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The algorithm for choosing a trajectory of spacecraft flight to the Moon is discussed. The characteristic velocity values needed for correcting the flight trajectory and a braking maneuver are estimated using the Monte Carlo method. The profile of insertion and flight to a near-circular polar orbit with an altitude of ~100 km of an artificial lunar satellite (ALS) is given. The case of two corrections applied during the flight and braking phases is considered. The flight to an ALS orbit is modeled in the geocentric geoequatorial nonrotating coordinate system with the influence of perturbations from the Earth, the Sun, and the Moon factored in. The characteristic correction costs corresponding to corrections performed at different time points are examined. Insertion phase errors, the errors of performing the needed corrections, and the errors of determining the flight trajectory parameters are taken into account. 相似文献
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讨论满足约束条件的月球卫星飞行轨道的设计问题，将约束条件分类为只与太阳，月球，地球，飞行器和观测站之间的相对位置有关的运行学约束条件以及涉及到飞行器轨道运行的动力学约束条件，在考虑月球卫星轨道的受力情况后，给出一种准确快速地计算和设计满足约束条件的标准飞行轨道的方法，并应用于不同约束条件下月球卫星的轨道预设计，初步讨论了轨道设计的误差分析，轨道跟踪及实时精密定轨等正在进行的其它相关工作。 相似文献
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New Challenges to Trajectory Design by the Use of Electric Propulsion and Other New Means of Wandering in the Solar System 总被引：1，自引：0，他引：1
Giuseppe D. Racca 《Celestial Mechanics and Dynamical Astronomy》2003,85(1):1-24
The design of spacecraft trajectories is a crucial part of a space mission design. Often the mission goal is tightly related to the spacecraft trajectory. A geostationary orbit is indeed mandatory for a stationary equatorial position. Visiting a solar system planet implies that a proper trajectory is used to bring the spacecraft from Earth to the vicinity of the planet. The first planetary missions were based on conventional trajectories obtained with chemical engine rockets. The manoeuvres could be considered 'impulsive' and clear limitations to the possible missions were set by the energy required to reach certain orbits. The gravity-assist trajectories opened a new way of wandering through the solar system, by exploiting the gravitational field of some planets. The advent of other propulsion techniques, as electric or ion propulsion and solar sail, opened a new dimension to the planetary trajectory, while at the same time posing new challenges. These 'low thrust' propulsion techniques cannot be considered 'impulsive' anymore and require for their study mathematical techniques which are substantially different from before. The optimisation of such trajectories is also a new field of flight dynamics, which involves complex treatments especially in multi-revolution cases as in a lunar transfer trajectory. One advantage of these trajectories is that they allow to explore regions of space where different bodies gravitationally compete with each other. We can exploit therefore these gravitational perturbations to save fuel or reduce time of flight. The SMART-1 spacecraft, first European mission to the Moon, will test for the first time all these techniques. The paper is a summary report on various activities conducted by the project team in these areas. 相似文献
5.
An accurate determination of the landing trajectory of Chang'e-3(CE-3)is significant for verifying orbital control strategy, optimizing orbital planning, accurately determining the landing site of CE-3 and analyzing the geological background of the landing site. Due to complexities involved in the landing process, there are some differences between the planned trajectory and the actual trajectory of CE-3. The landing camera on CE-3 recorded a sequence of the landing process with a frequency of 10 frames per second. These images recorded by the landing camera and high-resolution images of the lunar surface are utilized to calculate the position of the probe, so as to reconstruct its precise trajectory. This paper proposes using the method of trajectory reconstruction by Single Image Space Resection to make a detailed study of the hovering stage at a height of 100 m above the lunar surface. Analysis of the data shows that the closer CE-3 came to the lunar surface, the higher the spatial resolution of images that were acquired became, and the more accurately the horizontal and vertical position of CE-3 could be determined. The horizontal and vertical accuracies were7.09 m and 4.27 m respectively during the hovering stage at a height of 100.02 m. The reconstructed trajectory can reflect the change in CE-3's position during the powered descent process. A slight movement in CE-3 during the hovering stage is also clearly demonstrated. These results will provide a basis for analysis of orbit control strategy,and it will be conducive to adjustment and optimization of orbit control strategy in follow-up missions. 相似文献
6.
Rodney L. Anderson Jeffrey S. Parker 《Celestial Mechanics and Dynamical Astronomy》2013,115(3):311-331
In this study, transfer trajectories from the Earth to the Moon that encounter the Moon at various flight path angles are examined, and lunar approach trajectories are compared to the invariant manifolds of selected unstable orbits in the circular restricted three-body problem. Previous work focused on lunar impact and landing trajectories encountering the Moon normal to the surface, and this research extends the problem with different flight path angles in three dimensions. The lunar landing geometry for a range of Jacobi constants is computed, and approaches to the Moon via invariant manifolds from unstable orbits are analyzed for different energy levels. 相似文献
7.
