Spacecraft Injection into a Heliocentric Terrestrial Orbit

We consider the problem of injection of a spacecraft into the heliocentric Earth's orbit ahead and/or behind the Earth by 60° and 120° in heliographic longitude. The range of solar and astrophysical problems for which these orbits are necessary is reviewed. The variants of injection into heliocentric orbits work from a low around-Earth orbit with one turn-on of the engine in this orbit and one turn-on at the end of the injection trajectory. In this case, it turns out to be more profitable to put spacecraft into orbit for three or even four revolutions of the Earth about the Sun. The velocities necessary for the start from a low around-Earth orbit, the velocities at the final point of injection, and the fuel mass (relative to the spacecraft mass) necessary for injection are estimated. The problems for which injection to similar orbits is executed, using the low-thrust engine and with a combined regime of injection, are also considered.     

Controlling the motion of a spacecraft when approaching a large object of space debris

The problem of calculating the parameters of maneuvering a spacecraft as it approaches a large object of space debris (LOSD) in close near-circular noncoplanar orbits has been considered. In [1–4], the results of analyzing the problem of the flyby of the separated LOSD groups have been presented. It has been assumed that a collector spacecraft approaches the LOSD and captures it or it is inserted into the nozzle of a small spacecraft that has a proper propulsion system (PS). However, in these papers, the flight from one object to another was only analyzed and the problem of approaching to LOSD with a given accuracy was not considered. This paper is a supplement to the cycle of papers [1–4]. It is assumed that, the final stage of approaching the LOSD is implemented by maneuvering in many orbits (up to several dozens) with low-thrust engines, but the PS operating time is fairly small compared with the orbit period in order to make it possible to use impulse approximation in the calculations.     

嫦娥五号平动点拓展任务轨道方案研究

中国嫦娥五号探测器成功实现月球采样返回任务,为最大限度利用任务资源,研究了利用嫦娥五号轨道器的平动点拓展任务轨道方案,设计了平动点轨道及其转移轨道.首先,给出了任务轨道设计的轨道动力学模型,包括圆型限制性三体问题模型和精确力模型.其次,基于嫦娥二号和嫦娥5T1平动点拓展任务设计经验,介绍了平动点轨道直接转移与入轨等轨道...     

Space patrol: Variants of the optical barrier scheme

Different variants of the space patrol system to be designed for discovering and cataloging space objects hazardous for the Earth have been investigated. The basic idea of this system is to create an optical barrier using the telescopes deployed in a heliocentric orbit. Difficulties (as well as ways of overcoming them) of this program are analyzed, associated with form and position of the orbit of a space object relative to the patrol spacecraft, determination of orbit parameters, and mutual motion of space objects and the telescopes on spacecraft. The barrier’s schemes with scanning vertical or horizontal belts are considered. Some examples of observational conditions are presented for space objects crossing the barrier region: angular positions, velocities, distances, and numbers of days during which they are observed in the barrier region. The barrier’s characteristics are given for telescopes deployed in the orbits of the Earth and Venus.     

Mission design and analysis of European astrophysics missions orbiting libration points

The main characteristics of the trajectory design of space observatory missions in the Earth–Sun libration point region is highlighted, based on experiences gained in work performed by the authors on ESA missions. Free transfers always lead to large-amplitude orbits around L₂, their properties (amplitudes, phases, non-linear behaviour) are related to the conditions at perigee. Launch scenarios with different degrees of freedom in the perigee geometry and different strategies of sharing the apogee raising between launcher and spacecraft propulsion for Soyuz (with circular parking orbit or direct injection) and Ariane 5 launches from French Guiana will be discussed. Besides the orbit selection and transfer analysis, an important aspect of libration missions is the maintenance of the operational orbit. For some missions it is required to maximise the time between maintenance manoeuvres, and for some the thrust authority is limited. In both cases the exponential nature of the state transition matrix has to be considered. If the equivalent velocity error in the unstable direction becomes too large, the orbit can become unrecoverable, leading to a departure from the environment of the Lagrange point within a few months.     

