Super NiCd Open-Circuit Storage and Low Earth Orbit (LEO) Life Test Evaluation

NASA Technical Reports Server (NTRS)

Baer, Jean Marie; Hwang, Warren C.; Ang, Valerie J.; Hayden, Jeff; Rao, Gopalakrishna; Day, John H. (Technical Monitor)

2002-01-01

This presentation discusses Air Force tests performed on super NiCd cells to measure their performance under conditions simulating Low Earth Orbit (LEO) conditions. Super NiCd cells offer potential advantages over existing NiCd cell designs including advanced cell design with improved separator material and electrode making processes, but handling and storage requires active charging. These tests conclude that the super NiCd cells support generic Air Force qualifications for conventional LEO missions (up to five years duration) and that handling and storage may not actually require active charging as previously assumed. Topics covered include: Test Plan, Initial Characterization Tests, Open-Circuit Storage Tests, and post storage capacities.

Aktiv De-Orbiting Onboard System from Leo of Upper Stages of Launchers

NASA Astrophysics Data System (ADS)

Trushlyakov, V.; Shalay, V.; Shatrov, J.; Jakovlev, M.; Kostantino, A.

2009-03-01

The active de-orbiting onboard system (VDOS) of upper separable parts (USP) stage of launchers from LEO into orbits of utilization with term of existence orbital lifetimes till 25 years is offered. ADOS it is based on use of power resources of not produced rests of liquid fuel onboard USP launchers with liquid propulsion module (LPM). Following systems enter in structure VDOS: the gas jet propulsion system consisting of a system of gasification, chambers of gas engines (GE), a control system. For gasification of the rests of liquid fuel the heat-carrier received in the autonomous gas generator is used. The gasification propellant components from each tank with temperature and the pressure determined by strength of the corresponding tank, move in chambers of the GE established on a top of a fuel compartment. After separation of a payload execute twist USP for preservation of its position in the space by activity of the GE. Ways of increase of a system effectiveness of gasification are offered by superposition on the entered heat-carrier of ultrasonic oscillations, and also introduction in gaseous fuel nanopowder of aluminum. The volume of adaptations of construction USP, connected with introduction VDOS does not exceed 5 % from weight of a dry construction.

Precise Orbit Determination Of Low Earth Satellites At AIUB Using GPS And SLR Data

NASA Astrophysics Data System (ADS)

Jaggi, A.; Bock, H.; Thaller, D.; Sosnica, K.; Meyer, U.; Baumann, C.; Dach, R.

2013-12-01

An ever increasing number of low Earth orbiting (LEO) satellites is, or will be, equipped with retro-reflectors for Satellite Laser Ranging (SLR) and on-board receivers to collect observations from Global Navigation Satellite Systems (GNSS) such as the Global Positioning System (GPS) and the Russian GLONASS and the European Galileo systems in the future. At the Astronomical Institute of the University of Bern (AIUB) LEO precise orbit determination (POD) using either GPS or SLR data is performed for a wide range of applications for satellites at different altitudes. For this purpose the classical numerical integration techniques, as also used for dynamic orbit determination of satellites at high altitudes, are extended by pseudo-stochastic orbit modeling techniques to efficiently cope with potential force model deficiencies for satellites at low altitudes. Accuracies of better than 2 cm may be achieved by pseudo-stochastic orbit modeling for satellites at very low altitudes such as for the GPS-based POD of the Gravity field and steady-state Ocean Circulation Explorer (GOCE).

Near-real time orbit determination for the GPS, CHAMP, GRACE, TerraSAR-X, and TanDEM-X satellites

NASA Astrophysics Data System (ADS)

Michalak, Grzegorz; Koenig, Rolf

The GFZ German Research Centre for Geosciences developed a near-real time (NRT) orbit gen-eration system for GPS and Low Earth Orbiting (LEO) satellites to support radio occultation data processing for the CHAMP, GRACE, Terra-SAR-X and the upcoming TanDEM-X mis-sions and fast baseline determination for the TanDEM-X mission. Precise NRT orbits are being generated for the CHAMP and GRACE-A satellites since August 2006 and for TerraSAR-X since August 2007. For each LEO, the system consists of three independent chains delivering NRT orbits with different latencies and accuracies. The first chain generates in a preceding step NRT GPS orbits and clock biases and based thereon LEO orbits with delays of 30 minutes counted from the last measurement point to the time the orbit product is available. The orbit accuracies can be assessed via Satellite Laser Ranging (SLR) to 7 cm. The second chain is based on predicted GPS orbits from the International GNSS Service (IGS) but endowed with in-house estimated clock biases. This chain generates orbits with the same latency of 30 minutes but with better accuracies of 5 cm SLR RMS. The third chain, the least accurate but the fastest, is based on predicted IGS GPS orbits and clocks and delivers LEO orbits with latencies of 13 minutes and accuracies of 10 cm SLR RMS. The system design is such that it can easily be extended to cope with new satellites like TanDEM-X requiring precise and fast available orbits.

Low-Earth-Orbit (LEO) Life Cycle Evaluation of Nickel-Zinc Batteries

NASA Technical Reports Server (NTRS)

Coates, D.; Ferreira, E.; Nyce, M.; Charkey, A.

1997-01-01

The conclusion of the Low-Earth-Orbit (LEO) life cycle evaluation of nickel-zinc batteries are: that composite nickel electrode provide excellent performance at a reduced weight and lower cost; calcium / zinc electrode minimizes shape change; unioptimized cell designs yield 60 Wh/kg; nickel-zinc delivers 600 cycles at 80% DOD; long cycle life obtainable at low DOD; high rate capability power density; long-term failure mechanism is stack dry; and anomalous overcharge (1120%) greatly affected cell performance but did not induce failure and was recoverable.

Accuracy of Satellite Optical Observations and Precise Orbit Determination

NASA Astrophysics Data System (ADS)

Shakun, L.; Koshkin, N.; Korobeynikova, E.; Strakhova, S.; Dragomiretsky, V.; Ryabov, A.; Melikyants, S.; Golubovskaya, T.; Terpan, S.

The monitoring of low-orbit space objects (LEO-objects) is performed in the Astronomical Observatory of Odessa I.I. Mechnikov National University (Ukraine) for many years. Decades-long archives of these observations are accessible within Ukrainian network of optical observers (UMOS). In this work, we give an example of orbit determination for the satellite with the 1500-km height of orbit based on angular observations in our observatory (Int. No. 086). For estimation of the measurement accuracy and accuracy of determination and propagation of satellite position, we analyze the observations of Ajisai satellite with the well-determined orbit. This allows making justified conclusions not only about random errors of separate measurements, but also to analyze the presence of systematic errors, including external ones to the measurement process. We have shown that the accuracy of one measurement has the standard deviation about 1 arcsec across the track and 1.4 arcsec along the track and systematical shifts in measurements of one track do not exceed 0.45 arcsec. Ajisai position in the interval of the orbit fitting is predicted with accuracy better than 30 m along the orbit and better than 10 m across the orbit for any its point.

Wholesale debris removal from LEO

NASA Astrophysics Data System (ADS)

Levin, Eugene; Pearson, Jerome; Carroll, Joseph

2012-04-01

Recent advances in electrodynamic propulsion make it possible to seriously consider wholesale removal of large debris from LEO for the first time since the beginning of the space era. Cumulative ranking of large groups of the LEO debris population and general limitations of passive drag devices and rocket-based removal systems are analyzed. A candidate electrodynamic debris removal system is discussed that can affordably remove all debris objects over 2 kg from LEO in 7 years. That means removing more than 99% of the collision-generated debris potential in LEO. Removal is performed by a dozen 100-kg propellantless vehicles that react against the Earth's magnetic field. The debris objects are dragged down and released into short-lived orbits below ISS. As an alternative to deorbit, some of them can be collected for storage and possible in-orbit recycling. The estimated cost per kilogram of debris removed is a small fraction of typical launch costs per kilogram. These rates are low enough to open commercial opportunities and create a governing framework for wholesale removal of large debris objects from LEO.

Radiation exposure and performance of multiple burn LEO-GEO orbit transfer trajectories

NASA Technical Reports Server (NTRS)

Gorland, S. H.

1985-01-01

Many potential strategies exist for the transfer of spacecraft from low Earth orbit (LEO) to geosynchronous (GEO) orbit. One strategy has generally been utilized, that being a single impulsive burn at perigee and a GEO insertion burn at apogee. Multiple burn strategies were discussed for orbit transfer vehicles (OTVs) but the transfer times and radiation exposure, particularly for potentially manned missions, were used as arguments against those options. Quantitative results concerning the trip time and radiation encountered by multiple burn orbit transfer missions in order to establish the feasibility of manned missions, the vulnerability of electronics, and the shielding requirements are presented. The performance of these multiple burn missions is quantified in terms of the payload and propellant variances from the minimum energy mission transfer. The missions analyzed varied from one to eight perigee burns and ranged from a high thrust, 1 g acceleration, cryogenic hydrogen-oxygen chemical prpulsion system to a continuous burn, 0.001 g acceleration, hydrogen fueled resistojet propulsion system with a trip time of 60 days.

Modeling of LEO Orbital Debris Populations in Centimeter and Millimeter Size Regimes

NASA Technical Reports Server (NTRS)

Xu, Y.-L.; Hill, . M.; Horstman, M.; Krisko, P. H.; Liou, J.-C.; Matney, M.; Stansbery, E. G.

