Problems of design and development of advanced superheavy launch vehicles

The article analyzes problems of design and development of advanced superheavy launch vehicles. Mass and energy characteristics and design layout of launch vehicles are substantiated. Delivery methods of bulky superheavy launch vehicle components to the spacecraft launch site are discussed. Methods of reduction of financial and technical risks of development and operation of superheavy launch vehicles are analyzed. The problem of environmental impacts of superheavy launch vehicle launches is posed.     

A Concept for Designing a System for a Double-Launch Spacecraft System as Part of the Upper Stages of the Fregat

Solar System Research - Abstract—A concept for designing a double-launch system (SDL) for the spacecraft SDZ-La2 onboard the Fregat upper stage and the launch vehicle Soyuz-2 and SDZ-La5...     

Ballistic representation for kinematic access

This work uses simple two-body orbital dynamics to initially determine the kinematic access for a ballistic vehicle. Primarily this analysis was developed to assess when a rocket body might conjunct with an orbiting satellite platform. A family of access opportunities can be represented as a volume for a specific rocket relative to its launch platform. Alternately, the opportunities can be represented as a geographical footprint relative to aircraft or satellite position that encompasses all possible launcher locations for a specific rocket. A thrusting rocket is treated as a ballistic vehicle that receives all its energy at launch and follows a coasting trajectory. To do so, the rocket’s burnout energy is used to find its equivalent initial velocity for a given launcher’s altitude. Three kinematic access solutions are then found that account for spherical Earth rotation. One solution finds the maximum range for an ascent-only trajectory while another solution accommodates a descending trajectory. In addition, the ascent engagement for the descending trajectory is used to depict a rapid access scenario. These preliminary solutions are formulated to address ground-, sea-, or air-launched vehicles.     

Limited By Cost: The Case Against Humans In The Scientific Exploration Of Space

Human space flight represents a heady mix of bravery and drama which can be inspirational to nations and to humankind but at huge economic cost. Due to the current high launch costs only a handful of people have ventured beyond low Earth orbit and walked on the Moon, propelled by aspirations related more to the Cold War than to science. Problems with reusable launch vehicle development mean that severe launch cost limitations will exist for some time. Meanwhile, cheaper robotic probes have visited all the planets except Pluto, flown by comets, landed on Mars, Venus and an asteroid, have probed Jupiter's atmosphere and studied the Universe beyond our own solar system with telescopes. Using these data we are determining mankind's place in the Universe. Public interest in the historic Eros landing eclipsed a simultaneous space walk at the fledgling International Space Station and the Mars Pathfinder landing generated hundreds of millions of website hits in a few days. Given the fact that hundreds of Mars missions could be flown for the still-escalating cost of the International Space Station, the unsuitability of human bodies for deep space exploration, and the advances in 3-d and virtual reality techniques, we discuss whether human exploration needs a place in a realistic, useful and inspirational space programme. This revised version was published online in July 2006 with corrections to the Cover Date.     

Implementation of the Concept for Creating a Spacecraft Double-Launch System as a Part of the Upper Stages of the Fregat Family

Solar System Research - Abstract—Results of the implementation of a single concept for the development of the double launch system (SDZ) are presented, in particular the SDZ-La5 for the...     

Prospective spacecraft for venus research: Venera-D design

A new phase of Venus research has started in Russia; the Federal Space Program includes the Venera-D design with the launch of the spacecraft scheduled for 2016. The mission comprises an orbiter, a descent vehicle, and balloon probes. The balloon probes will be placed at different altitudes in the cloud layer and under the clouds, where they are intended to last for a long time in the atmosphere of Venus. The successful implementation of the design will allow solving of quite a number of scientific tasks for comparative planetology.     

Finding KBO Flyby Targets for New Horizons

Development of the New Horizons mission to Pluto and the Kuiper Belt is now fully funded by NASA (Stern and Spencer, this volume). If all goes well, New Horizons will be launched in January 2006, followed by a Jupiter gravity assist in 2007, with Pluto arrival expected in either 2015 or 2016, depending on the launch vehicle chosen. A backup launch date of early 2007, without a Jupiter flyby, would give a Pluto arrival in 2019 or 2020. In either case, a flyby of at least one Kuiper Belt object (KBO) is planned following the Pluto encounter, sometime before the spacecraft reaches a heliocentric distance of 50 AU, in 2021 or 2023 for the 2006 launch, and 2027 or 2029 for the 2007 launch. However, none of the almost 1000 currently-known KBOs will pass close enough to the spacecraft trajectory to be targeted by New Horizons, so the KBO flyby depends on finding a suitable target among the estimated 500,000 KBOs larger than 40 km in diameter. This paper discusses the issues involved in finding one or more KBO targets for New Horizons. The New Horizons team plans its own searches for mission KBOs but will welcome other U.S, or international team who wish to become involved in exchange for mission participation at the KBO.     