Understanding plume impingement by retrorockets on the surface of the Moon is paramount for safe lunar outpost design in NASA’s planned return to the Moon for the Constellation Program. Visual inspection, Scanning Electron Microscopy, and surface scanned topology have been used to investigate the damage to the Lunar Surveyor III spacecraft that was caused by the Apollo 12 Lunar Module’s close proximity landing. Two parts of the Surveyor III craft returned by the Apollo 12 astronauts, Coupons 2050 and 2051, which faced the Apollo 12 landing site, show that a fine layer of lunar regolith coated the materials and was subsequently removed by the Apollo 12 Lunar Module landing rocket. The coupons were also pitted by the impact of larger soil particles with an average of 103 pits/cm2. The average entry size of the pits was 83.7 μm (major diameter) × 74.5 μm (minor diameter) and the average estimated penetration depth was 88.4 μm. Pitting in the surface of the coupons correlates to removal of lunar fines and is likely a signature of lunar material imparting localized momentum/energy sufficient to cause cracking of the paint. Comparison with the lunar soil particle size distribution and the optical density of blowing soil during lunar landings indicates that the Surveyor III spacecraft was not exposed to the direct spray of the landing Lunar Module, but instead experienced only the fringes of the spray of soil. Had Surveyor III been exposed to the direct spray, the damage would have been orders of magnitude higher. 相似文献
8.
The American Lunar Reconnaissance Orbiter spacecraft acquired high-resolution images of the landing sites of the Apollo manned spaceships and the Luna automatic space probes. In the images taken with the LROC Narrow-Angle Camera, the traces of anthropogenic influence on the lunar surface are seen in these places. However, such traces are not always noticeable sufficiently well, since they are masked by inhomogeneities in the brightness of the examined surface region caused by its topographic features and albedo variations. To increase the potential of identifying the disturbances of the initial structure of the lunar surface, the data should be analyzed with so-called phase-ratio imaging. Its essence is that the ratio of two coinciding images of the same surface region obtained at different phase angles is calculated. This method was applied to the analysis of the landing site of the Soviet Luna-17 space probe that transported the Lunokhod-1 rover to the lunar surface. The structural disturbance caused by the impact of jet flows from the probe’s engines and the tracks of the Lunokhod-1 wheels, which are faintly discernible in the usual images, has been detected. 相似文献
9.
From the lunar surface survey performed with a narrow-angle camera of the Lunar Reconnaissance Orbiter (LRO) spacecraft, the
distributions of the phase ratios of the Apollo 11 and 12 landing sites and the Ranger 9 impact site were mapped. In the acquired images, the traces of the structural disturbances of the lunar regolith layer caused
by the jet flows are seen. In the Ranger 9 impact site, one can see the crater of about 15 m across with a ray system, which is hardly noticeable in the brightness
picture, but has a high contract in the phase ratio picture. The character of the photometric anomaly of the rays of this
crater shows that they are formed by the ejected stones composing the rugged relief, which induces a strong shadow effect.
At the same time, the influence of jet flows from the rocket engines smooths the relief and leads to the photometric anomaly
of the opposite sign. The estimate of the maturity degree of the lunar regolith in the Apollo 11 and 12 landing sites obtained from the SELENE spectral survey suggests that the depth of the influence of the rocket engines on
the soil is small, and the surface of the impact crater formed by the Ranger 9 spacecraft contains a large amount of the immature soil. 相似文献
10.
W. E. Williamson 《Celestial Mechanics and Dynamical Astronomy》1972,5(2):174-188
The perturbation method, a numerical method for solving two point boundary value problems (TPBVP), is modified to attempt to improve inherent instability and sensitivity problems associated with the method. The desired solution to the TPBVP is divided into two time intervals. The differential equations required to define a solution to the two point boundary value problem are integrated independently over these shorter segments rather than consecutively over the entire trajectory. The independent integration of the differential equations over approximately half of the trajectory instead of the entire trajectory substantially decreases sensitivity and stability properties associated with the numerical integration. The equations for both time segments can be integrated simultaneously. By this procedure, a system of twice the dimension of the original problem is integrated for a period of time equal to half of the time interval for the original problem. To show the effectiveness of the method, two impulse trajectories which minimize the total velocity increment required to transfer a spacecraft from an Earth orbit into a lunar orbit are calculated. 相似文献
11.