2010-01-01

The building of the NASA Orbital Debris Engineering Model, whether ORDEM2000 or its recently updated version ORDEM2010, uses as its foundation a number of model debris populations, each truncated at a minimum object-size ranging from 10 micron to 1 m. This paper discusses the development of the ORDEM2010 model debris populations in LEO (low Earth orbit), focusing on centimeter (smaller than 10 cm) and millimeter size regimes. Primary data sets used in the statistical derivation of the cm- and mm-size model populations are from the Haystack radar operated in a staring mode. Unlike cataloged objects of sizes greater than approximately 10 cm, ground-based radars monitor smaller-size debris only in a statistical manner instead of tracking every piece. The mono-static Haystack radar can detect debris as small as approximately 5 mm at moderate LEO altitudes. Estimation of millimeter debris populations (for objects smaller than approximately 6 mm) rests largely on Goldstone radar measurements. The bi-static Goldstone radar can detect 2- to 3-mm objects. The modeling of the cm- and mm-debris populations follows the general approach to developing other ORDEM2010-required model populations for various components and types of debris. It relies on appropriate reference populations to provide necessary prior information on the orbital structures and other important characteristics of the debris objects. NASA's LEO-to-GEO Environment Debris (LEGEND) model is capable of furnishing such reference populations in the desired size range. A Bayesian statistical inference process, commonly adopted in ORDEM2010 model-population derivations, changes a priori distribution into a posteriori distribution and thus refines the reference populations in terms of data. This paper describes key elements and major steps in the statistical derivations of the cm- and mm-size debris populations and presents results. Due to lack of data for near 1-mm sizes, the model populations of 1- to 3.16-mm

Mission Analysis for LEO Microwave Power-Beaming Station in Orbital Launch of Microwave Lightcraft

NASA Technical Reports Server (NTRS)

Myrabo, L. N.; Dickenson, T.

2005-01-01

A detailed mission analysis study has been performed for a 1 km diameter, rechargeable satellite solar power station (SPS) designed to boost 20m diameter, 2400 kg Micr,oWave Lightcraft (MWLC) into low earth orbit (LEO) Positioned in a 476 km daily-repeating oi.bit, the 35 GHz microwave power station is configured like a spinning, thin-film bicycle wheel covered by 30% efficient sola cells on one side and billions of solid state microwave transmitter elements on the other, At the rim of this wheel are two superconducting magnets that can stor,e 2000 G.J of energy from the 320 MW, solar array over a period of several orbits. In preparation for launch, the entire station rotates to coarsely point at the Lightcraft, and then phases up using fine-pointing information sent from a beacon on-board the Lightcraft. Upon demand, the station transmits a 10 gigawatt microwave beam to lift the MWLC from the earth surface into LEO in a flight of several minutes duration. The mission analysis study was comprised of two parts: a) Power station assessment; and b) Analysis of MWLC dynamics during the ascent to orbit including the power-beaming relationships. The power station portion addressed eight critical issues: 1) Drag force vs. station orbital altitude; 2) Solar pressure force on the station; 3) Station orbital lifetime; 4) Feasibility of geo-magnetic re-boost; 5) Beta angle (i..e., sola1 alignment) and power station effective area relationship; 6) Power station percent time in sun vs, mission elapsed time; 7) Station beta angle vs.. charge time; 8) Stresses in station structures.. The launch dynamics portion examined four issues: 1) Ascent mission/trajecto1y profile; 2) MWLC/power-station mission geometry; 3) MWLC thrust angle vs. time; 4) Power station pitch rate during power beaming. Results indicate that approximately 0 58 N of drag force acts upon the station when rotated edge-on to project the minimum frontal area of 5000 sq m. An ion engine or perhaps an electrodynamic

Application of the Constrained Admissible Region Multiple Hypothesis Filter to Initial Orbit Determination of a Break-up

NASA Astrophysics Data System (ADS)

Kelecy, Tom; Shoemaker, Michael; Jah, Moriba

2013-08-01

A break-up in Low Earth Orbit (LEO) is simulated for 10 objects having area-to-mass ratios (AMR's) ranging from 0.1-10.0 m2/kg. The Constrained Admissible Region Multiple Hypothesis Filter (CAR-MHF) is applied to determining and characterizing the orbit and atmospheric drag parameters (CdA/m) simultaneously for each of the 10 objects with no a priori orbit or drag information. The results indicate that CAR-MHF shows promise for accurate, unambiguous and autonomous determination of the orbit and drag states.

Mapping the Space Radiation Environment in LEO Orbit by the SATRAM Timepix Payload On Board the Proba-V Satellite

NASA Astrophysics Data System (ADS)

Granja, Carlos; Polansky, Stepan; Sospisil, Stanislav; Owens, Alan; Mellab, Karim

2016-08-01

The compact spacecraft payload SATRAM is operating in LEO orbit since 2013 on board the Proba-V satellite from ESA and provides high-resolution wide-range radiation monitoring of the satellite environment. Equipped with the pixel detector Timepix, the technology demonstration payload determines the composition (particle types) and spectral characterization (stopping power) of the mixed radiation field with quantum imaging sensitivity, charged particle tracking, energy loss and directionality capability. With a polar orbit (sun synchronous, 98° inclination) and altitude of 820 km the space radiation field is continuously sampled over the entire planet every few days. Results are given in the form of spatial- and time- correlated maps of dose rate and particle flux. Comparison is made between quiescent and geomagnetic storm activity periods.

Computer modeling of high-voltage solar array experiment using the NASCAP/LEO (NASA Charging Analyzer Program/Low Earth Orbit) computer code

NASA Astrophysics Data System (ADS)

Reichl, Karl O., Jr.

1987-06-01

The relationship between the Interactions Measurement Payload for Shuttle (IMPS) flight experiment and the low Earth orbit plasma environment is discussed. Two interactions (parasitic current loss and electrostatic discharge on the array) may be detrimental to mission effectiveness. They result from the spacecraft's electrical potentials floating relative to plasma ground to achieve a charge flow equilibrium into the spacecraft. The floating potentials were driven by external biases applied to a solar array module of the Photovoltaic Array Space Power (PASP) experiment aboard the IMPS test pallet. The modeling was performed using the NASA Charging Analyzer Program/Low Earth Orbit (NASCAP/LEO) computer code which calculates the potentials and current collection of high-voltage objects in low Earth orbit. Models are developed by specifying the spacecraft, environment, and orbital parameters. Eight IMPS models were developed by varying the array's bias voltage and altering its orientation relative to its motion. The code modeled a typical low Earth equatorial orbit. NASCAP/LEO calculated a wide variety of possible floating potential and current collection scenarios. These varied directly with both the array bias voltage and with the vehicle's orbital orientation.

Wave optics-based LEO-LEO radio occultation retrieval

NASA Astrophysics Data System (ADS)

Benzon, Hans-Henrik; Høeg, Per

2016-06-01

This paper describes the theory for performing retrieval of radio occultations that use probing frequencies in the XK and KM band. Normally, radio occultations use frequencies in the L band, and GPS satellites are used as the transmitting source, and the occultation signals are received by a GPS receiver on board a Low Earth Orbit (LEO) satellite. The technique is based on the Doppler shift imposed, by the atmosphere, on the signal emitted from the GPS satellite. Two LEO satellites are assumed in the occultations discussed in this paper, and the retrieval is also dependent on the decrease in the signal amplitude caused by atmospheric absorption. The radio wave transmitter is placed on one of these satellites, while the receiver is placed on the other LEO satellite. One of the drawbacks of normal GPS-based radio occultations is that external information is needed to calculate some of the atmospheric products such as the correct water vapor content in the atmosphere. These limitations can be overcome when a proper selected range of high-frequency waves are used to probe the atmosphere. Probing frequencies close to the absorption line of water vapor have been included, thus allowing the retrieval of the water vapor content. Selecting the correct probing frequencies would make it possible to retrieve other information such as the content of ozone. The retrieval is performed through a number of processing steps which are based on the Full Spectrum Inversion (FSI) technique. The retrieval chain is therefore a wave optics-based retrieval chain, and it is therefore possible to process measurements that include multipath. In this paper simulated LEO to LEO radio occultations based on five different frequencies are used. The five frequencies are placed in the XK or KM frequency band. This new wave optics-based retrieval chain is used on a number of examples, and the retrieved atmospheric parameters are compared to the parameters from a global European Centre for Medium

A Dynamic/Anisotropic Low Earth Orbit (LEO) Ionizing Radiation Model

NASA Technical Reports Server (NTRS)

Badavi, Francis F.; West, Katie J.; Nealy, John E.; Wilson, John W.; Abrahms, Briana L.; Luetke, Nathan J.

2006-01-01

The International Space Station (ISS) provides the proving ground for future long duration human activities in space. Ionizing radiation measurements in ISS form the ideal tool for the experimental validation of ionizing radiation environmental models, nuclear transport code algorithms, and nuclear reaction cross sections. Indeed, prior measurements on the Space Transportation System (STS; Shuttle) have provided vital information impacting both the environmental models and the nuclear transport code development by requiring dynamic models of the Low Earth Orbit (LEO) environment. Previous studies using Computer Aided Design (CAD) models of the evolving ISS configurations with Thermo Luminescent Detector (TLD) area monitors, demonstrated that computational dosimetry requires environmental models with accurate non-isotropic as well as dynamic behavior, detailed information on rack loading, and an accurate 6 degree of freedom (DOF) description of ISS trajectory and orientation.

Updating the NASA LEO Orbital Debris Environment Model with Recent Radar and Optical Observations and in Situ Measurements

NASA Technical Reports Server (NTRS)

Liou, J.-C.; Anz-Meador, P.; Matney, M. J.; Kessler, D. J.; Theall, J.; Johnson, N. L.

