, has been proposed and its predicted high performance validated under an ESA research, development and test programme. The DS4G concept is able to operate at very high specific impulse and thrust density values well in excess of conventional 3-grid ion thrusters at the expense of a higher power-to-thrust ratio. This makes it a possible candidate for ambitious missions requiring very high delta-V capability and high power. Such missions include 100 kW-level multi-ton probes based on nuclear and solar electric propulsion (SEP) to distant Kuiper Belt Object and inner Oort cloud objects, and to the Local Interstellar medium. In this paper, the DS4G concept is introduced and its application to this mission class is investigated. Benefits of using the DS4G over conventional thrusters include reduced transfer time and increased payload mass, if suitably advanced lightweight power system technologies are developed.A mission-level optimisation is performed (launch, spacecraft system design and low-thrust trajectory combined) in order to find design solutions with minimum transfer time, maximum scientific payload mass, and to explore the influence of power system specific mass. It is found that the DS4G enables an 8-ton spacecraft with a payload mass of 400 kg, equipped with a 65 kW nuclear reactor with specific mass 25 kg/kW (e.g. Topaz-type with Brayton cycle conversion) to reach 200 AU in 23 years after an Earth escape launch by Ariane 5. In this scenario, the optimum specific impulse for the mission is over 10,000 s, which is well within the capabilities of a single 65 kW DS4G thruster. It is also found that an interstellar probe mission to 200 AU could be accomplished in 25 years using a “medium-term” SEP system with a lightweight 155 kW solar array (2 kg/kW specific mass) and thruster PPU (3.7 kg/kW) and an Earth escape launch on Ariane 5. In this case, the optimum specific impulse is lower at 3500 s which is well within conventional gridded ion thruster capability.     

太阳系边际探测飞行任务规划

针对太阳系边际探测任务，开展了星际多目标飞越的任务规划，采用小推力混合优化设计方法完成了基于借力飞行及电推进技术的行星际转移轨道联合优化设计，对比研究了面向日球层鼻尖和尾部探测的星际多目标探测飞行方案。研究表明，探测器在2024-2025年发射，可飞抵日球层鼻尖区域，在2027-2030年发射可飞抵日球层尾部区域，并可在2049年1月1日前飞离日心100 AU，实现太阳系边际空间的科学探测。其中日球层鼻尖探测任务探测器飞抵100 AU的位置位于鼻尖中心区域，可与旅行者1号、2号探测器形成有效互补。文章所用任务规划方法，可为太阳系边际探测的自主任务规划技术提供基础，相关研究成果能够为未来中国首次太阳系边际探测任务的实施提供有价值的参考。     

基于借力飞行技术的星际多目标交会转移轨道设计

针对深空探测中的转移轨道问题,提出一种星际多目标交会转移轨道设计方法。该方法基于Tisserand原理,采用能量曲线确定交会目标的序列,由Pork-Chop图确定设计参数的可行域和时序,避免了传统方法在交会目标序列和初始轨道段设计中的缺陷。将多目标交会转移轨道设计问题归结为一种非线性多约束多变量搜索寻优问题,通过梯度下降法,选取目标函数的负梯度方向作为每步迭代的搜索方向,逐步逼近目标函数的极小值点,从而使转移轨道优化问题得到简化。最后,将本文的设计方法用于解决欧空局的ROSETTA任务深空转移轨道的设计,设计结果与欧空局公布的结果一致,从而验证了该设计方法的可行性和正确性。     

Optimal nodal flyby with near-Earth asteroids using electric sail

The aim of this paper is to quantify the performance of an Electric Solar Wind Sail for accomplishing flyby missions toward one of the two orbital nodes of a near-Earth asteroid. Assuming a simplified, two-dimensional mission scenario, a preliminary mission analysis has been conducted involving the whole known population of those asteroids at the beginning of the 2013 year. The analysis of each mission scenario has been performed within an optimal framework, by calculating the minimum-time trajectory required to reach each orbital node of the target asteroid. A considerable amount of simulation data have been collected, using the spacecraft characteristic acceleration as a parameter to quantify the Electric Solar Wind Sail propulsive performance. The minimum time trajectory exhibits a different structure, which may or may not include a solar wind assist maneuver, depending both on the Sun-node distance and the value of the spacecraft characteristic acceleration. Simulations show that over 60% of near-Earth asteroids can be reached with a total mission time less than 100 days, whereas the entire population can be reached in less than 10 months with a spacecraft characteristic acceleration of 1 mm/s².     

Revolutionary systems and technologies for missions to the outer planets

The Advanced Deep Space System Development Program is managed by the Jet Propulsion Laboratory for NASA and is also called X2000. X2000 is organized to create advanced flight and ground systems for the exploration of the outer planets and beyond; it has been created to develop the engineering elements of flight and ground systems. Payloads will be developed by another team. Each X2000 delivery gets its requirements from a set of planned missions, or “mission customers”.The X2000 First Delivery Project supports missions to the Sun (to 4 solar radii), Europa (looking for a liquid ocean), Mars (in support of several Mars missions including a sample return), a comet (including a sample return), and Pluto followed by a trip into the Kuiper belt. This set of missions leads to some outstanding requirements:

• 1.1. Long-life (10–12 years)
• 2.2. Total Ionizing Dose of 4 Mrad (for a Europa Orbiter)
• 3.3. Average power consumption less than or equal to 150 Watts
• 4.4. Autonomous operations that result in an extreme reduction in operations costs

This paper describes the X2000 first delivery and its technologies following a brief overview of the program.     

