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

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.

Stereoscopic observations from meteorological satellites

Reliability testing of modern tantalum capacitors shows the presence of both types, infant mortality (IM) and wear-out (WO) failures. To assure reliable operation, the probability of IM failures should be reduced to below the specified level, and the time of WO inception should be greater than the required useful life of the parts at operation conditions. In this work, several types of MnO2 and polymer cathode tantalum capacitors have been tested at highly accelerated conditions to assess voltage acceleration factors, determine adequate burn-in conditions, and assess the useful life. A modified time dependent dielectric breakdown (TDDB) model has been suggested to simulate both types of failure. A physical mechanism of degradation that is based on migration and reactions of oxygen vacancies explains increase of the defect related, IM failures with the level of stress.

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 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.

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

Orbital debris minimization and mitigation techniques

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.

Currently, most Landsat satellites are deployed in the low earth orbit (LEO) to obtain high resolution data of the Earth surface and atmosphere. However, the return channels of LEO satellites are unstable and discontinuous intrinsically resulting from the high orbital velocity, long revisit interval and limited ranges of ground-based radar receivers. Space-based information networks (SBIN), in which data can be delivered by the cooperative transmission of relay satellites, can greatly expand the spatial transport connection ranges of LEO satellites. While different types of these relay satellites deployed in orbits of different altitudes represent distinctive performances when they are participating in forwarding. In this chapter, we consider the cooperative mechanism of relay satellites deployed in the geosynchronous orbit (GEO) and LEO according to their different transport performances and orbital characteristics. To take full advantage of the transmission resource of different kinds of cooperative relays, we propose a multiple access and bandwidth resource allocation strategy for GEO relay, in which the relay can receive and transmit simultaneously according to channel characteristics of space-based systems. Moreover, a time slot allocation strategy which is based on the slotted time division multiple access is introduced for the system with LEO relays. Based on the queueing theoretic formulation, the stability of the proposed systems and protocols is analyzed and the maximum stable throughput region is derived as well, which provide the guidance for the design of the system optimal control. Simulation results exhibit multiple factors that affect the stable throughput and verify the theoretical analysis.KeywordsSpace-based Information NetworksCooperative CommunicationResource AllocationQueueing TheoryStability Analysis

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:: ~ •• ~: ••• . -

Impact of the orbital eccentricity on the attitude performance before and after the deorbiting phase for Alsat-1

Space-based remote sensing of the Earth is conducted from a fleet of spacecraft in two basic orbital positions, near-polar low-Earth orbits and geosynchronous orbits, with each offering its own advantages and disadvantages. Low-Earth orbits provide high-resolution observations at the expense of large-scale contextual information, while geosynchronous orbits provide near-global, continuous coverage at reduced resolutions. However, due to the rapidly decreasing horizontal resolution data-products derived from geosynchronous orbits are of degraded value beyond approximately 55 degrees of latitude. A novel mission design is introduced to enable continuous observation of all longitudes at latitudes between 55 and 90 degrees with an observation zenith angle of less than 60 degrees, without the use of composite images. A single Soyuz launch is used to deliver three spacecraft to 12-hr, highly eccentric true-polar orbits with apogee at 40170 km and electric propulsion is used to maintain the orbit apse-line coincident with the Earth’s poles. It is shown that the science payload mass can be traded against the mission duration, with a payload mass varying between 120 – 90 kg for mission durations between 3 – 5 years, respectively. It is further shown that the payload would have approximately of 2kW of power available during operations as the electric propulsion system is not operated at these times. Whilst the payload mass is less than a typical remote sensing platform in geosynchronous orbit it is considered that the concept would offer an excellent technology demonstrator mission for operational missions, whilst also enabling unique and valuable science.

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.

Multi-scale analysis of mechanical properties and failure behavior of SiCf/Ti composite thin-walled tubes under compression loading

This paper examines the problem of detecting and classifying objects in Earth orbit using a space-augmented space surveillance network (SA-SSN). A SA-SSN uses a combination of ground- and space-based sensors to monitor activities over a range of space orbits from low earth orbits up to an altitude higher than the geosynchronous orbit. We develop a cost-aware Bayesian risk analysis approach for object detection and classification, using range-angle sensors with intermittent information-sharing between the sensors. The problem is formulated in a simplified two-dimensional setting where the SA-SSN is composed of four ground-based sensor and a space-based orbiting sensor satellite. This is done in order to reduce computational complexity while retaining the basic nontrivial elements of the problem. We will demonstrate that objects in geosynchronous orbits can be detected and perfectly classified (under appropriate sensor models) if they intermittently cross the field of view of some sensor in the SA-SSN, and that performance degrades for objects located in non-geosynchronous orbits. We will conclude the paper with future research directions on how to address the detection and classification of objects in non-geosynchronous orbits.