2000-01-01

The Low Earth Orbit (LEO, between 200 and 2000 km altitudes) debris environment has been constantly measured by NASA Johnson Space Center's Liquid Mirror Telescope (LMT) since 1996 (Africano et al. 1999, NASA JSC-28826) and by Haystack and Haystack Auxiliary radars at MIT Lincoln Laboratory since 1990 (Settecerri et al. 1999, NASA JSC-28744). Debris particles as small as 3 mm can be detected by the radars and as small as 3 cm can be measured by LMT. Objects about 10 cm in diameter and greater are tracked and catalogued by the US Space Surveillance Network. Much smaller (down to several micrometers) natural and debris particle populations can be estimated based on in situ measurements, such as Long Duration Exposure Facility, and based on analyses of returned surfaces, such as Hubble Space Telescope solar arrays, European Retrievable Carrier, and Space Shuttles. To increase our understanding of the current LEO debris environment, the Orbital Debris Program Office at NASA JSC has initiated an effort to improve and update the ORDEM96 model (Kessler et al. 1996, NASA TM-104825) utilizing the recently available data. This paper gives an overview of the new NASA orbital debris engineering model, ORDEM2000.

Mapping the space radiation environment in LEO orbit by the SATRAM Timepix payload on board the Proba-V satellite

DOE Office of Scientific and Technical Information (OSTI.GOV)

Granja, Carlos, E-mail: carlos.granja@utef.cvut.cz; Polansky, Stepan

Detailed spatial- and time-correlated maps of the space radiation environment in Low Earth Orbit (LEO) are produced by the spacecraft payload SATRAM operating in open space on board the Proba-V satellite from the European Space Agency (ESA). Equipped with the hybrid semiconductor pixel detector Timepix, the compact radiation monitor payload provides the composition and spectral characterization of the mixed radiation field with quantum-counting and imaging dosimetry sensitivity, energetic charged particle tracking, directionality and energy loss response in wide dynamic range in terms of particle types, dose rates and particle fluxes. With a polar orbit (sun synchronous, 98° inclination) at themore » altitude of 820 km the payload samples the space radiation field at LEO covering basically the whole planet. First results of long-period data evaluation in the form of time-and spatially-correlated maps of total dose rate (all particles) are given.« less

Mapping the space radiation environment in LEO orbit by the SATRAM Timepix payload on board the Proba-V satellite

NASA Astrophysics Data System (ADS)

Granja, Carlos; Polansky, Stepan

2016-07-01

Detailed spatial- and time-correlated maps of the space radiation environment in Low Earth Orbit (LEO) are produced by the spacecraft payload SATRAM operating in open space on board the Proba-V satellite from the European Space Agency (ESA). Equipped with the hybrid semiconductor pixel detector Timepix, the compact radiation monitor payload provides the composition and spectral characterization of the mixed radiation field with quantum-counting and imaging dosimetry sensitivity, energetic charged particle tracking, directionality and energy loss response in wide dynamic range in terms of particle types, dose rates and particle fluxes. With a polar orbit (sun synchronous, 98° inclination) at the altitude of 820 km the payload samples the space radiation field at LEO covering basically the whole planet. First results of long-period data evaluation in the form of time-and spatially-correlated maps of total dose rate (all particles) are given.

NASA STD-4005: The LEO Spacecraft Charging Design Standard

NASA Technical Reports Server (NTRS)

Ferguson, Dale C.

2006-01-01

Power systems with voltages higher than about 55 volts may charge in Low Earth Orbit (LEO) enough to cause destructive arcing. The NASA STD-4005 LEO Spacecraft Charging Design Standard will help spacecraft designers prevent arcing and other deleterious effects on LEO spacecraft. The Appendices, an Information Handbook based on the popular LEO Spacecraft Charging Design Guidelines by Ferguson and Hillard, serve as a useful explanation and accompaniment to the Standard.

Satellite laser ranging to low Earth orbiters: orbit and network validation

NASA Astrophysics Data System (ADS)

Arnold, Daniel; Montenbruck, Oliver; Hackel, Stefan; Sośnica, Krzysztof

2018-04-01

Satellite laser ranging (SLR) to low Earth orbiters (LEOs) provides optical distance measurements with mm-to-cm-level precision. SLR residuals, i.e., differences between measured and modeled ranges, serve as a common figure of merit for the quality assessment of orbits derived by radiometric tracking techniques. We discuss relevant processing standards for the modeling of SLR observations and highlight the importance of line-of-sight-dependent range corrections for the various types of laser retroreflector arrays. A 1-3 cm consistency of SLR observations and GPS-based precise orbits is demonstrated for a wide range of past and present LEO missions supported by the International Laser Ranging Service (ILRS). A parameter estimation approach is presented to investigate systematic orbit errors and it is shown that SLR validation of LEO satellites is not only able to detect radial but also along-track and cross-track offsets. SLR residual statistics clearly depend on the employed precise orbit determination technique (kinematic vs. reduced-dynamic, float vs. fixed ambiguities) but also reveal pronounced differences in the ILRS station performance. Using the residual-based parameter estimation approach, corrections to ILRS station coordinates, range biases, and timing offsets are derived. As a result, root-mean-square residuals of 5-10 mm have been achieved over a 1-year data arc in 2016 using observations from a subset of high-performance stations and ambiguity-fixed orbits of four LEO missions. As a final contribution, we demonstrate that SLR can not only validate single-satellite orbit solutions but also precise baseline solutions of formation flying missions such as GRACE, TanDEM-X, and Swarm.

Transfers from Earth to LEO and LEO to interplanetary space using lasers

NASA Astrophysics Data System (ADS)

Phipps, Claude R.; Bonnal, Christophe; Masson, Fréderic; Boustie, Michel; Berthe, Laurent; Schneider, Matthieu; Baton, Sophie; Brambrink, Erik; Chevalier, Jean-Marc; Videau, Laurent; Boyer, Séverine A. E.

2018-05-01

New data on some materials at 80ps pulse duration and 1057 nm wavelength give us the option of proportionally combining them to obtain arbitrary values between 35 (aluminum) and 800 N/MW (POM, polyoxymethylene) for momentum coupling coefficient Cm. Laser ablation physics lets us transfer to LEO from Earth, or to interplanetary space using repetitively pulsed lasers and Cm values appropriate for each mission. We discuss practical results for lifting small payloads from Earth to LEO, and space missions such as a cis-Mars orbit with associated laser system parameters.

Passive Optical Link Budget for LEO Space Surveillance

NASA Astrophysics Data System (ADS)

Wagner, P.; Hasenohr, T.; Hampf, D.; Sproll, F.; Humbert, L.; Rodmann, J.; Riede, W.

The rising space debris population is becoming an increasing risk for space assets. Even objects with the size of 10mm can cause major damages to active spacecraft. Especially the orbits around 800km high are densely populated with space debris objects. To assess the risk of collisions with active satellites, the Earth orbits need to be surveyed permanently. Space debris laser ranging systems for example can deliver highly accurate positional data for precise orbit determination. Therefor a priori information about the objects coarse trajectory is needed. Such initial orbit information can be provided by wide angle optical sensors. The Institute of Technical Physics at the German Aerospace Center in Stuttgart runs an observatory to perform passive as well as laser optical measurements to LEO objects. In order to detect unknown objects, a wide-angle imaging system with a field of view in the range of 5° to 15° equipped with an astronomical CCD camera and a commercial off the shelf (COTS) lens was designed to continuously observe the night sky for LEO objects. This paper presents the passive optical link budget for observing LEO objects to show the benefits and limits of the physical performance of an optical surveillance system. A compact COTS system is able to detect objects with a couple of decimeters in size while a large aperture telescope can detect objects with diameters below 10cm. Additionally, data captured by a passive optical staring system with a 10 cm aperture was analyzed. It is shown that 90% of all objects with a radar cross section larger than 2m² are detected with such a system during twilight conditions. The smallest detected LEO object with this system has a size of 0.32m x 0.32m x 0.26m. These measurements are compared to the developed link budget which allows an estimation of the performance of larger systems.

LCTS on ALPHASAT and Sentinel 1a: in orbit status of the LEO to geo data relay system

NASA Astrophysics Data System (ADS)

Zech, H.; Heine, F.; Troendle, D.; Pimentel, P. M.; Panzlaff, K.; Motzigemba, M.; Meyer, R.; Philipp-May, S.

2017-11-01

The performance of sensors for Earth Observation Missions is constantly improving. This drives the need for a reliable, high-speed data transfer capability from a Low Earth Orbit (LEO) spacecraft (S/C) to ground. In addition, for the transfer of time-critical data to ground, a low latency between data generation in orbit and data reception at the respective mission control center is of high importance. Laser communication between Satellites for high data transmission in combination with a GEO data relay system for reducing the latency time addresses these requirements.

Tanker Argus: Re-supply for a LEO Cryogenic Propellant Depot

NASA Astrophysics Data System (ADS)

St. Germain, B.; Olds, J.; Kokan, T.; Marcus, L.; Miller, J.