A preliminary assessment of an orbiter in the Haumean system: How quickly can a planetary orbiter reach such a distant target?

With the recent discoveries of planetary objects beyond Neptune and Pluto, the vast majority of all sizeable Solar System planetary objects lie now beyond Uranus, where insertion into orbit after a reasonably short travel is still not within the current capabilities of our spacecraft. Being able to go and stop at a transneptunian dwarf planet would represent a step stone for ambitious long-term goals. The pressure to send spacecraft to these bodies will grow, as, among the tens or hundreds of large objects, some will emerge as high priorities for science and exploration missions. It is subsequently necessary to prepare the technologies required for such spacecraft. In addition, being able to achieve a fast journey to a distant object will benefit also missions to closer targets.Thales Alenia Space has carried out a preliminary parameter exploration of such a mission with a challenging target: an orbiter in the Haumean system. The main parameters are the characteristics of the propulsion and power system, as well as the masses of the spacecraft. The exploration has inferred the technological improvement needed for reaching these objects within a reasonable time.     

Extrasolar solar-sail trajectories and dark matter

Hyper-thin, high-speed solar-photon sail space probes exploring the Sun?s Oort comet cloud could also be used to set an upper bound to the concentration of WIMPS (weakly interacting massive particles), one of the suggested (but unconfirmed) forms of dark matter within the vicinity of the solar system. Newton?s Shell Theorem would be applied to determine variations in apparent solar mass as the probe moves further out from the Sun. Application of this technique to the trajectories of Pioneer 10/11 reveals that the upper limit to WIMP concentration within ~60 AU of the Sun is ~0.2 Earth masses, as revealed in studies of the Pioneer Anomaly. If the published accuracy of the Pioneer acceleration measurements can be increased by an order of magnitude, probe trajectory measurements out to ~10,000 AU may confirm or falsify the hypothesis that WIMP mass within the solar vicinity is ~3X star mass. It is shown that a space-manufactured ~40-nm thick beryllium hollow-body solar sail deployed from a ~0.07 AU perihelion is a candidate spacecraft for such a mission. Possible science-team organization strategy for a ~100-year mission to ~10,000 AU is discussed.     

Solution of the flyby problem for large space debris at sun-synchronous orbits

the paper considers the flyby problem related to large space debris (LSD) objects at low earth orbits. The data on the overall dimensions of known last and upper stages of launch vehicles makes it possible to single out five compact groups of such objects from the NORAD catalog in the 500–2000 km altitude interval. The orbits of objects of each group have approximately the same inclinations. The features of the mutual distribution of the orbital planes of LSD objects in the group are shown in a portrait of the evolution of deviations of the right ascension of ascending nodes (RAAN). In the case of the first three groups (inclinations of 71°, 74°, and 81°), the straight lines of relative RAAN deviations of object orbits barely intersect each other. The fourth (83°) and fifth (97°–100°) LSD groups include a considerable number of objects whose orbits are described by straight lines (diagonals), which intersect other lines many times. The use of diagonals makes it possible to significantly reduce the temporal and total characteristic velocity expenditures required for object flybys, but it complicates determination of the flyby sequence. Diagonal solutions can be obtained using elements of graph theory. A solution to the flyby problem is presented for the case of group 5, formed of LSD objects at sun-synchronous orbits.     

太阳帆日心定点悬浮转移轨道设计

研究了太阳帆航天器日心定点悬浮轨道(HFDO)的转移轨道设计问题,以球坐标形式建立了太阳帆的动力学模型,基于该模型给出在日心悬浮轨道基础上实现定点悬浮的条件,提出了一种实现日心定点悬浮的转移轨道设计方法。首先,确定定点悬浮的位置;然后,设计经过该位置的绕日极轨轨道;最后,实施轨道减速实现定点悬浮,并给出了解析形式的轨道控制律。结合太阳极地观测任务,设计了定点悬浮在太阳北极1AU处的太阳帆转移轨道。仿真结果表明:该轨道转移方案总耗时3.5年,太阳帆定点到黄北极距日心1AU处,此后只要保持太阳光垂直照射帆面,即可维持稳定的悬浮状态。     

Measurement and stability of the pointing of the BepiColombo Laser Altimeter under thermal load

The BepiColombo Laser Altimeter (BELA) has been selected to fly on ESA?s BepiColombo mission to Mercury. The instrument will be the first European laser altimeter designed for interplanetary flight. This paper describes the setup used to characterize the angular movements of BELA under the simulated environmental conditions that the instrument will encounter when orbiting Mercury. The system comprises a laser transmitter and a receiving telescope, which can move with respect to each other under thermal load. Tests performed using the Engineering Qualification Model show that the setup is accurate enough to characterize angular movements of the instrument components to an accuracy of ≈10 μrad. The qualification instrument is thermally stable to operate during all mission phases around Mercury proving that the transmitter and receiver sections will remain within the alignment requirements during its mission.     

Optimization of passive orbit with the use of gravity maneuver

A high-precision method of calculating gravitational interactions is applied in order to determine optimal trajectories. A number of problems, necessary for determination of optimal parameters at a launch of a spacecraft and during its flyby near celestial bodies, are considered. The spacecraft trajectory was determined by numerical integration of the equations of passive motion of the spacecraft and of the equations of motion for planets, the Sun, and the Moon. The optimal trajectory of the spacecraft approaching the Sun is determined by fitting its initial conditions.