Novel mission concepts for polar coverage: An overview of recent developments and possible future applications

The paper provides a survey of novel mission concepts for continuous, hemispheric polar observation and direct-link polar telecommunications. It is well known that these services cannot be provided by traditional platforms: geostationary satellites do not cover high-latitude regions, while low- and medium-orbit Sun-synchronous spacecraft only cover a narrow swath of the Earth at each passage. Concepts that are proposed in the literature are described, including the pole-sitter concept (in which a spacecraft is stationary above the pole), spacecraft in artificial equilibrium points in the Sun–Earth system and non-Keplerian polar Molniya orbits. Additionally, novel displaced eight-shaped orbits at Lagrangian points are presented. For many of these concepts, a continuous acceleration is required and propulsion systems include solar electric propulsion, solar sail and a hybridisation of the two. Advantages and drawbacks of each mission concept are assessed, and a comparison in terms of high-latitude coverage and distance, spacecraft mass, payload and lifetime is presented. Finally, the paper will describe a number of potential applications enabled by these concepts, focusing on polar Earth observation and telecommunications.     

The management of the geostationary environment

The population in the geostationary orbit is increasing at the rate of about 25 spacecraft a year and operating lifetimes are increasing. The size of the spacecraft is increasing, as is the power level. The only way to protect the operational arc is to reboost spacecraft at end of life to a burial orbit. While most operators do some reboost maneuver at end of mission there has been no agreed upon criterion for the maneuver. The ITU-R S. 1003¹ recommends reboost of not less than 300 km with the apogee as high as possible. The Interagency Debris Coordination Working Group (IADC) has recently achieved a consensus on a recommendation that the minimum maneuver be 235 km + Cr 1000 A/M. The concept is that this accommodates the ± 3 7.5 km variance in normal radial positioning and a 167.5 km corridor above the arc for repositioning or supersynchronous delivery and establishes a criterion by which the dispositioned spacecraft will never enter that zone after its completion of the maneuver. It also deals with the fact the area mass ratio of spacecraft has been evolving to higher values. Earlier spacecraft had characteristic values of 0.03 but the average now is closer to 0.05 and there are some as great as 0.10.

Disposition of the upper stage should be the same as the spacecraft if it is delivered to GSO. It is preferable to have the stage deliver the spacecraft supersynchronous and then have the spacecraft maneuver down to the GSO.     

Mission analysis and systems design of a near-term and far-term pole-sitter mission

This paper provides a detailed mission analysis and systems design of a near-term and far-term pole-sitter mission. The pole-sitter concept was previously introduced as a solution to the poor temporal resolution of polar observations from highly inclined, low Earth orbits and the poor high-latitude coverage from geostationary orbit. It considers a spacecraft that is continuously above either the north or south pole and, as such, can provide real-time, continuous and hemispherical coverage of the polar regions. Being on a non-Keplerian orbit, a continuous thrust is required to maintain the pole-sitter position. For this, two different propulsion strategies are proposed, which result in a near-term pole-sitter mission using solar electric propulsion (SEP) and a far-term pole-sitter mission where the SEP thruster is hybridized with a solar sail. For both propulsion strategies, minimum propellant pole-sitter orbits are designed. In order to maximize the spacecraft mass at the start of the operations phase of the mission, the transfer from Earth to the pole-sitter orbit is designed and optimized assuming either a Soyuz or an Ariane 5 launch. The maximized mass upon injection into the pole-sitter orbit is subsequently used in a detailed mass budget analysis that will allow for a trade-off between mission lifetime and payload mass capacity. Also, candidate payloads for a range of applications are investigated. Finally, transfers between north and south pole-sitter orbits are considered to overcome the limitations in observations due to the tilt of the Earth's rotational axis that causes the poles to be alternately situated in darkness. It will be shown that in some cases these transfers allow for propellant savings, enabling a further extension of the pole-sitter mission.     

Return to Venus of the Japanese Venus Climate Orbiter AKATSUKI

Japanese Venus Climate Orbiter/AKATSUKI was proposed in 2001 with strong support by international Venus science community and approved as an ISAS (The Institute of Space and Astronautical Science) mission soon after the proposal. The mission life we expected was more than two Earth years in Venus orbit. AKATSUKI was successfully launched at 06:58:22JST on May 21, 2010, by H-IIA F17. After the separation from H-IIA, the telemetry from AKATSUKI was normally detected by DSN Goldstone station (10:00JST) and the solar cell paddles’ deployment was confirmed. After a successful cruise, the malfunction happened on the propulsion system during the Venus orbit insertion (VOI) on Dec. 7, 2010. The engine shut down before the planned reduction in speed to achieve. The spacecraft did not enter the Venus orbit but entered an orbit around the Sun with a period of 203 days. Most of the fuel still had remained, but the orbital maneuvering engine was found to be broken and unusable. However, we have found an alternate way of achieving orbit by using only the reaction control system (RSC). We had adopted the alternate way for orbital maneuver and three minor maneuvers in Nov. 2011 were successfully done so that AKATSUKI would meet Venus in 2015. We are considering several scenarios for VOI using only RCS.     