The Argus reusable launch vehicle (RLV) concept is a single-stage-to-orbit conical, winged bodied vehicle powered by two liquid hydrogen/liquid oxygen supercharged ejector ramjets. The 3rd generation Argus launch vehicle utilizes advanced vehicle technologies along with a Maglev launch assist track. A tanker version of the Argus RLV is envisioned to provide an economical means of providing liquid fuel and oxidizer to an orbiting low-Earth orbit (LEO) propellant depot. This depot could then provide propellant to various spacecraft, including reusable orbital transfer vehicles used to ferry space solar power satellites to geo-stationary orbit. Two different tanker Argus configurations were analyzed. The first simply places additional propellant tanks inside the payload bay of an existing Argus reusable launch vehicle. The second concept is a modified Argus RLV in which the payload bay is removed and the vehicle propellant tanks are stretched to hold extra propellant. An iterative conceptual design process was used to design both Argus vehicles. This process involves various disciplines including aerodynamics, trajectory analysis, weights &structures, propulsion, operations, safety, and cost/economics. The payload bay version of tanker Argus, which has a gross mass of 256.3MT, is designed to deliver a 9.07MT payload to LEO. This payload includes propellant and the tank structure required to secure this propellant in the payload bay. The modified, pure tanker version of Argus has a gross mass of 218.6MT and is sized to deliver a full 9.07MT of propellant to LEO. The economic analysis performed for this study involved the calculation of many factors including the design/development and recurring costs of each vehicle. These results were used along with other economic assumptions to determine the "per kilogram" cost of delivering propellant to orbit. The results show that for a given flight rate the "per kilogram" cost is cheaper for the pure tanker version of Argus

Impedance measurements on a spiral-wound nickel/metal hydride cell cycled in a simulated Leo orbit

NASA Technical Reports Server (NTRS)

Reid, Margaret A.

1993-01-01

A spiral-wound size C cell was cycled at 25 C in a low earth orbit (LEO) regime at 50 percent depth of discharge (DOD) with approximately five percent over-charge. The nominal capacity was 3.5 AH. The cell was cycled for 2000 cycles. Capacity checks and impedance measurements over the complete range of state of charge were made upon receipt and after 500, 1000, and 2000 cycles. The capacity of the cell was essentially unchanged until after the impedance measurements at 2000 cycles. Only small changes in the impedance parameters were observed, but there was somewhat more scatter in the data after 2000 cycles. When the cell was returned to LEO cycling after 2000 cycles, only 38 percent of the capacity could be obtained. It is believed that the cell failed because of an equipment failure at the end of the final impedance measurements which allowed an over-discharge.

Measuring atmospheric density using GPS-LEO tracking data

NASA Astrophysics Data System (ADS)

Kuang, D.; Desai, S.; Sibthorpe, A.; Pi, X.

2014-01-01

We present a method to estimate the total neutral atmospheric density from precise orbit determination of Low Earth Orbit (LEO) satellites. We derive the total atmospheric density by determining the drag force acting on the LEOs through centimeter-level reduced-dynamic precise orbit determination (POD) using onboard Global Positioning System (GPS) tracking data. The precision of the estimated drag accelerations is assessed using various metrics, including differences between estimated along-track accelerations from consecutive 30-h POD solutions which overlap by 6 h, comparison of the resulting accelerations with accelerometer measurements, and comparison against an existing atmospheric density model, DTM-2000. We apply the method to GPS tracking data from CHAMP, GRACE, SAC-C, Jason-2, TerraSAR-X and COSMIC satellites, spanning 12 years (2001-2012) and covering orbital heights from 400 km to 1300 km. Errors in the estimates, including those introduced by deficiencies in other modeled forces (such as solar radiation pressure and Earth radiation pressure), are evaluated and the signal and noise levels for each satellite are analyzed. The estimated density data from CHAMP, GRACE, SAC-C and TerraSAR-X are identified as having high signal and low noise levels. These data all have high correlations with anominal atmospheric density model and show common features in relative residuals with respect to the nominal model in related parameter space. On the contrary, the estimated density data from COSMIC and Jason-2 show errors larger than the actual signal at corresponding altitudes thus having little practical value for this study. The results demonstrate that this method is applicable to data from a variety of missions and can provide useful total neutral density measurements for atmospheric study up to altitude as high as 715 km, with precision and resolution between those derived from traditional special orbital perturbation analysis and those obtained from onboard

Leo Commercial Market Projects (1997)

DOT National Transportation Integrated Search

1997-07-25

The Federal Aviation Administration's Associate Administrator for Commercial : Space Transportation (AST) has prepared projections of the low Earth orbit : (LEO) commercial payload and launch markets for the period between 1997 and : 2006. This asses...

Overview of battery usage in NASA/GSFC LEO and GEO missions

NASA Technical Reports Server (NTRS)

Yi, Thomas

1989-01-01

In July, 1989, Cosmic Background Explorer (COBE) will be launched from a Delta rocket to study the big bang theory. The COBE, which is in a LEO/Polar orbit, will have two 20 Ah NiCd batteries, and 18 cells per battery, made by McDonnell Douglas Company. In December, 1989, National Oceanic and Atmospheric Administration (NOAA-D) will be launched from an Atlas rocket for weather observation purposes. NOAA-D, which is in a LEO/Polar morning orbit, will have two 26.5 Ah NiCd batteries, and 17 cells per battery, made by Ge-Astro East Windor. NOAA-I, which is scheduled for May, 1991 launch in a LEO/Polar afternoon orbit, will have three 26.5 Ah NiCd batteries, 17 cells per battery, made by GE-Astro East Windor. In April, 1990, Gamma Ray Observatory (GRO) will be launched from STS37 to study the gamma ray radiation phenomenon. GRO, which is in a LEO orbit, will have two modular power systems (MPS) made by McDonnell Douglas, each MPS consisting of three 50 Ah NiCd batteries, 22 cells per battery. In July, 1990, Geostationary Operational Environmental Satellite (GOES-I) will be launched from an Atlas I rocket for weather observation purposes. GOES-I, which is in a GEO orbit, will have two 12 Ah NiCd batteries, 28 cells per battery, made by Ford Aerospace and Communications Company. In December, 1990, Tracking and Data Relay Satellite (TDRS-E) will be launched from STS43 for communication purposes. TDRS-E, which is in a GEO orbit, will have three 40 Ah NiCd batteries, 24 cells per battery, made by TRW. In August, 1991, Extreme Ultraviolet Explorer (EUVE) will be launched from a Delta rocket. EUVE, which is in a LEO orbit, will have one modular power system (MPS) made by McDonnell Douglas. In December, 1991, Upper Atmosphere Research Satellite (UARS) will be launched from STS50 to study the Earth's ozone layer and other environmental concerns. UARS, which is in a 56 deg inclination LEO orbit, will have one modular power systems (MPS) made by McDonnell Douglas.

Leo Satellite Communication through a LEO Constellation using TCP/IP Over ATM

NASA Technical Reports Server (NTRS)

Foore, Lawrence R.; Konangi, Vijay K.; Wallett, Thomas M.

1999-01-01

The simulated performance characteristics for communication between a terrestrial client and a Low Earth Orbit (LEO) satellite server are presented. The client and server nodes consist of a Transmission Control Protocol /Internet Protocol (TCP/IP) over ATM configuration. The ATM cells from the client or the server are transmitted to a gateway, packaged with some header information and transferred to a commercial LEO satellite constellation. These cells are then routed through the constellation to a gateway on the globe that allows the client/server communication to take place. Unspecified Bit Rate (UBR) is specified as the quality of service (QoS). Various data rates are considered.

Optical Observation of LEO Debris Caused by Feng Yun 1C

NASA Astrophysics Data System (ADS)

Kurosaki, Hirohisa; Yanagisawa, Toshifumi; Nakajima, Atsushi

Many pieces of space debris are in low earth orbit (LEO), and may be a serious problem in the near future. They are very hazardous to spacecraft such as the ISS, in which humans stay for long periods. In January 2007, China performed an experimental destruction of the meteorological satellite FengYun-1C in low earth orbit using a ballistic missile. Optical instruments for space debris observation were installed on Mt. Nyukasa in Nagano Prefecture by JAXA, and the resulting low earth orbit debris was observed with the small telescope there. We have developed an image-processing technique, the line-detection method, to extract such effects as the streaks created by meteors, LEO satellites, and LEO debris. We succeeded in detecting the trajectories of specified FengYun-1C debris whose TLE were known. In this paper, the detection and observation of low earth orbit debris are discussed.

Leo Commercial Market Projections (1996)

DOT National Transportation Integrated Search

1996-04-05

The Office of the Associate Administrator for Commercial Space Transportation OCST) of the Federal Aviation Administration (FAA) has prepared the attached projections of the Low Earth Orbit (LEO) commercial payload and launch markets for the period b...

An electrically conductive thermal control surface for spacecraft encountering Low-Earth Orbit (LEO) atomic oxygen indium tin oxide-coated thermal blankets

NASA Technical Reports Server (NTRS)

Bauer, J. L.

1987-01-01

An organic black thermal blanket material was coated with indium tin oxide (ITO) to prevent blanket degradation in the low Earth orbit (LEO) atomic oxygen environment. The blankets were designed for the Galileo spacecraft. Galileo was initially intended for space shuttle launch and would, therefore, have been exposed to atomic oxygen in LEO for between 10 and 25 hours. Two processes for depositing ITO are described. Thermooptical, electrical, and chemical properties of the ITO film are presented as a function of the deposition process. Results of exposure of the ITO film to atomic oxygen (from a shuttle flight) and radiation exposure (simulated Jovian environment) are also presented. It is shown that the ITO-protected thermal blankets would resist the anticipated LEO oxygen and Jovian radiation yet provide adequate thermooptical and electrical resistance. Reference is made to the ESA Ulysses spacecraft, which also used ITO protection on thermal control surfaces.

