Selection and trajectory design to mission secondary targets

Recently, with new trajectory design techniques and use of low-thrust propulsion systems, missions have become more efficient and cheaper with respect to propellant. As a way to increase the mission’s value and scientific return, secondary targets close to the main trajectory are often added with a small change in the transfer trajectory. As a result of their large number, importance and facility to perform a flyby, asteroids are commonly used as such targets. This work uses the Primer Vector theory to define the direction and magnitude of the thrust for a minimum fuel consumption problem. The design of a low-thrust trajectory with a midcourse asteroid flyby is not only challenging for the low-thrust problem solution, but also with respect to the selection of a target and its flyby point. Currently more than 700,000 minor bodies have been identified, which generates a very large number of possible flyby points. This work uses a combination of reachability, reference orbit, and linear theory to select appropriate candidates, drastically reducing the simulation time, to be later included in the main trajectory and optimized. Two test cases are presented using the aforementioned selection process and optimization to add and design a secondary flyby to a mission with the primary objective of 3200 Phaethon flyby and 25143 Itokawa rendezvous.     

Mars reconnaissance lander: Vehicle and mission design

There is enormous potential for more mobile planetary surface science. This is especially true in the case of Mars because the ability to cross challenge terrain, access areas of higher elevation, visit diverse geological features and perform long traverses of up to 200 km supports the search for past water and life. Vehicles capable of a ballistic ‘hop’ have been proposed on several occasions, but those proposals using in-situ acquired propellants are the most promising for significant planetary exploration. This paper considers a mission concept termed Mars Reconnaissance Lander using such a vehicle. We describe an approach where planetary science requirements that cannot be met by a conventional rover are used to derive vehicle and mission requirements.The performance of the hopper vehicle was assessed by adding estimates of gravity losses and mission mass constraints to recently developed methods. A baseline vehicle with a scientific payload of 16.5 kg and conservatively estimated sub-system masses is predicted to achieve a flight range of 0.97 km. Using a simple consideration of system reliability, the required cumulative range of 200 km could be achieved with a probability of around 80%. Such a range is sufficient to explore geologically diverse terrains. We therefore plot an illustrative traverse in Hypanis Valles/Xanthe Terra, which encounters crater wall sections, periglacial terrain, aqueous sedimentary deposits and a traverse up an ancient fluvial channel. Such a diversity of sites could not be considered with a conventional rover. The Mars Reconnaissance Lander mission and vehicle presents some very significant engineering challenges, but would represent a valuable complement to rovers, static landers and orbital observations.     

A graph based methodology for mission design

A space mission design methodology is presented, where initial and final orbits are connected through segments of periodic orbits. After a discretization of the solution space, the problem of mission design is transformed into an equivalent combinatorial optimization problem. Specifically, a graph is constructed that represents periodic orbits connected by the execution of impulsive maneuvers. A low computational complexity algorithm for this transformation is introduced. An efficient combinatorial optimization algorithm that solves the shortest path problem is described. Subject to the initial discretization of the solution space, an optimal sequence of coastal arcs is determined for a low total Delta-V mission. Finally, the proposed methodology is applied to the design of a hypothetical Saturn?CTitan system mission.     

Odyssey: a solar system mission

The Solar System Odyssey mission uses modern-day high-precision experimental techniques to test the laws of fundamental physics which determine dynamics in the solar system. It could lead to major discoveries by using demonstrated technologies and could be flown within the Cosmic Vision time frame. The mission proposes to perform a set of precision gravitation experiments from the vicinity of Earth to the outer Solar System. Its scientific objectives can be summarized as follows: (1) test of the gravity force law in the Solar System up to and beyond the orbit of Saturn; (2) precise investigation of navigation anomalies at the fly-bys; (3) measurement of Eddington’s parameter at occultations; (4) mapping of gravity field in the outer solar system and study of the Kuiper belt. To this aim, the Odyssey mission is built up on a main spacecraft, designed to fly up to 13 AU, with the following components: (a) a high-precision accelerometer, with bias-rejection system, measuring the deviation of the trajectory from the geodesics, that is also giving gravitational forces; (b) Ka-band transponders, as for Cassini, for a precise range and Doppler measurement up to 13 AU, with additional VLBI equipment; (c) optional laser equipment, which would allow one to improve the range and Doppler measurement, resulting in particular in an improved measurement (with respect to Cassini) of the Eddington’s parameter. In this baseline concept, the main spacecraft is designed to operate beyond the Saturn orbit, up to 13 AU. It experiences multiple planetary fly-bys at Earth, Mars or Venus, and Jupiter. The cruise and fly-by phases allow the mission to achieve its baseline scientific objectives [(1) to (3) in the above list]. In addition to this baseline concept, the Odyssey mission proposes the release of the Enigma radio-beacon at Saturn, allowing one to extend the deep space gravity test up to at least 50 AU, while achieving the scientific objective of a mapping of gravity field in the outer Solar System [(4) in the above list].      