NanoSail-D: A solar sail demonstration mission

In the early to mid-2000s, NASA made substantial progress in the development of solar sail propulsion systems. Solar sail propulsion uses the solar radiation pressure exerted by the momentum transfer of reflected photons to generate a net force on a spacecraft. To date, solar sail propulsion systems were designed for large robotic spacecraft. Recently, however, NASA has been investigating the application of solar sails for small satellite propulsion. The NanoSail-D is a subscale solar sail system designed for possible small spacecraft applications. The NanoSail-D mission flew on board the ill-fated Falcon Rocket launched August 2, 2008, and due to the failure of that rocket, never achieved orbit. The NanoSail-D flight spare is ready for flight and a suitable launch arrangement is being actively pursued. This paper will present an introduction solar sail propulsion systems and an overview of the NanoSail-D spacecraft.     

Halo orbits in the vicinity of the L 2 libration point in the Sun-Earth system

Methods are proposed for constructing the orbits of spacecraft remaining for long periods of time in the vicinity of the L ₂ libration point in the Sun-Earth system (so-called halo orbits), and the trajectories of uncontrolled flights from low near-Earth orbits to halo orbits. Halo orbits and flight trajectories are constructed in two stages: A suitable solution to a circular restricted three-body problem is first constructed and then transformed into the solution for a restricted four-body problem in view of the real motions of the Sun, Earth, and Moon. For a halo orbit, its prototype in the first stage is a combination of a periodic Lyapunov solution in the vicinity of the L ₂ point and lying in the plane of large-body motion, with the solution for the linear second-order system describing small deviations of the spacecraft from this plane along the periodic solution. The desired orbit is found as the solution to the three-body problem best approximating the prototype in the mean square. The constructed orbit serves as a similar prototype in the second stage. In both stages, the approximating solution is constructed by continuation along a parameter that is the length of the approximation interval. Flight trajectories are constructed in a similar manner. The prototype orbit in the first stage is a combination of a solution lying in the plane of large-body motion and a solution for a linear second-order system describing small deviations of the spacecraft from this plane. The planar solution begins near the Earth and over time tends toward the Lyapunov solution existing in the vicinity of the L ₂ point. The initial conditions of both prototypes and the approximating solutions correspond to the spacecraft’s departure from a low near-Earth orbit at a given distance, perigee, and inclination.     

Optimization of space orbits design for Earth orbiting missions

In Earth orbiting space missions, the orbit selection dictates the mission parameters like the ground resolution, the area coverage, and the frequency of coverage parameters. To achieve desired mission parameters, usually Earth regions of interest are identified and the spacecraft is maneuvered continuously to visit only these regions. This method is expensive, it requires a propulsion system onboard the spacecraft, working throughout the mission lifetime. It also requires a longer time to cover all the regions of interest, due to the very weak thrust forces compared to that of the Earth's gravitational field. This paper presents a methodology to design natural orbits, in which the regions of interest are visited without the use of propulsion systems, depending only on the gravitational forces. The problem is formulated as an optimization problem. A genetic algorithm along with a second order gradient method is implemented for optimization. The design process takes into consideration the gravitational second zonal harmonic, and hence allows for the design of repeated Sun-synchronous orbits. The field of view of the payload is also taken into consideration in the optimization process. Numerical results are presented that demonstrates the efficiency of the proposed method.     

层次分析法在空间推进系统综合评价中的应用

在建立航天器推进系统综合评定体系的基础上，动用层次分析法对已经应用飞行的推进系统和有未来应用潜力的推进系统进行了评估。评估结果表明，从能量特性，使用性能、系统可靠性、经济性、系统特性五个准则上看，过氧化氢／煤油无毒推进系统是未来航天器在轨推进的最优选择。     

How we will go to Mars

This article studies the efficiency of ejecting waste generated by the life support system (LSS) of a manned spacecraft to reduce initial mass on low earth orbit. The spacecraft is used for a long-duration interplanetary mission and is equipped with either a chemical or a nuclear-thermal propulsion system. For this study we simulate an optimal control problem for a given spacecraft maneuver. An impulsive approximation of the optimal interplanetary spacecraft trajectory is assumed, which allows us to reduce the general optimal control problem to hierarchic structure of 'outer' and 'inner' subproblems. This structure is analyzed using the Pontryagin's Maximum principle. Numerical results, illustrating the efficiency of waste ejection are shown for typical Earth-Mars transfer trajectories. This results confirm in theory that using a waste ejection system makes an early manned Mars mission possible without having to design and build new, advanced biological LSS.     