Determining the Rotation Periods of an Inactive LEO Satellite and the First Korean Space Debris on GEO, KOREASAT 1

NASA Astrophysics Data System (ADS)

Choi, Jin; Jo, Jung Hyun; Kim, Myung-Jin; Roh, Dong-Goo; Park, Sun-Youp; Lee, Hee-Jae; Park, Maru; Choi, Young-Jun; Yim, Hong-Suh; Bae, Young-Ho; Park, Young-Sik; Cho, Sungki; Moon, Hong-Kyu; Choi, Eun-Jung; Jang, Hyun-Jung; Park, Jang-Hyun

2016-06-01

Inactive space objects are usually rotating and tumbling as a result of internal or external forces. KOREASAT 1 has been inactive since 2005, and its drift trajectory has been monitored with the optical wide-field patrol network (OWL-Net). However, a quantitative analysis of KOREASAT 1 in regard to the attitude evolution has never been performed. Here, two optical tracking systems were used to acquire raw measurements to analyze the rotation period of two inactive satellites. During the optical campaign in 2013, KOREASAT 1 was observed by a 0.6 m class optical telescope operated by the Korea Astronomy and Space Science Institute (KASI). The rotation period of KOREASAT 1 was analyzed with the light curves from the photometry results. The rotation periods of the low Earth orbit (LEO) satellite ASTRO-H after break-up were detected by OWL-Net on April 7, 2016. We analyzed the magnitude variation of each satellite by differential photometry and made comparisons with the star catalog. The illumination effect caused by the phase angle between the Sun and the target satellite was corrected with the system tool kit (STK) and two line element (TLE) technique. Finally, we determined the rotation period of two inactive satellites on LEO and geostationary Earth orbit (GEO) with light curves from the photometry. The main rotation periods were determined to be 5.2 sec for ASTRO-H and 74 sec for KOREASAT 1.

An Update on the Effectiveness of Postmission Disposal in LEO

NASA Technical Reports Server (NTRS)

Liou, J.-C; Krisko, Paula

2013-01-01

Orbital debris mitigation measures have been developed to reduce the growth of the future debris population by the international space community over the past two decades. A major component in the overall mitigation strategy is postmission disposal (PMD). A key PMD element for the low Earth orbit (LEO, the region below 2000 km altitude) satellites is the 25-year decay rule. It is intended to limit the long-term presence of massive intact objects - rocket bodies (R/Bs) and spacecraft (S/C) in the environment. The effectiveness of the 25-year rule was well demonstrated and documented during the development of the mitigation measures. The orbital debris population in LEO, unfortunately, has significantly increased since that time. The objectives of this paper are to provide an updated assessment based on the 2012 LEO environment and to highlight the importance of the global compliance of the 25-year decay rule.

Precise Orbit Determination for LEO Spacecraft Using GNSS Tracking Data from Multiple Antennas

NASA Technical Reports Server (NTRS)

Kuang, Da; Bertiger, William; Desai, Shailen; Haines, Bruce

2010-01-01

To support various applications, certain Earth-orbiting spacecrafts (e.g., SRTM, COSMIC) use multiple GNSS antennas to provide tracking data for precise orbit determination (POD). POD using GNSS tracking data from multiple antennas poses some special technical issues compared to the typical single-antenna approach. In this paper, we investigate some of these issues using both real and simulated data. Recommendations are provided for POD with multiple GNSS antennas and for antenna configuration design. The observability of satellite position with multiple antennas data is compared against single antenna case. The impact of differential clock (line biases) and line-of-sight (up, along-track, and cross-track) on kinematic and reduced-dynamic POD is evaluated. The accuracy of monitoring the stability of the spacecraft structure by simultaneously performing POD of the spacecraft and relative positioning of the multiple antennas is also investigated.

Autonomous Low Earth Orbit Satellite and Orbital Debris Tracking Using Mid Aperture COTS Optical Trackers

NASA Astrophysics Data System (ADS)

Ehrhorn, B.; Azari, D.

Low Earth Orbit (LEO) and Orbital Debris tracking have become considerably important with regard to Space Situational Awareness (SSA). This paper discusses the capabilities of autonomous LEO and Orbital Debris Tracking Systems using commercially available (mid aperture 20-24 inch) telescopes, tracking gimbals, and CCD imagers. RC Optical Systems has been developing autonomous satellite trackers that allow for unattended acquisition, imaging, and orbital determination of LEOs using low cost COTS equipment. The test setup from which we are gathering data consists of an RC Optical Systems Professional Series Elevation over Azimuth Gimbal with field de-rotation, RC Optical Systems 20 inch Ritchey-Chretien Telescope coupled to an e2v CCD42-40 CCD array, and 77mm f/4 tracking lens coupled to a KAF-0402ME CCD array. Central to success of LEO acquisition and open loop tracking is accurate modeling of Gimbal and telescope misalignments and flexures. Using pro-TPoint and a simple automated mapping routine we have modeled our primary telescope to achieve pointing and tracking accuracies within a population standard deviation of 1.3 arc-sec (which is 1.1 arc-sec RMS). Once modeled, a mobile system can easily and quickly be calibrated to the sky using a simple 6-10 star map to solve for axis tilt and collimation coefficients. Acquisition of LEO satellites is accomplished through the use of a wide field imager. Using a 77mm f/4 lens and 765 x 510 x 9mu CCD array yields a 1.28 x 0.85 degree field of view in our test setup. Accurate boresite within the acquisition array is maintained throughout the full range of motion through differential tpoint modeling of the main and acquisition imagers. Satellite identification is accomplished by detecting a stationary centroid as a point source and differentiating from the background of streaked stars in a single frame. We found 100% detection rate of LEO with radar cross sections (RCS) of > 0.5 meter*meter within the acquisition array, and

System using leo satellites for centimeter-level navigation

NASA Technical Reports Server (NTRS)

Rabinowitz, Matthew (Inventor); Parkinson, Bradford W. (Inventor); Cohen, Clark E. (Inventor); Lawrence, David G. (Inventor)

2002-01-01

Disclosed herein is a system for rapidly resolving position with centimeter-level accuracy for a mobile or stationary receiver [4]. This is achieved by estimating a set of parameters that are related to the integer cycle ambiguities which arise in tracking the carrier phase of satellite downlinks [5,6]. In the preferred embodiment, the technique involves a navigation receiver [4] simultaneously tracking transmissions [6] from Low Earth Orbit Satellites (LEOS) [2] together with transmissions [5] from GPS navigation satellites [1]. The rapid change in the line-of-sight vectors from the receiver [4] to the LEO signal sources [2], due to the orbital motion of the LEOS, enables the resolution with integrity of the integer cycle ambiguities of the GPS signals [5] as well as parameters related to the integer cycle ambiguity on the LEOS signals [6]. These parameters, once identified, enable real-time centimeter-level positioning of the receiver [4]. In order to achieve high-precision position estimates without the use of specialized electronics such as atomic clocks, the technique accounts for instabilities in the crystal oscillators driving the satellite transmitters, as well as those in the reference [3] and user [4] receivers. In addition, the algorithm accommodates as well as to LEOS that receive signals from ground-based transmitters, then re-transmit frequency-converted signals to the ground.

Low earth orbiting Nadir Etalon Sounding Spectrometer instrument concept for temperature, moisture and trace species, LeoNESS

NASA Technical Reports Server (NTRS)

Kumer, J. B.; Sterritt, L. W.; Roche, A. E.; Rosenberg, W. J.; Morrow, H. E.; Shenk, W. E.; Susskind, J.

1992-01-01

A concept for a low earth orbiting nadir etalon spectrometer sounder (LeoNESS) is described which can achieve retrieval of temperature, H2O, surface, boundary conditions, cloudiness, and trace species with an accuracy that meets or exceeds the AIRS specifications. Options employing 65-K and 30-K detectors are examined; the former may be implemented via passive radiative cooling. The concept, which is derived from the Cryogenic Limb Array Etalon Spectrometer, has the potential for improving the horizontal and vertical resolution.

Sizing of "Mother Ship and Catcher" Concepts for LEO Small Debris Capture

NASA Technical Reports Server (NTRS)

Bacon, John B.

2009-01-01

Most Low Earth Orbit (LEO) debris lies in a limited number of inclination "bands" associated with launch latitudes, or with specific useful orbit inclinations (such as polar orbits). Such narrow inclination bands generally have a uniform spread over all possible Right Ascensions of Ascending Node (RAANs), creating a different orbit plane for nearly every piece of debris. This complicates concept of rendezvous and capture for debris removal. However, a low-orbiting satellite will always phase in RAAN faster than debris objects in higher orbits at the same inclination, potentially solving the problem. Such a base can serve as a single space-based launch facility (a "mother ship") that can tend and then send tiny individual catcher devices for each debris object, as the facility drifts into the same RAAN as the higher object. This presentation will highlight characteristic system requirements of such an architecture, including structural and navigation requirements, power, mass and dV budgets for both the mother ship and the mass-produced common catcher devices that would clean out selected inclination bands. The altitude and inclination regime over which a band is to be cleared, the size distribution of the debris, and the inclusion of additional mission priorities all affect the sizing of the system. It is demonstrated that major LEO hazardous debris reductions can be realized in each band with a single LEO launch of a single mother ship, with simple attached catchers of total mass less than typical commercial LEO launch capability.