The lunar gravity mission MAGIA: preliminary design and performances

The importance of an accurate model of the Moon gravity field has been assessed for future navigation missions orbiting and/or landing on the Moon, in order to use our natural satellite as an intermediate base for next solar system observations and exploration as well as for lunar resources mapping and exploitation. One of the main scientific goals of MAGIA mission, whose Phase A study has been recently funded by the Italian Space Agency (ASI), is the mapping of lunar gravitational anomalies, and in particular those on the hidden side of the Moon, with an accuracy of 1 mGal RMS at lunar surface in the global solution of the gravitational field up to degree and order 80. MAGIA gravimetric experiment is performed into two phases: the first one, along which the main satellite shall perform remote sensing of the Moon surface, foresees the use of Precise Orbit Determination (POD) data available from ground tracking of the main satellite for the determination of the long wavelength components of gravitational field. Improvement in the accuracy of POD results are expected by the use of ISA, the Italian accelerometer on board the main satellite. Additional gravitational data from recent missions, like Kaguya/Selene, could be used in order to enhance the accuracy of such results. In the second phase the medium/short wavelength components of gravitational field shall be obtained through a low-to-low (GRACE-like) Satellite-to-Satellite Tracking (SST) experiment. POD data shall be acquired during the whole mission duration, while the SST data shall be available after the remote sensing phase, when the sub-satellite shall be released from the main one and both satellites shall be left in a free-fall dynamics in the gravity field of the Moon. SST range-rate data between the two satellites shall be measured through an inter-satellite link with accuracy compliant with current state of art space qualified technology. SST processing and gravitational anomalies retrieval shall benefit from a second ISA accelerometer on the sub-satellite in order to decouple lunar gravitational signal from other accelerations. Experiment performance analysis shows that the stated scientific requirements can be achieved with a low mass and low cost sub-satellite, with a SST gravimetric mission of just few months.     

A laser altimeter for BepiColombo mission: Instrument design and performance model

This paper presents a definition study of a laser altimeter for the topographic exploration of Mercury. The reference scenario is the BepiColombo mission, a cornerstone mission of European Space Agency (ESA) planned for 2012. BepiColombo will offer the chance to make a remarkable new contribution to our knowledge of the Solar System, by venturing into the hot region near the Sun and exploring Mercury, the most enigmatic of the earth's sisters among the terrestrial planets. First images of Mercury surface were acquired by Mariner 10 in 1974 and 1975 offering a coverage and resolution comparable to Earth-based telescopic coverage of the Moon before spaceflight. BepiColombo mission can be very beneficial by using an optical rangefinder for Mercury exploration. In fact starting from the first missions in 1970s until today, laser altimeters have been demonstrating to be particularly appropriate as part of the scientific payload whenever the topography of earth, lunar and planetary surface is the scientific objective of a space mission.Our system design is compliant to Mercury Polar Orbiter (MPO) of the mission. System performance analysis is carried out simulating main hermean topographic features and the potential targets on the planet by means of analytical models and computer codes and several plot are presented to analyse the performance of the instrument.     

Preferred design and error analysis for the future dedicated deep-space Mars-SST satellite gravity mission

Because the precise measurement of the Martian gravitational field plays a significant role in the future Mars exploration program, the future dedicated Mars satellite-to-satellite tracking (Mars-SST) gravity mission in China is investigated in detail for producing the next generation of the Mars gravity field model with high accuracy. Firstly, a new semi-numerical synthetical error model of the cumulative Martian geoid height influenced by the major error sources of the space-borne instruments is precisely established and efficiently verified. Secondly, the deep space network in combination with the satellite-to-satellite tracking in the low-low (DSN-SST-LL) mode is a preferred design owing to the high precision determination of the gravity maps, the low technical complexity of the satellite system and the successful experiences with the Earth’s Gravity Recovery and Climate Experiment (GRACE) projects and the lunar Gravity Recovery and Interior Laboratory (GRAIL) program. Finally, the future twin Mars-SST satellites plan to adopt the optimal matching accuracy indices of the satellite-equipped sensors (e.g., \(10^{-7}\) m/s in the inter-satellite range-rate from the interferometric laser ranging system (ILRS), 35 m in the orbital position tracked by the DSN and \(3\times 10^{-11}\) m/s² in the non-conservative force from the drag-free control system (DFCS)) and the preferred orbital parameters (e.g., the orbital altitude of \(100\pm 50\) km and the inter-satellite range of \(50\pm 10\) km).     

