Cooperative Resource Allocation in Heterogeneous Space-Based Networks

Recently, wireless multimedia services have been growing because of the spread of various wireless applications. Hence, time division duplex (TDD) systems and their crossed-slot interference problems in a cellular environment have been attracting growing interests. Considering large signaling overhead between cells, decentralized time slot allocation (TSA) strategy is suitable for practical implementation. Thus, Fixed-TSA strategy which fixes the same ratio of uplink and downlink time slots in all cells is adopted in commercial WiBro systems (IEEE802.16e Mobile WiMax) to mitigate the crossed-slot interference. However, the Fixed-TSA strategy reduces the flexibility in time slot allocation due to a strong constraint set by a predefined boundary. Moreover, the mathematical derivation of the optimal value for the predefined boundary has not been presented yet. In this paper, we propose two decentralized TSA strategies. The first one is Enhanced Fixed-TSA strategy, which dynamically adapts the predefined boundary of the conventional Fixed-TSA strategy according to traffic conditions. The second one is region-based Decentralized Time Slot Allocation (RED-TSA) strategy, which utilizes partial location information of MSs to reduce crossedslot interference. We fully analyze the proposed TSA strategies in terms of new call blocking probability, average bit error probability, and the overall system throughput. Numerical results show that the proposed RED-TSA strategy provides the highest system throughput by compromising both new call blocking performance and average bit error performance, whereas it requires additional location information and complicated computation. On the contrary, the proposed Enhanced Fixed-TSA strategy requires reasonable computational complexity while it provides almost the same system throughput with that of the proposed RED-TSA strategy. In a nutshell, the proposed Enhanced Fixed- TSA strategy is more appropriate for the practical systems. Moreover, it can be directly applied to the current commercial WiMax/WiBro systems with minimum changes.

Spacecraft attitude control subsystem: Reliability, multi-state analyses, and comparative failure behavior in LEO and GEO

We analyze spectral efficiency of dynamic time division duplex in a fixed wireless cellular network. Conventionally, spectral efficiency has been analyzed for static and fully loaded systems. In this paper, we investigate how asymmetric and dynamic traffic affect the spectral efficiency of Time-Division-Duplex (TDD) systems. Recently, dynamic TDD (D-TDD) has gained much attention as an efficient duplex scheme for high-speed data communications, because adaptive switching ability enables the system to obtain statistical multiplexing gain by exploiting dynamic and asymmetric data traffic. However, it has been noted that a rather strong co-channel interference can be present due to adaptive switching in a cellular network that uses frequency reuse. Thus, benefits of dynamic TDD may not be justifiable unless a proper countermeasure is employed. To suppress the effect of strong co-channel interference, we employ time slot allocation (TSA) strategy along with sector antenna layouts, as proposed in our previous work Jeong and Kavehrad (IEEE and Transactions on communication, Vol. 50, no.10 pp. 1627–1636, 2002). We note that higher spectral efficiency is obtained in D-TDD systems by employing TSA strategy. We also evaluated spectral efficiency of D-TDD system employing adaptive modulation, assuming that traffic is delay tolerant. It is observed that five times higher spectral efficiency can be obtained by employing adaptive modulation. The effect of variance of ratio of offered load on uplink and downlink is also evaluated. Our computer simulation results show that spectral efficiency of D-TDD system with time slot allocation algorithm is more than that of static TDD (S-TDD) over a large range of traffic variation. In conclusion, D-TDD system can take advantage of statistical multiplexing gain of dynamic traffic by adaptively positioning the boundary to the varying traffic bandwidth in its two-way transmissions when TSA strategy is employed for suppression of strong co-channel interference.

A strawman crystal-silicon 5-GW satellite power system (SPS) concept formed the basis of a study of construction concepts for building a complete SPS in low Earth orbit (LEO) or geosynchronous orbit (GEO). Construction scenarios were evolved, including factory-in-space concepts and operations. Design implications imposed on the SPS satellite as a consequence of in-orbit assembly operations, and related attitude control requirements during assembly in LEO or GEO environments, were also evaluated. Results are presented indicating that complete assembly of an operational SPS in LEO, followed by transport to GEO, does not appear technically desirable. The best mix, however, of LEO vs GEO construction activity remains to be resolved.