Orbital Debris Research in the United States

NASA Technical Reports Server (NTRS)

Stansbery, Gene

2009-01-01

The presentation includes information about growth of the satellite population, the U.S. Space Surveillance Network, tracking and catalog maintenance, Haystack and HAX radar observation, Goldstone radar, the Michigan Orbital Debris Survey Telescope (MODEST), spacecraft surface examinations and sample of space shuttle impacts. GEO/LEO observations from Kwajalein Atoll, NASA s Orbital Debris Engineering Model (ORDEM2008), a LEO-to-GEO Environment Debris Model (LEGEND), Debris Assessment Software (DAS) 2.0, the NASA/JSC BUMPER-II meteoroid/debris threat assessment code, satellite reentry risk assessment, optical size and shape determination, work on more complicated fragments, and spectral studies.

Geocenter Coordinates from a Combined Processing of LEO and Ground-based GPS Observations

NASA Astrophysics Data System (ADS)

Männel, Benjamin; Rothacher, Markus

2017-04-01

The GPS observations provided by the global IGS (International GNSS Service) tracking network play an important role for the realization of a unique terrestrial reference frame that is accurate enough to allow the monitoring of the Earth's system. Combining these ground-based data with GPS observations tracked by high-quality dual-frequency receivers on-board Low Earth Orbiters (LEO) might help to further improve the realization of the terrestrial reference frame and the estimation of the geocenter coordinates, GPS satellite orbits and Earth rotation parameters (ERP). To assess the scope of improvement, we processed a network of 50 globally distributed and stable IGS-stations together with four LEOs (GRACE-A, GRACE-B, OSTM/Jason-2 and GOCE) over a time interval of three years (2010-2012). To ensure fully consistent solutions the zero-difference phase observations of the ground stations and LEOs were processed in a common least-square adjustment, estimating GPS orbits, LEO orbits, station coordinates, ERPs, site-specific tropospheric delays, satellite and receiver clocks and ambiguities. We present the significant impact of the individual LEOs and a combination of all four LEOs on geocenter coordinates derived by using a translational approach (also called network shift approach). In addition, we present geocenter coordinates derived from the same set of GPS observations by using a unified approach. This approach combines the translational and the degree-one approach by estimating translations and surface deformations simultaneously. Based on comparisons against each other and against geocenter time series derived by other techniques the effect of the selected approach is assessed.

Real-Time Optical Surveillance of LEO/MEO with Small Telescopes

NASA Astrophysics Data System (ADS)

Zimmer, P.; McGraw, J.; Ackermann, M.

J.T. McGraw and Associates, LLC operates two proof-of-concept wide-field imaging systems to test novel techniques for uncued surveillance of LEO/MEO/GEO and, in collaboration with the University of New Mexico (UNM), uses a third small telescope for rapidly queued same-orbit follow-up observations. Using our GPU-accelerated detection scheme, the proof-of-concept systems operating at sites near and within Albuquerque, NM, have detected objects fainter than V=13 at greater than 6 sigma significance. This detection approximately corresponds to a 16 cm object with albedo of 0.12 at 1000 km altitude. Dozens of objects are measured during each operational twilight period, many of which have no corresponding catalog object. The two proof-of-concept systems, separated by ~30km, work together by taking simultaneous images of the same orbital volume to constrain the orbits of detected objects using parallax measurements. These detections are followed-up by imaging photometric observations taken at UNM to confirm and further constrain the initial orbit determination and independently assess the objects and verify the quality of the derived orbits. This work continues to demonstrate that scalable optical systems designed for real-time detection of fast moving objects, which can be then handed off to other instruments capable of tracking and characterizing them, can provide valuable real-time surveillance data at LEO and beyond, which substantively informs the SSA process.

Materials selection for long life in LEO: A critical evaluation of atomic oxygen testing with thermal atom systems

NASA Technical Reports Server (NTRS)

Koontz, S. L.; Kuminecz, J.; Leger, L.; Nordine, P.

1988-01-01

The use of thermal atom test methods as a materials selection and screening technique for low-Earth orbit (LEO) spacecraft is critically evaluated. The chemistry and physics of thermal atom environments are compared with the LEO environment. The relative reactivities of a number of materials determined to be in thermal atom environments are compared to those observed in LEO and in high quality LEO simulations. Reaction efficiencies measured in a new type of thermal atom apparatus are one-hundredth to one-thousandth those observed in LEO, and many materials showing nearly identical reactivities in LEO show relative reactivities differing by as much as a factor of 8 in thermal atom systems. A simple phenomenological kinetic model for the reaction of oxygen atoms with organic materials can be used to explain the differences in reactivity in different environments. Certain specific thermal test environments can be used as reliable materials screening tools. Using thermal atom methods to predict material lifetime in LEO requires direct calibration of the method against LEO data or high quality simulation data for each material.

Geocenter variations derived from a combined processing of LEO- and ground-based GPS observations

NASA Astrophysics Data System (ADS)

Männel, Benjamin; Rothacher, Markus

2017-08-01

GNSS observations provided by the global tracking network of the International GNSS Service (IGS, Dow et al. in J Geod 83(3):191-198, 2009) play an important role in the realization of a unique terrestrial reference frame that is accurate enough to allow a detailed monitoring of the Earth's system. Combining these ground-based data with GPS observations tracked by high-quality dual-frequency receivers on-board low earth orbiters (LEOs) is a promising way to further improve the realization of the terrestrial reference frame and the estimation of geocenter coordinates, GPS satellite orbits and Earth rotation parameters. To assess the scope of the improvement on the geocenter coordinates, we processed a network of 53 globally distributed and stable IGS stations together with four LEOs (GRACE-A, GRACE-B, OSTM/Jason-2 and GOCE) over a time interval of 3 years (2010-2012). To ensure fully consistent solutions, the zero-difference phase observations of the ground stations and LEOs were processed in a common least-squares adjustment, estimating all the relevant parameters such as GPS and LEO orbits, station coordinates, Earth rotation parameters and geocenter motion. We present the significant impact of the individual LEO and a combination of all four LEOs on the geocenter coordinates. The formal errors are reduced by around 20% due to the inclusion of one LEO into the ground-only solution, while in a solution with four LEOs LEO-specific characteristics are significantly reduced. We compare the derived geocenter coordinates w.r.t. LAGEOS results and external solutions based on GPS and SLR data. We found good agreement in the amplitudes of all components; however, the phases in x- and z-direction do not agree well.

Conceptual design of an orbital debris collector

NASA Technical Reports Server (NTRS)

Odonoghue, Peter (Editor); Brenton, Brian; Chambers, Ernest; Schwind, Thomas; Swanhart, Christopher; Williams, Thomas

1991-01-01

The current Lower Earth Orbit (LEO) environment has become overly crowded with space debris. An evaluation of types of debris is presented in order to determine which debris poses the greatest threat to operation in space, and would therefore provide a feasible target for removal. A target meeting these functional requirements was found in the Cosmos C-1B Rocket Body. These launchers are spent space transporters which constitute a very grave risk of collision and fragmentation in LEO. The motion and physical characteristics of these rocket bodies have determined the most feasible method of removal. The proposed Orbital Debris Collector (ODC) device is designed to attach to the Orbital Maneuvering Vehicle (OMV), which provides all propulsion, tracking, and power systems. The OMV/ODC combination, the Rocket Body Retrieval Vehicle (RBRV), will match orbits with the rocket body, use a spin table to match the rotational motion of the debris, capture it, despin it, and remove it from orbit by allowing it to fall into the Earth's atmosphere. A disposal analysis is presented to show how the debris will be deorbited into the Earth's atmosphere. The conceptual means of operation of a sample mission is described.

Plasma Flowfields Around Low Earth Orbit Objects: Aerodynamics to Underpin Orbit Predictions

NASA Astrophysics Data System (ADS)

Capon, Christopher; Boyce, Russell; Brown, Melrose

2016-07-01

Interactions between orbiting bodies and the charged space environment are complex. The large variation in passive body parameters e.g. size, geometry and materials, makes the plasma-body interaction in Low Earth Orbit (LEO) a region rich in fundamental physical phenomena. The aerodynamic interaction of LEO orbiting bodies with the neutral environment constitutes the largest non-conservative force on the body. However in general, study of the LEO plasma-body interaction has not been concerned with external flow physics, but rather with the effects on surface charging. The impact of ionospheric flow physics on the forces on space debris (and active objects) is not well understood. The work presented here investigates the contribution that plasma-body interactions have on the flow structure and hence on the total atmospheric force vector experienced by a polar orbiting LEO body. This work applies a hybrid Particle-in-Cell (PIC) - Direct Simulation Monte Carlo (DSMC) code, pdFoam, to self-consistently model the electrostatic flowfield about a cylinder with a uniform, fixed surface potential. Flow conditions are representative of the mean conditions experienced by the Earth Observing Satellite (EOS) based on the International Reference Ionosphere model (IRI-86). The electron distribution function is represented by a non-linear Boltzmann electron fluid and ion gas-surface interactions are assumed to be that of a neutralising, conducting, thermally accommodating solid wall with diffuse reflections. The variation in flowfield and aerodynamic properties with surface potential at a fixed flow condition is investigated, and insight into the relative contributions of charged and neutral species to the flow physics experienced by a LEO orbiting body is provided. This in turn is intended to help improve the fidelity of physics-based orbit predictions for space debris and other near-Earth space objects.