Design and on-orbit status of the trigger system for the DAMPE mission

DArk Matter Particle Explorer(DAMPE), the first Chinese astronomical satellite, was successfully launched at the Jiuquan Satellite Launch Center on 2015 Dec. 17. DAMPE consists of four subdetectors: Plastic Scintillator array Detector(PSD), Silicon-Tungsten tracKer-converter(STK), Bismuth Germanium Oxide(BGO) imaging calorimeter and NeUtron Detector(NUD). The global hardware trigger signal, which is generated by hits from the BGO calorimeter and the trigger logic board in the data acquisition system(DAQ), is responsible for event selection and DAQ synchronization of DAMPE. On orbit,to improve the detection efficiency, different trigger logics are used for event selection in different regions of latitude. The DAMPE trigger system compresses the average on-orbit trigger rate to 60 Hz and reduces science data mass to less than 13 GB per day to meet the requirement for the satellite's data link. The whole trigger system has run stably up to now, ensuring excellent on-orbit operation of DAMPE.     

The THESEUS space mission: science goals,requirements and mission concept

Experimental Astronomy - THESEUS, one of the two space mission concepts being studied by ESA as candidates for next M5 mission within its Comsic Vision programme, aims at fully exploiting Gamma-Ray...     

Optical design of visible emission line coronagraph on Indian space solar mission Aditya-L1

The ground based observations of the coronal emission lines using a coronagraph are affected by the short duration of clear sky and varying sky transparency. These conditions do not permit to study small amplitude variations in the coronal emission reliably necessary to investigate the process or processes involved in heating the coronal plasma and dynamics of solar corona. The proposed Visible Emission Line Coronagraph (VELC) over comes these limitations and will provide continuous observation 24 h a day needed for detailed studies of solar corona and drivers for space weather predictions. VELC payload onboard India’s Aditya-L1 space mission is an internally occulted solar coronagraph for studying the temperature, velocity, density and heating of solar corona. To achieve the proposed science goals, an instrument which is capable of carrying out simultaneous imaging, spectroscopy and spectro-polarimetric observations of the solar corona close to the solar limb is required. VELC is designed with salient features of (a) Imaging solar corona at 500 nm with an angular resolution of 5 arcsec over a FOV of 1.05Ro to 3Ro (Ro:Solar radius) (b) Simultaneous multi-slit spectroscopy at 530.3 nm [Fe XIV],789.2 nm [Fe XI] and 1074.7 nm [Fe XIII] with spectral dispersion of 28mÅ, 31mÅ and 202mÅ per pixel respectively, over a FOV of 1.05Ro to 1.5Ro. (c) Multi-slit dual beam spectro-polarimetry at 1074.7 nm. All the components of instrument have been optimized in view of the scientific objectives and requirements of space payloads. In this paper we present the details of optical configuration and the expected performance of the payload.     

The Solar-A mission

The Institute for Space and Astronautical Sciences (ISAS) is developing a satellite dedicated to high-energy observations of solar flares. The Solar-A will be launched in August–September, 1991, from the Kagoshima Space Center on board a M3S-II vehicle. The instrument complement emphasizes hard X-ray and soft X-ray imaging, and contains instruments supplied in part by U.S. and U.K. experimenters. This paper describes the instrumentation and the tentative observing program.     

The Solar-A mission

The Institute for Space and Astronautical Sciences (ISAS) is developing a satellite dedicated to high-energy observations of solar flares. The Solar-A will be launched in August–September, 1991, from the Kagoshima Space Center on board a M3S-II vehicle. The instrument complement emphasizes hard X-ray and soft X-ray imaging, and contains instruments supplied in part by U.S. and U.K. experimenters. This paper describes the instrumentation and the tentative observing program.

     

The INTEGRAL mission

INTEGRAL, the International Gamma-Ray Astrophysics Laboratory, to be launched in 2001, is the second medium-size scientific mission (M2) of the ESA long term programme Horizon 2000. INTEGRAL addresses the fine spectroscopy and accurate positioning of celestial gamma-ray sources in the energy range 10 keV to 10 MeV. The observational requirements will be met by a payload utilising coded mask imaging in combination with detector pixel arrays (Imaging) and cooled Germanium detectors (Spectroscopy). INTEGRAL is an ESA led mission in collaboration with Russia and USA. Most of the observing time will be made available to the general scientific community.This paper is largely based on the INTEGRAL Phase A study report (ESA SCI(93)1), written by the INTEGRAL Phase A Science Working Team: S. Bergeson-Willis, T.J.-L. Courvoisier, A.J. Dean, Ph. Durouchoux, B. McBreen, N. Eismont, N. Gehrels, J.E. Grindlay, W.A. Mahoney, J.L. Matteson, O. Pace, T.A. Prince, V. Schönfelder, G.K. Skinner, R. Sunyaev, B.N. Swanenburg, B.J. Teegarden, P. Ubertini, G. Vedrenne, G.E. Villa, S. Volonté, and C. Winkler.     