Gravity field recovery of inter-satellite links between Beidou navigation satellite system (BDS) and LEO based on geodesy and time reference in space (GETRIS)

The improvement of BDS Observation Geometry with LEO constellations in Orbit Determination

Stereoscopic observations from meteorological satellites

Today most operational Earth observing satellites reside in low Earth orbits (LEO) at less than 1,000 km altitude, and in geostationary Earth orbits (GEO) at ~35,800 km altitude. These orbits have been the venues of choice for observations, albeit for very different reasons. LEO provides high spatial resolution with low temporal resolution while GEO provides for low spatial resolution, but high temporal resolution. NOAA utilizes both venues for their environmental satellites. The NOAA Polar-orbiting Operational Environmental Satellites (POES) reside in LEO Sun synchronous orbits at approximately 830 km in altitude, as do the Defense Meteorological Satellite Program (DMSP) satellites of the Department of Defense. In the near future the POES and DMSP satellites will be merged into a new satellite system referred to as the National Polar-orbiting Operational Environmental Satellite System (NPOESS). The NOAA Geostationary Operational Environmental Satellite (GOES) system, as the name specifies, resides at the other preferred observational venue of GEO. The Jet Propulsion Laboratory (JPL), under contract to NOAA, has been studying the characteristics of medium Earth orbits (MEO), at altitudes between 1000 and 35,800 km, as an observation venue to answer the question as to whether MEO might capture the attributes of the two traditional venues. This on-going study initially focused on determining the optimal altitude for MEO observations, through numerous trade studies involving altitude, instrument complexity, coverage, radiation environment, data temporality, revisit time, data rates, downlink requirements and other parameters including cost and launch complexity. Once the optimal altitude of 10,400 km had been determined the study proceeded to explore single through multiple MEO satellite constellation performance capabilities using two instrument types, a visible through infrared (IR) imager and IR sounder as the satellites’ payload. The MEO performance capabilities were compared to comparable LEO and GEO satellite constellation capabilities. This portion of the study concluded that indeed for global coverage a constellation of satellites operating in the MEO venue could capture the attributes of those operating in the LEO and GEO venues. Three 8-satellite constellations configurations - Walker, ICO, and Equatorial-Polar (EP) - then were studied to develop more constellation coverage statistics including robustness to individual satellite failure. That study phase concluded that the EP constellation was superior to both the ICO and Walker configurations. The study is presently examining if, and to what extent, the equatorial portion of the EP constellation might provide substantive supplemental data to that collected by the NPOESS and GOES satellite constellations.

Low earth orbit (LEO) satellite system becomes an emerging non-terrestrial network solution for future radio communications. To maximize the system capacity, an LEO satellite could utilize circularly polarized planar antenna array with the frequency reuse technique to form multiple earthmoving beams. With the increasing number of LEO satellites, inline interference from the geosynchronous equatorial orbit (GEO) satellite is an important issue to be resolved when the spectrum resource is shared. With the frequency hopping phenomena in earth-moving LEO networks, the inline interference could be minimized if the GEO users also perform the similar frequency hopping via coordination between LEO and GEO satellite networks. This paper proposes eight spectrum sharing schemes for GEO satellite networks to produce the frequency hopping sequence in earth-moving LEO network. We evaluate the system capacity, fairness between LEO and GEO users, and time costs of all schemes. Simulation results validate the advantage obtained from coordination, and show that the simple greedy scheme performs as good as the optimal scheme due to fixed frequency hopping patterns.