Riccardo Marson Mauro Pontani Ettore Perozzi Paolo Teofilatto 《Celestial Mechanics and Dynamical Astronomy》2010,106(2):117-142
The possibility of communicating with the far side of the Moon is essential for keeping a continuous radio link with lunar
orbiting spacecraft, as well as with manned or unmanned surface facilities in locations characterized by poor coverage from
Earth. If the exploration and the exploitation of the Moon must be sustainable in the medium/long term, we need to develop
the capability to realize and service such facilities at an affordable cost. Minimizing the spacecraft mass and the number
of launches is a driving parameter to this end. The aim of this study is to show how Space Manifold Dynamics can be profitably
applied in order to launch three small spacecraft onboard the same launch vehicle and send them to different orbits around
the Moon with no significant difference in the Delta-V budgets. Internal manifold transfers are considered to minimize also
the transfer time. The approach used is the following: we used the linearized solution of the equations of motion in the Circular
Restricted Three Body Problem to determine a first–guess state vector associated with the Weak Stability Boundary regions,
either around the collinear Lagrangian point L1 or around the Moon. The resulting vector is then used as initial state in
a numerical backward-integration sequence that outputs a trajectory on a manifold. The dynamical model used in the numerical
integration is four-body and non-circular, i.e. the perturbations of the Sun and the lunar orbital eccentricity are accounted
for. The trajectory found in this way is used as the principal segment of the lunar transfer. After separation, with minor
maneuvers each satellite is injected into different orbits that lead to ballistic capture around the Moon. Finally, one or
more circularization maneuvers are needed in order to achieve the final circular orbits. The whole mission profile, from launch
to insertion into the final lunar orbits, is modeled numerically with the commercial software STK. 相似文献
12.
The problem of optimizing the interplanetary trajectories of a spacecraft (SC) with a solar electric propulsion system (SEPS) is examined. The problem of investigating the permissible power minimum of the solar electric propulsion power plant required for a successful flight is studied. Permissible ranges of thrust and exhaust velocity are analyzed for the given range of flight time and final mass of the spacecraft. The optimization is performed according to Portnyagin’s maximum principle, and the continuation method is used for reducing the boundary problem of maximal principle to the Cauchy problem and to study the solution/ parameters dependence. Such a combination results in the robust algorithm that reduces the problem of trajectory optimization to the numerical integration of differential equations by the continuation method. 相似文献
13.
Yu. G. Shkuratov M. A. Kreslavsky D. G. Stankevich V. G. Kaydash P. Pinet V. V. Shevchenko B. H. Foing J.-L. Josset 《Solar System Research》2003,37(4):251-259
We describe the future SMART-1 European Space Mission whose objective is to study the lunar surface from a polar lunar orbit. In particular, it is anticipated that selected regions of the Moon will be photographed using the AMIE camera with a mean spatial resolution of about 100 m in three spectral channels (0.75, 0.92, and 0.96 m) over a wide range of phase angles. Since these spectral channels and the AMIE resolution are close to those of the UVVIS camera onboard the Clementine spacecraft, the simultaneous processing of SMART-1 and Clementine data can be planned, for example, to obtain phase-ratio images. These images carry information on the structural features of the lunar surface. In particular, UVVIS/Clementine data revealed a photometric anomaly at the Apollo-15 landing site associated with the blowing of the lunar regolith by the lander engine. Anomalies were found in the ejection zones of several fresh craters. 相似文献
14.
Golubev Yu. F. Tuchin A. G. Grushevskii A. V. Koryanov V. V. Tuchin D. A. Morskoy I. M. Simonov A. V. Dobrovolskii V. S. 《Solar System Research》2016,50(7):597-603
The development of a methodology for designing trajectories of spacecraft intended for the contact and remote studies of Jupiter and its natural satellites is considered. This methodology should take into account a number of specific features. Firstly, in order to maintain the propellant consumption at an acceptable level, the flight profile, ensuring the injection of the spacecraft into orbit around the Jovian moon, should include a large number of gravity assist maneuvers both in the interplanetary phase of the Earth-to-Jupiter flight and during the flight in the system of the giant planet. Secondly, the presence of Jupiter’s powerful radiation belts also imposes fairly strict limitations on the trajectory parameters.
相似文献15.
This article continues our study of spacecraft guidance and control for a soft Moon landing (see our article “Main braking phase for a soft Moon landing as a form of trajectory correction”). Rationale is given for the objectives of the subsequent (final touchdown) phases. Analytical relations for the main parameters are obtained, and the impact of various disturbing factors is estimated. A methodology is proposed for calculating the main parameters for the whole braking sequence from the sighting altitude of the main braking phase termination to braking engine thrust and its throttle range. 相似文献
16.