TandEM: Titan and Enceladus mission

TandEM was proposed as an L-class (large) mission in response to ESA’s Cosmic Vision 2015–2025 Call, and accepted for further studies, with the goal of exploring Titan and Enceladus. The mission concept is to perform in situ investigations of two worlds tied together by location and properties, whose remarkable natures have been partly revealed by the ongoing Cassini–Huygens mission. These bodies still hold mysteries requiring a complete exploration using a variety of vehicles and instruments. TandEM is an ambitious mission because its targets are two of the most exciting and challenging bodies in the Solar System. It is designed to build on but exceed the scientific and technological accomplishments of the Cassini–Huygens mission, exploring Titan and Enceladus in ways that are not currently possible (full close-up and in situ coverage over long periods of time). In the current mission architecture, TandEM proposes to deliver two medium-sized spacecraft to the Saturnian system. One spacecraft would be an orbiter with a large host of instruments which would perform several Enceladus flybys and deliver penetrators to its surface before going into a dedicated orbit around Titan alone, while the other spacecraft would carry the Titan in situ investigation components, i.e. a hot-air balloon (Montgolfière) and possibly several landing probes to be delivered through the atmosphere.     

Data acquisition, control, communication and computation system of solar X-ray spectrometer (SOXS) mission

The Solar X-ray Spectrometer (SOXS) mission onboard GSAT-2 Indian Spacecraft was launched on 08 May 2003 using GSLV-D2 rocket by Indian Space Research Organization (ISRO). SOXS aims to study solar flares, which are the most violent and energetic phenomena in the solar system, in the energy range of 4–56 keV with high spectral and temporal resolution. SOXS employs state-of-the-art semiconductor devices, viz., Si-Pin and CZT detectors to achieve sub-keV energy resolution requirements. In this paper, we present an overview of data acquisition, control, communication and computation of low energy payload of the SOXS mission.     

The Planck Surveyor mission and gravitational lenses

An extremely sensitive all-sky survey will be carried out in the millimetre/submillimetre waveband by the forthcoming ESA mission Planck Surveyor . The main scientific goal of the mission is to make very accurate measurements of the spatial power spectrum of primordial anisotropies in the cosmic microwave background radiation; however, hundreds of thousands of distant dusty galaxies and quasars will also be detected. These sources are much more likely to be gravitationally lensed by intervening galaxies compared with sources discovered in surveys in other wavebands. Here the number of lenses expected in the survey is estimated, and techniques for discriminating between lensed and unlensed sources are discussed. A practical strategy for this discrimination is presented, based on exploiting the remarkable sensitivity and resolving power of large ground-based millimetre/submillimetre-wave interferometer arrays. More than a thousand gravitational lenses could be detected: a sample that would be an extremely valuable resource in observational cosmology.     

Mariner 9 mission profile and project history

Mariner 9 was inserted into orbit about Mars on November 14, 1971, to study that planet for a period of 90 days. Observations and measurements made by Mariner 9 continued beyond the planned 90 days, providing data to meet all science objectives. The new knowledge of Mars gained from this mission has made obsolete all previous concepts of Mars.A general background of the Mariner Mars 1971 Project and the significant events of the Mariner 9 mission are presented.     

Galilean satellites and the Galileo space mission

The Galileo spacecraft arrived at Jupiter in December 1995 to start its two-year mission of exploring the Jovian system, The spacecraft will complete eleven orbits around Jupiter and have ten more close encounters with the outer three Galilean satellites, after the initial close approach to lo on December 7, 1995, Since the lo encounter occurred closer to lo than originally designed, the spacecraft energy change was greater than nominally planned and resulted in an initial spacecraft orbital period about 7 days less than that designed in the nominal tour, A 100-km change in the Io-encounter distance results in an 8-day change in initial period of the spacecraft. Hence the first Ganymede encounter was moved forward one week, and the aim points for the first two Ganymede encounters were altered, but all other encounters would occur on their nominal dates and at the nominal altitudes, This was accomplished without expending spacecraft fuel and resulted in the first Ganymede flyby occurring on June 27, 1996 rather than the nominally scheduled July 4.Earth- and spacecraft-based data were employed in developing ephemerides in support of the Galileo space mission. An analysis of CCD astrometric observations from 1992–1994, of photographic observations from 1967–1993, of mutual event astrometric data from 1973–1991, of Jovian eclipse timing data from 1652-1983, of Doppler data from 1987–1991, and of optical navigation data from the Voyager spacecraft encounter in 1979, produced the satellite ephemerides for the Galileo space mission.