Orbital debris minimization and mitigation techniques

The paper investigates the utility of how nuclear thermal rocket (NTR) upper stage propulsion could substantially reduce the interplanetary transit time while greatly enhancing the economics of performing the manned Mars mission. au ECCV GTO IMLEO Isp kg kg/kw km km/sec kw LEO MEV mw NTR sec !N Nomenclature astronomical unit Earth capture crew vehicle geosynchronous transfer orbit initial mass in low Earth orbit specific impulse kilogram kilograms per kilowatt kilometer kilometers per second kilowatt low Earth orbit Mars excursion vehicle megawatt nuclear thermal rocket seconds delta V 1. Background Initial responses to the launch cost issue have been to suppose the development of heavy lift launch vehicles that are assumed to have fundamental launch cost savings. In general, most studies have isolated launch vehicle systems from the interplanetary vehicle assembled in low Earth orbit. This paper investigates a potential synergism between applying nuclear thermal rockets as both upper stages of the launch vehicle and as departure propulsion for a Mars vehicle departure stage. II. InterPlanetary Vehicle Description The mission to Mars is a propulsion technology dominated mission. The velocity change requirements (llV or delta V) are very large and very sensitive to trip time constraints. Due to this, there has been a heavy emphasis on high specific impulse (Isp) propulsion systems for performing the manned missions to Mars. The fundamental problem of high Isp systems is that they are generally thermally limited by material temperature limits. This issue is eliminated by electric propulsion systems that use electro-magnetic forces for accelerating a working fluid, but are thrust limited due to practical limitations in power supply mass. More exotic propulsion systems tend to use plasma working fluids that can only be contained by such methods as magnetic nozzles. The Space Exploration Initiative (SET) has a long range goal to perform a manned mission to Mars. Preliminary design goals of an interplanetary vehicle for this mission stress extremely high propulsion requirements to minimize transit times while allowing significant stay times on Mars. Related to this key trade between propulsion performance and trip time is the space transportation economics of 20 or more Earth launches required of the current launch vehicle architecture to place adequate vehicle mass in low Earth orbit to perform a single manned mission to Mars. The large initial mass in low Earth orbit (IMLEO) to perform the manned Mars mission is recognized as a dominant mission cost driver using the current launch infrastructure. IMLEO has been used as a preliminary figure of merit to evaluate various design approaches. * Senior Advanced Propulsion Engineer, Member AIAA Most analysis to date has focused on nuclear thermal rockets (Isp ... 1000 sec) and megawatt class electric propulsion systems (Isp ... 5000+ sec). The vehicle strategies have been to split the mission between a one way only cargo mission to Mars and a fast transfer manned vehicle. Mars rendezvous is essential for mission success, and techniques to slow at Earth return have used Earth capture crew vehicles (ECCV) with very high excess velocities (9+ km/sec) that are ballistically braked. Analyzed here is a very simple single vehicle. It uses an NTR driven, high thrust, Earth departure stage that is dropped once the vehicle has entered a fast Mars transfer ellipse. Once the vehicle has exited the deep Earth gravity well, it relies solely on a 5000 second Isp electric propulsion system for capture at Mars, departure from Mars, and final mission capture at Mars with zero excess velocity (a conservative assumption) The payload to Mars is broken into two elements. The ECCV is assumed to weigh 40,000 kg for a crew of four, and is the crew's interplanetary travel habitation module. The Mars excursion vehicle (MEV) is the element used to drop the crew from Mars orbit to the planet surface, be the crew's planetary habitation module, carry all science equipment, and carry the ascent module for returning to Mars orbit. The MEV has a baseline weight of 84,000 kg for a nominal 30 day planetary stay time and is scaled with the squareroot of normalized stay time. The MEV is not returned to Earth. Consumables are estimated at 10 kg per person per mission day. optimized for minimum l1V by adjusting the up leg perihelion and aphelion while maintaining the constraints of Earth rendezvous and travel time. Figure 2 (with the radius in millions of kilometers) shows a transfer ellipse for the up leg for the 320 day travel time, 80 day Mars stay time, 0 degree mean Mars true anomaly case. Figure 3 shows the down leg of that mission. The mission l1V requirement of 29.1 krn/sec is divided into 4.2 krn/sec for Earth departure (in heliocentric space), 6.0 krn/sec for Mars capture, 7.4 krn/sec for Mars departure, and 11.5 krn/sec for Earth capture. Transfer Ellipse Up. III. Interplanetary Mission Modeling The problem of analyzing the interplanetary mission can be driven by very exacting issues such as actual planetary configurations for particular mission opportunities, the requirements for travel time, and the Mars stay time. The issue that needs near term resolution, however, is the general efficacy of various propulsion and/or vehicle configuration options for the Mars mission in general. To answer these type of questions, it was necessary to generalize the Mars transfer problem. To do so, simplifying assumptions were that the Mars and Earth orbits were coplanar, and that the Earth orbit was circular. The nominal planetary configurations for Mars, Earth, and Venus are shown in Figure 1. The radius units are in astronomical units (Earth radius =1 au) and the angular units are the true anomaly of the orbital positions in degrees (Mars perihelion = 0 degrees). tojMJoC:: ~ •• ~: ••• . -

The primary mission of the PSSC Testbed is to provide space flight data on two types of advanced solar cells, XTJ and ZTJ, produced by the two U.S. solar cell manufacturers, Spectrolab and Emcore, respectively. The PSSC Testbed flight data confirmed that the picosatellite can be used for measuring l-V characteristics of solar cells. Continuing monitoring of the l-V measurement will allow determining the actual degradation of the solar cells due to exposure to the space environment. Mission success for the initial low earth orbit (LEO) space flight provided proof of concept for the PSSC Testbed. Subsequent mission will fly in a geosynchronous transfer orbit (GTO), in which a 450-day radiation exposure of the solar cells will be equivalent to that of 15 years radiation exposure in geosynchronous orbit.