QSAT: The Satellite for Polar Plasma Observation

This paper introduces QSAT, the satellite for polar plasma observation. The QSAT project began in 2006 as an initiative by graduate students of Kyushu University, and has the potential to contribute greatly to IHY (International Heliophysical Year) by showing to the world the beauty, importance, and relevance of space science. The primary objectives of the QSAT mission are (1) to investigate plasma physics in the Earth’s aurora zone in order to better understand spacecraft charging, and (2) to conduct a comparison of the field-aligned current observed in orbit with ground-based observations. The QSAT project can provide education and research opportunities for students in an activity combining space sciences and satellite engineering. The QSAT satellite is designed to be launched in a piggyback fashion with the Japanese launch vehicle H-IIA. The spacecraft bus is being developed at the Department of Aeronautics and Astronautics of Kyushu University with collaboration of Fukuoka Institute of Technology. Regarding the payload instruments, the Space Environment Research Center of Kyushu University is developing the magnetometers, whereas the Laboratory of Spacecraft Environment Interaction Engineering of Kyushu Institute of Technology is developing the plasma probes. We aim to be ready for launch in 2009 or later.     

Non-numeric computation for high eccentricity orbits

Geocentric orbits of large eccentricity (e=0.9 to 0.95) are significantly perturbed in cislunar space by the Sun and Moon. The time-history of the height of perigee, subsequent to launch, is particularly critical. The determination of ‘launch windows’ is mostly concerned with preventing the height of perigee from falling below its low initial value before the mission lifetime has elapsed. Between the extremes of high accuracy digital integration of the equations of motion and of using an approximate, but very fast, stability criteria method, this paper is concerned with the development of a method of intermediate complexity using non-numeric computation. The computer is used as the theory generator to generalize Lidov's theory using six osculating elements. Symbolic integration is completely automatized and the output is a set of condensed formulae well suited for repeated applications in launch window analysis. Examples of applications are given.     

Earth-Based Support for the Titan Saturn System Mission

The Titan Saturn System Mission (TSSM) concept is composed of a TSSM orbiter provided by NASA that would carry two Titan in situ elements provided by ESA: the montgolfière and the probe/lake lander. One overarching goal of TSSM is to explore in situ the atmosphere and surface of Titan. The mission has been prioritized as the second Outer Planets Flagship Mission, the first one being the Europa Jupiter System Mission (EJSM). TSSM would launch around 2023–2025 arriving at Saturn 9 years later followed by a 4-year science mission in the Saturn system. Following delivery of the in situ elements to Titan, the TSSM orbiter would explore the Saturn system via a 2-year tour that includes Enceladus and Titan flybys before entering into a dedicated orbit around Titan. The Titan montgolfière aerial vehicle under consideration will circumnavigate Titan at a latitude of ~20° and at altitudes of ~10 km for a minimum of 6 months. The probe/lake lander will descend through Titan’s atmosphere and land on the liquid surface of Kraken Mare (~75° north latitude). As for any planetary space science mission, and based on the Cassini–Huygens experience, Earth-based observations will be synergistic and enable scientific optimization of the return of such a mission. Some specific examples of how this can be achieved (through VLBI and Doppler tracking, continuous monitoring of atmospheric and surface features, and Direct-to-Earth transmission) are described in this paper.     

Design and analysis of Weak Stability Boundary trajectories to Moon

An algorithm is developed to find Weak Stability Boundary transfer trajectories to Moon in high fidelity force model using forward propagation. The trajectory starts from an Earth Parking Orbit (circular or elliptical). The algorithm varies the control parameters at Earth Parking Orbit and on the way to Moon to arrive at a ballistic capture trajectory at Moon. Forward propagation helps to satisfy launch vehicle’s maximum payload constraints. Using this algorithm, a number of test cases are evaluated and detailed analysis of capture orbits is presented.     

Lunar Gravity-Assist Maneuver As a Way of Reducing the Orbit Amplitude in the Spectrum–Röntgen–Gamma Project

Spectrum–Röntgen–Gamma (SRG) is a space observatory designed to observe astrophysical objects in the X-ray range of the electromagnetic spectrum. SRG is planned to be launched in 2019 by a Proton-M launch vehicle with a DM3 upper stage. The spacecraft will be delivered to an orbit around the Sun–Earth collinear libration point L2 located at a distance of ~1.5 million km from the Earth. Although the SRG launch scheme has already been determined at present, in this paper we consider an alternative spacecraft transfer scenario using a lunar gravity-assist maneuver. The proposed scenario allows a oneimpulse transfer from a low Earth orbit to a small-amplitude orbit around the libration point to be performed while fulfilling the technical constraints and the scientific requirements of the mission.     