Thermospheric density estimation and responses to the March 2013 geomagnetic storm from GRACE GPS-determined precise orbits

NASA Astrophysics Data System (ADS)

Calabia, Andres; Jin, Shuanggen

2017-02-01

The thermospheric mass density variations and the thermosphere-ionosphere coupling during geomagnetic storms are not clear due to lack of observables and large uncertainty in the models. Although accelerometers on-board Low-Orbit-Earth (LEO) satellites can measure non-gravitational accelerations and derive thermospheric mass density variations with unprecedented details, their measurements are not always available (e.g., for the March 2013 geomagnetic storm). In order to cover accelerometer data gaps of Gravity Recovery and Climate Experiment (GRACE), we estimate thermospheric mass densities from numerical derivation of GRACE determined precise orbit ephemeris (POE) for the period 2011-2016. Our results show good correlation with accelerometer-based mass densities, and a better estimation than the NRLMSISE00 empirical model. Furthermore, we statistically analyze the differences to accelerometer-based densities, and study the March 2013 geomagnetic storm response. The thermospheric density enhancements at the polar regions on 17 March 2013 are clearly represented by POE-based measurements. Although our results show density variations better correlate with Dst and k-derived geomagnetic indices, the auroral electroject activity index AE as well as the merging electric field Em picture better agreement at high latitude for the March 2013 geomagnetic storm. On the other side, low-latitude variations are better represented with the Dst index. With the increasing resolution and accuracy of Precise Orbit Determination (POD) products and LEO satellites, the straightforward technique of determining non-gravitational accelerations and thermospheric mass densities through numerical differentiation of POE promises potentially good applications for the upper atmosphere research community.

Operation of commercial R3000 processors in the low earth orbit (LEO) space environment

NASA Astrophysics Data System (ADS)

Kaschmitter, J. L.; Shaeffer, D. L.; Colella, N. J.; McKnett, C. L.; Coakley, P. G.

1991-12-01

Spacecraft processors must operate with minimal degradation of performance in the LEO radiation environment, which includes the effects of total accumulated ionizing dose and single event phenomena (SEP) caused by protons and cosmic rays. Commercially available microprocessors can offer a number of advantages relative to radiation-hardened devices but are not normally designed to tolerate effects induced by the LEO environment. Extensive testing of the MIPS R3000 Reduced Instruction Set Computer (RISC) microprocessor family for operation in LEO environments is reported. The authors have characterized total dose and SEP effects for altitudes and inclinations of interest to systems operating in LEO, and they postulate techniques for detection and alleviation of SEP effects based on experimental results.

Low Earth orbital atomic oxygen and ultraviolet radiation effects on polymers

NASA Technical Reports Server (NTRS)

Dever, Joyce A.

1991-01-01

Because atomic oxygen and solar ultraviolet radiation present in the low earth orbital (LEO) environment can alter the chemistry of polymers resulting in degradation, their effects and mechanisms of degradation must be determined in order to determine the long term durability of polymeric surfaces to be exposed on missions such as Space Station Freedom. The effects of atomic oxygen on polymers which contain protective coatings must also be explored, since unique damage mechanisms can occur in areas where the protective coatings has failed. Mechanisms can be determined by utilizing results from previous LEO missions, by performing ground based LEO simulation tests and analysis, and by carrying out focussed space experiments. A survey is presented of the interactions and possible damage mechanisms for environmental atomic oxygen and UV radiation exposure of polymers commonly used in LEO.

Human Mars Mission: Launch Window from Earth Orbit. Pt. 1

NASA Technical Reports Server (NTRS)

Young, Archie

1999-01-01

The determination of orbital window characteristics is of major importance in the analysis of human interplanetary missions and systems. The orbital launch window characteristics are directly involved in the selection of mission trajectories, the development of orbit operational concepts, and the design of orbital launch systems. The orbital launch window problem arises because of the dynamic nature of the relative geometry between outgoing (departure) asymptote of the hyperbolic escape trajectory and the earth parking orbit. The orientation of the escape hyperbola asymptotic relative to earth is a function of time. The required hyperbola energy level also varies with time. In addition, the inertial orientation of the parking orbit is a function of time because of the perturbations caused by the Earth's oblateness. Thus, a coplanar injection onto the escape hyperbola can be made only at a point in time when the outgoing escape asymptote is contained by the plane of parking orbit. Even though this condition may be planned as a nominal situation, it will not generally represent the more probable injection geometry. The general case of an escape injection maneuver performed at a time other than the coplanar time will involve both a path angle and plane change and, therefore, a DELTA V penalty. Usually, because of the DELTA V penalty the actual departure injection window is smaller in duration than that determined by energy requirement alone. This report contains the formulation, characteristics, and test cases for five different launch window modes for Earth orbit. These modes are: (1) One impulsive maneuver from a Highly Elliptical Orbit (HEO) (2) Two impulsive maneuvers from a Highly Elliptical Orbit (HEO) (3) One impulsive maneuver from a Low Earth Orbit (LEO) (4) Two impulsive maneuvers from LEO (5) Three impulsive maneuvers from LEO.

Determining the material type of man-made orbiting objects using low-resolution reflectance spectroscopy

NASA Astrophysics Data System (ADS)

Jorgensen, Kira; Africano, John L.; Stansbery, Eugene G.; Kervin, Paul W.; Hamada, Kris M.; Sydney, Paul F.

2001-12-01

The purpose of this research is to improve the knowledge of the physical properties of orbital debris, specifically the material type. Combining the use of the fast-tracking United States Air Force Research Laboratory (AFRL) telescopes with a common astronomical technique, spectroscopy, and NASA resources was a natural step toward determining the material type of orbiting objects remotely. Currently operating at the AFRL Maui Optical Site (AMOS) is a 1.6-meter telescope designed to track fast moving objects like those found in lower Earth orbit (LEO). Using the spectral range of 0.4 - 0.9 microns (4000 - 9000 angstroms), researchers can separate materials into classification ranges. Within the above range, aluminum, paints, plastics, and other metals have different absorption features as well as slopes in their respective spectra. The spectrograph used on this telescope yields a three-angstrom resolution; large enough to see smaller features mentioned and thus determine the material type of the object. The results of the NASA AMOS Spectral Study (NASS) are presented herein.

Effects of LEO Environment on Tensile Properties of PEEK Films

NASA Astrophysics Data System (ADS)

Nakamura, Takashi; Nakamura, Hiroshi; Shimamura, Hiroyuki

2009-01-01

To clarify the effects of space environment on mechanical properties of polymer, exposure experiments were conducted utilizing the International Space Station Russian Service Module. Poly-ether-ether-ketone (PEEK) films under tensile stress were exposed to low Earth orbit (LEO) environment, and reference samples were irradiated with atomic oxygen (AO), electron beam (EB), and ultraviolet light (UV) in ground facilities. By comparing the results of flight and ground tests, the degradation behavior and the influential factors in LEO were investigated. The following results were obtained. (1) UV was found to be the harshest factor in LEO on tensile properties, since it decreased elongation to 15% of pristine sample after 46-months exposure. (2) AO in LEO eroded the specimen surface with a cone-like morphology and reduced the thickness; however, it had no significant effect on tensile properties. (3) EB irradiation in LEO had no measurable effects on the material properties.

Analysis of the Effect of UTI-UTC to High Precision Orbit

NASA Astrophysics Data System (ADS)

Shin, Dongseok; Kwak, Sunghee; Kim, Tag-Gon

1999-12-01

As the spatial resolution of remote sensing satellites becomes higher, very accurate determination of the position of a LEO (Low Earth Orbit) satellite is demanding more than ever. Non-symmetric Earth gravity is the major perturbation force to LEO satellites. Since the orbit propagation is performed in the celestial frame while Earth gravity is defined in the terrestrial frame, it is required to convert the coordinates of the satellite from one to the other accurately. Unless the coordinate conversion between the two frames is performed accurately the orbit propagation calculates incorrect Earth gravitational force at a specific time instant, and hence, causes errors in orbit prediction. The coordinate conversion between the two frames involves precession, nutation, Earth rotation and polar motion. Among these factors, unpredictability and uncertainty of Earth rotation, called UTI-UTC, is the largest error source. In this paper, the effect of UTI-UTC on the accuracy of the LEO propagation is introduced, tested and analzed. Considering the maximum unpredictability of UTI-UTC, 0.9 seconds, the meaningful order of non-spherical Earth harmonic functions is derived.

Instability of the Present LEO Satellite Populations

NASA Technical Reports Server (NTRS)

Liou, Jer-Chyi; Johnson, Nicholas L.

2006-01-01

Several studies conducted during 1991-2001 demonstrated, with some assumed launch rates, the future unintended growth potential of the Earth satellite population, resulting from random, accidental collisions among resident space objects. In some low Earth orbit (LEO) altitude regimes where the number density of satellites is above a critical spatial density, the production rate of new breakup debris due to collisions would exceed the loss of objects due to orbital decay. A new study has been conducted in the Orbital Debris Program Office at the NASA Lyndon B. Johnson Space Center, using higher fidelity models to evaluate the current debris environment. The study assumed no satellites were launched after December 2005. A total of 150 Monte Carlo runs were carried out and analyzed. Each Monte Carlo run simulated the current debris environment and projected it 200 years into the future. The results indicate that the LEO debris environment has reached a point such that even if no further space launches were conducted, the Earth satellite population would remain relatively constant for only the next 50 years or so. Beyond that, the debris population would begin to increase noticeably, due to the production of collisional debris. Detailed analysis shows that this growth is primarily driven by high collision activities around 900 to 1000 km altitude - the region which has a very high concentration of debris at present. In reality, the satellite population growth in LEO will undoubtedly be worse than this study indicates, since spacecraft and their orbital stages will continue to be launched into space. Postmission disposal of vehicles (e.g., limiting postmission orbital lifetimes to less than 25 years) will help, but will be insufficient to constrain the Earth satellite population. To preserve better the near-Earth environment for future space activities, it might be necessary to remove existing large and massive objects from regions where high collision activities are

Orbit to orbit transportation

NASA Technical Reports Server (NTRS)

Bergeron, R. P.