SummaryThis paper introduces architectures for next‐generation high throughput satellite (HTS) systems comprising various satellite payload options, ground terminal advances, and scalable system‐level software control and management techniques. It describes a model to estimate aggregate system capacity as a function of radio band, available spectrum, spot beams, waveforms, and payload capability, including antenna size, power, and digital/analog connectivity across various links and availability objectives. This system model has been used to evaluate aggregate capacity of representative Ka‐Band low earth orbit and geosynchronous orbit systems. A system implementation approach is described for next‐generation HTS systems based on widely used Industry standards. Modulation and coding techniques are based on Digital Video Broadcasting – S2 extensions (DVB‐S2X), which comprises spectrally efficient modulation schemes combined with low‐rate codes. Several implementation technologies are analyzed related to configurable onboard payload and ground‐based, software‐defined resource control and management, key enablers of next‐generation HTS systems. Basic architectural building blocks are introduced for design of end‐to‐end systems across low earth orbit, medium earth orbit, and geosynchronous orbit satellite constellations, with and without onboard processing and inter‐satellite links, and including several efficient scenarios to achieve lossless handovers. Copyright © 2016 John Wiley & Sons, Ltd.

It is noted that significant advances in laser communication technology can be utilized to provide low-risk, high-data-rate links from a low Earth orbit (LEO) to a geosynchronous Earth orbit (GEO) satellite or from a LEO satellite to the ground. The authors address the design of these links where data rates of 1 Gb/s and greater are required. The acquisition, the cooperative tracking, and the communication channel requirements, and implementation are developed and discussed, including atmospheric effects for a LEO satellite to ground link. Size, weight, and power estimates are included. Major components are identified. Overall system development risk is shown to be low, based on component maturity. >

To meet the ever-expanding needs of satellite customers, a new generation of launch vehicles in the Long-March (LM) family is being developed according to the principles of low-cost, high-reliability, and maximum flexibility. By reducing the pre-launch preparation time, the launch system will strengthen the capability of providing international commercial launch services in the future. The new generation launch vehicles consist of several standard modules featuring two types of newly developed engines and three new modules powered by non-toxic, non-polluting propellants. Among the three modules, 5-m module is powered by two 50-ton-thrust engines which burn liquid hydrogen (LH2) and liquid oxygen (LOX), the 3.35-m module is powered by two 120-ton-thrust engines which burn LOX and kerosene (KO), and the 2.25-m one is powered by one 120-ton-thrust engine. As a first step, a series of large-scale launch vehicles is formed with 5-m-dia common core. By modifying the modules, the next step is to combine medium launch vehicles with the 3.35-m-dia core and a small launch vehicle with the 2.25-m-dia core. Finally a new group of launch vehicles will be formed to accommodate a wide range of performance requirements, which provide a capability of delivering payloads ranging from 1.5 to14 tons to geosynchronous transfer orbit (GTO) and up to 25 tons to low earth orbit (LEO). This paper highlights the design principles, systems concept and a series of configurations for the new generation of the LM launch vehicles. Copyright  2003 by Beijing Institute of Astronautical Systems Engineering. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. Released to IAF/IAA/AIAA to publish in all forms. The Long-March family and its history The Long-March (LM) launch vehicle program was initiated in mid 1960s in China. In April 1970, the first launch vehicle LM-1 made its maiden flight and sent DFH-1 satellite into the earth orbit, which indicated that China entered the space era. In the past three decades, the LM family was developed rapidly and formed a series of launch vehicles consisting of LM-2C, LM-2E, LM-3, LM-3A, LM-3B, LM-4, etc. These launch vehicles are suitable for various tasks, such as LEO, SSO and GTO missions. The largest launch capacity of the LM family has reached 9,200 kg for LEO, and 5,200 kg for GTO, which basically meets the demands of various customers. Up to now, the LM family has accomplished 70 launches. From October 1996 to July 2003, 28 consecutive successful flights were made, which established a new record in the LM history. Requirements analysis Although LM family has made eye-catching achievements in many important fields of space transportation technology, the assessment of 1 54th International Astronautical Congress of the International Astronautical Federation, the International Academy of Astronautics, and the International Institute of Space Law 29 September 3 October 2003, Bremen, Germany IAC-03-V.1.04 Copyright © 2003 by the author(s). Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Released to IAF/IAA/AIAA to publish in all forms. launch services market clearly shows the need for the LM launch vehicles to be upgraded. − Lift-capability and flexibility need to be increased not only to match the continuing upward trend in terms of GEO communication satellites, but also enable new applications such as constellations, scientific missions and multiphase releases. − Present propellants have to be replaced by non-toxic, non-polluting ones in order to satisfy the demand of the environment protection. − Higher reliability and safety are required for commercial launches, especially for the manned mission. − Launch cost must be significantly reduced for the future market.