BepiColombo, ESA's Mercury Cornerstone mission

The paper presents the results of the definition studies performed for the European Space Agency (ESA) on system architectures and enabling technologies for “BepiColombo”, a Cornerstone class mission to be launched in the 2007–2009 time frame. The scientific mission comprises 1-year observations by a Mercury Planetary Orbiter (MPO), dedicated to remote sensing, and a Mercury Magnetospheric Orbiter (MMO), dedicated to particles and fields, plus short-duration in situ analysis by a Mercury surface element (MSE). A flexible approach to the programme has been developed, comprising two alternative launch scenarios. In the first option (2009), the 2500-kg class satellite composite, including two propulsion modules and three scientific modules, is launched by an Ariane-5. The trajectory design is based on Venus and Mercury gravity assists plus the thrust provided by a Solar Electric Propulsion Module (SEPM), that is jettisoned before being captured into Mercury orbit. Capture and orbit insertion, executed by successive manoeuvres of a Chemical Propulsion Module (CPM), occur less than 2.5 yr after launch. In the second scenario, the mission is split into two launches of a small launch vehicle. Two 1200-kg class composites are launched either in the same one-month window or at an interval of 1.6 yr. One composite comprises the SEPM, CPM, MMO and MSE and the other comprises duplicate SEPM+CPM and the MPO. The trajectory design follows the same principles as the Ariane-5 mission, with the SEPM thrust reduced by half and cruise duration ranging between 2.3 and 3.5 yr. Whatever be the implementation, the mission is expected to return about 1700 Gbit of scientific data during the one-year observation phase. The crucial aspects of the spacecraft design are associated with, and constrained by, the high-temperature and high-radiation environment. Basic feasibility has been demonstrated by an extensive design and analysis exercise, and the focus of the programme has now moved to a 3-year preparatory programme dedicated for developing the enabling technologies.     

Using space manifold dynamics to deploy a small satellite constellation around the Moon

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.     

An “ESA-Affordable” Laue-lens

With ESA’s INTEGRAL mission gamma-ray astronomy has advanced to the point where major scientific advances must be expected from detailed studies of the many new point sources. The interest in developing focusing telescopes operating in the soft gamma-ray regime up to 1 MeV is therefore mounting rapidly. Telescopes based on Laue diffraction of gamma-rays from crystals appear as one promising route, although the practical difficulties of realizing a large scale Laue lens are certainly not small. In this paper I have attempted to develop an optimized lens design considering the size and mass constraints of a specific medium size launch vehicle. The introduction of the lens mass as a primary design driver has some surprising effects for the choice of material for the crystals and new tradeoff considerations are introduced.     

Description of the trajectory functions of TOPEX/POSEIDON navigation

TOPEX/POSEIDON is a joint American/French ocean topography experiment. It was launched by an Ariane launch vehicle on August 10, 1992 to study and map ocean circulation. The primary functions of the navigation subsystem of the TOPEX/POSEIDON project are to establish and maintain a pre-designed reference orbit, and to measure, monitor, and predict the satellite ground track continuously. To fulfill these functions, trajectory analysis is required to design and generate all trajectory related products. This paper is concerned with the trajectory functions of TOPEX/POSEIDON navigation. It describes various activities of this support function.     

ADVANTAGES OF SEARCHING FOR ASTEROIDS FROM LOW EARTH ORBIT: THE NEOSSat MISSION

Space-based observatories have several advantages over ground-based observatories in searching for asteroids and comets. In particular, the Aten and Interior to Earth’s Orbit (IEO) asteroid classes may be efficiently sought at low solar elongations along the ecliptic plane. A telescope in low Earth orbit has a sufficiently long orbital baseline to determine the parallax for all Aten and IEO class asteroids discovered with this observing strategy. The Near Earth Object Space Surveillance Satellite (NEOSSat) mission will launch a microsatellite to exploit this observing strategy complementing ground-based search programmes.     

First light for the sodium laser guide star adaptive optics system on the Lijiang 1.8m telescope

A first generation sodium Laser Guide Star Adaptive Optics System(LGS-AOS) was developed and integrated into the Lijiang 1.8 m telescope in 2013. The LGS-AOS has three sub-systems:(1) a 20 W long pulsed sodium laser,(2) a 300-millimeter-diameter laser launch telescope, and(3) a 37-element compact adaptive optics system. On 2014 January 25, we obtained high resolution images of an mV8.18 star,HIP 43963, during the first light of the LGS-AOS. In this paper, the sodium laser, the laser launch telescope,the compact adaptive optics system and the first light results will be presented.     

Launch window study for the highly eccentric orbit satellite HEOS-1

A satellite with a high eccentricitye0.95 is strongly perturbed by the sun and the moon. This fact and mission constraints restrict considerably the possible launch times for such a satellte. The launch window calculations can be performed in two steps in order to save computing time. An approximate analytical solution provides a general survey of the launch opportunities. An accurate numerical approach is then necessary for the exact definition of the launch window. In the case of the orbit of HEOS-1 (satellite 68 10901), moreover the consideration of the injection errors has been of great importance.Presented at the Conference on Celestial Mechanics, Oberwolfach, Germany, August 17–23, 1969.     

The Soviet ASTRON mission: legacy

2013 marks the 30th anniversary since the launch of Soviet Spacecraft Astron that had been operated for 6 years as the largest ultraviolet telescope during its lifetime. The Astron orbital station was designed for the astrophysical observations. It was launched into orbit by Proton launch system on March 23, 1983. Astron had a 80 cm ultraviolet telescope with mass of 400 kg and a complex of X-ray spectrographs with mass of 300 kg on board as a payload. It’s high apogee orbit (with apogee 200000 km and perigee 2000 km) permitted the influences of the Earth’s umbra and radiation belts to be excluded from the measurements. The main astrophysical results are summarized in this paper.