F. J. T. Salazar C. F. de Melo E. E. N. Macau O. C. Winter 《Celestial Mechanics and Dynamical Astronomy》2012,114(1-2):201-213
An alternative transfer strategy to send spacecraft to stable orbits around the Lagrangian equilibrium points L4 and L5 based in trajectories derived from the periodic orbits around L1 is presented in this work. The trajectories derived, called Trajectories G, are described and studied in terms of the initial generation requirements and their energy variations relative to the Earth through the passage by the lunar sphere of influence. Missions for insertion of spacecraft in elliptic orbits around L4 and L5 are analysed considering the restricted three-body problem Earth–Moon-particle and the results are discussed starting from the thrust, time of flight and energy variation relative to the Earth. 相似文献
17.
Analysis of the gravity gradiometer developed by R. L. Forward and C. C. Bell at the Hughes Research Laboratories suggest than an accuracy, in the range 0.1 to 0.5 EU can be expected in a lunar orbiter application. This accuracy will allow gradient anomalies associated with mascons to be mapped with 1% accuracy and should reveal a great deal of new information about the lunar gravity field.The proposed experiment calls for putting such a gradiometer into a closely circular polar orbit at an average height of about 30 km above the lunar surface. This orbit allows the entire lunar surface to be covered in fourteen days, the gradiometer to be checked twice per revolution and results in successive passes above the lunar surface being spaced at about the resolution limit of about 30 km set both by the satellite altitude and instrumental integration time. Doppler tracking will be employed and the spacecraft will carry an electromagnetic altimeter. Gradient and altitude data from the far side of the Moon can be stored for replay when communication is re-established. 相似文献
18.
Michael Dellnitz Sina Ober-Blöbaum Marcus Post Oliver Schütze Bianca Thiere 《Celestial Mechanics and Dynamical Astronomy》2009,105(1-3):33-59
In this article, we introduce a novel three-step approach for solving optimal control problems in space mission design. We demonstrate its potential by the example task of sending a group of spacecraft to a specific Earth L 2 halo orbit. In each of the three steps we make use of recently developed optimization methods and the result of one step serves as input data for the subsequent one. Firstly, we perform a global and multi-objective optimization on a restricted class of control functions. The solutions of this problem are (Pareto-)optimal with respect to ΔV and flight time. Based on the solution set, a compromise trajectory can be chosen suited to the mission goals. In the second step, this selected trajectory serves as initial guess for a direct local optimization. We construct a trajectory using a more flexible control law and, hence, the obtained solutions are improved with respect to control effort. Finally, we consider the improved result as a reference trajectory for a formation flight task and compute trajectories for several spacecraft such that these arrive at the halo orbit in a prescribed relative configuration. The strong points of our three-step approach are that the challenging design of good initial guesses is handled numerically by the global optimization tool and afterwards, the last two steps only have to be performed for one reference trajectory. 相似文献
19.
Fen-Fei Wang Jian-Jun Liu Chun-Lai Li Xin Ren Ling-Li Mu Wei Yan Wen-Rui Wang Jing-Tao Xiao Xu Tan Xiao-Xia Zhang Xiao-Duan Zou Xing-Ye Gao 《中国天文和天体物理学报》2014,(12)
The establishment of a lunar control network is one of the core tasks in selenodesy, in which defining an absolute control point on the Moon is the most important step. However, up to now, the number of absolute control points has been very sparse. These absolute control points have mainly been lunar laser ranging retroreflectors, whose geographical location can be observed by observations on Earth and also identified in high resolution lunar satellite images. The Chang'e-3(CE-3) probe successfully landed on the Moon, and its geographical location has been monitored by an observing station on Earth. Since its positional accuracy is expected to reach the meter level, the CE-3 landing site can become a new high precision absolute control point. We use a sequence of images taken from the landing camera, as well as satellite images taken by CE-1 and CE-2, to identify the location of the CE-3 lander. With its geographical location known, the CE-3 landing site can be established as a new absolute control point, which will effectively expand the current area of the lunar absolute control network by 22%, and can greatly facilitate future research in the field of lunar surveying and mapping, as well as selenodesy. 相似文献
20.
B Kazeminejad M Pérez-Ayúcar M Sanchez-Nogales N Strange L Popken P Couzin 《Planetary and Space Science》2004,52(9):799-814
The Cassini spacecraft will arrive at Saturn in 2004 carrying the Huygens probe. The beginning of the Cassini tour at Saturn has been redesigned to achieve a different relative orbiter/probe geometry in order to compensate for the probe relay receiver design flaw that was discovered during tests in February 2000. This paper presents a numerical simulation of the Huygens atmospheric entry and descent trajectory and the Cassini flyby trajectory during the probe mission. A variety of parameters that are crucial for the probe system and its scientific payload have been calculated and analyzed together with an assessment of their uncertainties. Furthermore the orbiter/probe relay link was simulated in order to assess any potential data loss on the basis of an analytical model of the actual Huygens receiver onboard the Cassini spacecraft. The redesigned Cassini/Huygens mission satisfies all science and engineering requirements and assures the best possible radio link for the entire nominal mission duration. 相似文献