Applications of multi-body dynamical environments: The ARTEMIS transfer trajectory design

The application of forces in multi-body dynamical environments to permit the transfer of spacecraft from Earth orbit to Sun–Earth weak stability regions and then return to the Earth–Moon libration (L₁ and L₂) orbits has been successfully accomplished for the first time. This demonstrated that transfer is a positive step in the realization of a design process that can be used to transfer spacecraft with minimal Delta-V expenditures. Initialized using gravity assists to overcome fuel constraints; the ARTEMIS trajectory design has successfully placed two spacecrafts into Earth–Moon libration orbits by means of these applications.     

The use of three-impulse transfer to insert the spacecraft into the high Moon Artificial Satellite orbits

The problem of the optimal spacecraft’s insertion from the Earth into the high circular polar Moon Artificial Satellite’s orbit (MAS) with a radius of 4000–8000 km has been investigated. A comparison of single- and three-impulse insertion schemes has been performed. The analysis was made taking into account the disturbances from the lunar gravity field harmonics and the gravity fields of the Earth and the Sun, as well as the engine’s limited thrust. It has been shown that the three-impulse transfer from the initial selenocentric hyperbola of the approach into the considered final high MAS orbit is noticeably better with respect to the final mass than the ordinary single-impulse deceleration. The control parameters that implement this maneuver and provide nearly the same energy expenses as in the Keplerian case have been presented. It was found that, in contrast to the Keplerian case, in the considered case of the real gravity field, there is the optimal maximum distance of the maneuver. Recently, the Moon exploration problem became actual again.     

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.     

Mapping orbits with low station keeping costs for constellations of satellites based on the integral over the time of the perturbing forces

The present paper is concerned with the search for orbits that have potential to require low fuel consumption for station-keeping maneuvers for constellations of satellites. The method used to study this problem is based on the integral over the time of the undesired perturbing forces. This integral measures the change of velocity caused by the perturbation forces acting on the satellite, so mapping orbits that are less perturbed, which generates good candidates for orbits that requires low fuel consumption for station-keeping maneuvers. The integral over the time depends only on the orbit of the spacecraft and the dynamical system considered. The type of engine and the control technique applied to the spacecraft are not considered to search for those orbits. It can be a good strategy to be applied for a first mapping of orbits. For this search, it is analyzed the integral of orbits with different values of the Keplerian elements in order to find the best ones with respect to this criterion. The perturbations considered are the ones caused by the third body, which includes the Sun and the Moon, and the J₂ term of the geopotential. The results presented here show numerical simulations to obtain the integral of those perturbing forces for different orbits. The GPS and the Molniya constellations are used as examples for those calculations.     

Earth–Moon libration point orbit stationkeeping: Theory,modeling, and operations

Collinear Earth–Moon libration points have emerged as locations with immediate applications. These libration point orbits are inherently unstable and must be maintained regularly which constrains operations and maneuver locations. Stationkeeping is challenging due to relatively short time scales for divergence, effects of large orbital eccentricity of the secondary body, and third-body perturbations. Using the Acceleration Reconnection and Turbulence and Electrodynamics of the Moon's Interaction with the Sun (ARTEMIS) mission orbit as a platform, the fundamental behavior of the trajectories is explored using Poincaré maps in the circular restricted three-body problem. Operational stationkeeping results obtained using the Optimal Continuation Strategy are presented and compared to orbit stability information generated from mode analysis based in dynamical systems theory.     

单航天器无需机动对星座多星交会轨道全解

文献[1,2]已研究了单航天器无需变轨对Walker星座多星交会的轨道设计,在此基础上,依据交会点必是轨道交点的轨道特性,提出了轨道全解的解析法,可给出该轨道所能够交会卫星的最大数目,研究结果可为单航天器无需大机动变轨对星座多星接近的轨道应用提供理论参考。