1980-01-01

Orbital transfer vehicle propulsion options for SPS include both chemical (COTV) and electrical (EOTV) options. The proposed EOTV construction method is similar to that of the SPS and, by the addition of a transmitting antenna, may serve as a demonstration or precursor satellite option. The results of the studies led to the selection of a single stage COTV for crew and priority cargo transfer. An EOTV concept is favored for cargo transfer because of the more favorable orbital burden factor over chemical systems. The gallium arsenide solar array is favored over the silicon array because of its self annealing characteristics of radiation damage encountered during multiple transitions through the Van Allen radiation belt. Transportation system operations are depicted. A heavy lift launch vehicle (HLLV) delivers cargo and propellants to LEO, which are transferred to a dedicated EOTV by means of an intraorbit transfer vehicle (IOTV) for subsequent transfer to GEO. The space shuttle is used for crew transfer from Earth to LEO. At the LEO base, the crew module is removed from the shuttle cargo bay and mated to a COTV for transfer to GEO. Upon arrival at GEO, the SPS construction cargo is transferred from the EOTV to the SPS construction base by IOTV. Crew consumables and resupply propellants are transported to GEO by the EOTV. Transportation requirements are dominated by the vast quantity of materials to be transported to LEO and GEO.

Water vapor retrieval by LEO and GEO SAR: techniques and performance evaluation.

NASA Astrophysics Data System (ADS)

Fermi, Alessandro; Silvio Marzano, Frank; Monti Guarnieri, Andrea; Pierdicca, Nazzareno; Realini, Eugenio; Venuti, Giovanna

2016-04-01

The millimetric sensitivity of SAR interferometry has been proved fruitful in estimating water-vapor maps, that can then be processed into higher level ZWD and PWV products. In the paper, we consider two different SAR surveys: Low Earth Orbiting (LEO) SAR, like ESA Sentinel-1, and Geosynchronous Earth Orbiting SAR. The two system are complementary, where LEO coverage is world-wide, while GEO is regional. On the other hand, LEO revisit is daily-to weekly, whereas GEO provides images in minutes to hours. Finally, LEO synthetic aperture is so short, less than a second, that the water-vapor is mostly frozen, whereas in the long GEO aperture the atmospheric phase screen would introduce a total decorrelation, if not compensated for. In the paper, we first review the Differential Interferometric techniques to get differential delay maps - to be then converted into water-vapor products, and then evaluate the quality in terms of geometric resolution, sensitivity, percentage of scene coverage, revisit, by referring to L and C band system, for both LEO and GEO. Finally, we discuss an empirical model for time-space variogram, and show a preliminary validation by campaign conducted with Ground Based Radar, as a proxy of GEO-SAR, capable of continuous scanning wide areas (up to 15 km) with metric resolution.

New Light-Time Curve of Eclipsing Binary AM Leo

NASA Astrophysics Data System (ADS)

Gorda, S. Yu.; Matveeva, E. A.

2017-12-01

We present 72 photoelectric and CCD times of minima of eclipsing binary AM Leo obtained mainly during at Kourovka Astronomical Observatory of the Ural Federal University in Russia. We obtained new values of period of 50.5 years and eccentricity of 0.28 of the orbit of the eclipsing pair around the mass center of the system AM Leo with the third body. These results have been received taking into account the times of minima taken from literature and obtained from to .

STK/Lifetime as a Replacement for Heritage Orbital Lifetime Software

NASA Technical Reports Server (NTRS)

Dove, Edwin

2004-01-01

The Flight Dynamics Analysis Branch (FDAB) of NASNGSFC is tasked with determining the orbital lifetime of several developmental and operational satellites, which include the Hubble Space Telescope. A DOS based program developed by the FDAB many years ago, called PC Lifetime, is used to determine a satellite s lifetime and could soon be in need of a replacement. STK s Lifetime Object Tool is a possible candidate. Due to the reduced support of the PC Lifetime program, and the growing incompatibility of older programs with new operating systems, a comparative analysis was done to determine if STWLifetime could meet the stringent requirements that were laid before it. The use of highly accurate numerical propagators such as STK s High Precision Orbit Propagator ( OP) and the Goddard Trajectory Determination System (GTDS) provided a basis on which to compare STWLifetime s results. Several test cases were run, but the main four test cases would determine whether or not STWLifetime could be PC- Lifetime s replacement. These four cases include a geotransfer orbit, two circular LEOS, and a Poiar LEO. Following rigorous testmg procedures, a conclusion will be determined. STK has proved to be a versatile program on many satellite missions and the FDAB has high hopes that it can pass FDAB s requirements for orbital lifetime prediction.

Sizing of "Mother Ship and Catcher" Missions for LEO Small Debris and for GEO Large Object Capture

NASA Technical Reports Server (NTRS)

Bacon, John B.

2009-01-01

Most LEO debris lies in a limited number of inclination "bands" associated with specific useful orbits. Objects in such narrow inclination bands have all possible Right Ascensions of Ascending Node (RAANs), creating a different orbit plane for nearly every piece of debris. However, a low-orbiting satellite will always phase in RAAN faster than debris objects in higher orbits at the same inclination, potentially solving the problem. Such a low-orbiting base can serve as a "mother ship" that can tend and then send small, disposable common individual catcher/deboost devices--one for each debris object--as the facility drifts into the same RAAN as each higher object. The dV necessary to catch highly-eccentric orbit debris in the center of the band alternatively allows the capture of less-eccentric debris in a wider inclination range around the center. It is demonstrated that most LEO hazardous debris can be removed from orbit in three years, using a single LEO launch of one mother ship--with its onboard magazine of freeflying low-tech catchers--into each of ten identified bands, with second or potentially third launches into only the three highest-inclination bands. The nearly 1000 objects near the geostationary orbit present special challenges in mass, maneuverability, and ultimate disposal options, leading to a dramatically different architecture and technology suite than the LEO solution. It is shown that the entire population of near-GEO derelict objects can be gathered and tethered together within a 3 year period for future scrap-yard operations using achievable technologies and only two earth launches.

Regional positioning using a low Earth orbit satellite constellation

NASA Astrophysics Data System (ADS)

Shtark, Tomer; Gurfil, Pini

2018-02-01

Global and regional satellite navigation systems are constellations orbiting the Earth and transmitting radio signals for determining position and velocity of users around the globe. The state-of-the-art navigation satellite systems are located in medium Earth orbits and geosynchronous Earth orbits and are characterized by high launching, building and maintenance costs. For applications that require only regional coverage, the continuous and global coverage that existing systems provide may be unnecessary. Thus, a nano-satellites-based regional navigation satellite system in Low Earth Orbit (LEO), with significantly reduced launching, building and maintenance costs, can be considered. Thus, this paper is aimed at developing a LEO constellation optimization and design method, using genetic algorithms and gradient-based optimization. The preliminary results of this study include 268 LEO constellations, aimed at regional navigation in an approximately 1000 km × 1000 km area centered at the geographic coordinates [30, 30] degrees. The constellations performance is examined using simulations, and the figures of merit include total coverage time, revisit time, and geometric dilution of precision (GDOP) percentiles. The GDOP is a quantity that determines the positioning solution accuracy and solely depends on the spatial geometry of the satellites. Whereas the optimization method takes into account only the Earth's second zonal harmonic coefficient, the simulations include the Earth's gravitational field with zonal and tesseral harmonics up to degree 10 and order 10, Solar radiation pressure, drag, and the lunisolar gravitational perturbation.

Telemetry Tracking & Control (TT&C) - First TDRSS, then Commercial GEO & Big LEO and Now Through LEO

NASA Technical Reports Server (NTRS)

Morgan, Dwayne R.; Streich, Ron G.; Bull, Barton; Grant, Chuck; Power, Edward I. (Technical Monitor)

2001-01-01

The advent of low earth orbit (LEO) commercial communication satellites provides an opportunity to dramatically reduce Telemetry, Tracking and Control (TT&C) costs of launch vehicles, Unpiloted Aerial Vehicles (UAVs), Research Balloons and spacecraft by reducing or eliminating ground infrastructure. Personnel from the Goddard Space Flight Center's Wallops Flight Facility (GSFC\\WFF) have successfully used commercial Geostationary Earth Orbit (GEO) and Big LEO communications satellites for Long Duration Balloon Flight TT&C. The Flight Modem is a GSFC\\WFF Advanced Range Technology initiative (ARTI) designed to streamline TT&C capability in the user community of these scientific data gathering platforms at low cost. Making use of existing LEO satellites and adapting and ruggedized commercially available components; two-way, over the horizon communications may be established with these vehicles at great savings due to reduced infrastructure. Initially planned as a means for permitting GPS data for tracking and recovery of sounding rocket and balloon payloads, expectations are that the bandwidth can soon be expanded to allow more comprehensive data transfer. The system architecture which integrates antennas, GPS receiver, commercial satellite packet data modem and a single board computer with custom software is described and technical challenges are discussed along with the plan for their resolution. A three-phase testing and development plan is outlined and the current results are reported. Results and status of ongoing flight tests on aircraft and sounding rockets are reported. Future applications on these platforms and the potential for satellite support are discussed along with an analysis of cost effectiveness of this method vs. other tracking and data transmission schemes.

Data analysis and interpretation related to space system/environment interactions at LEO altitude

NASA Technical Reports Server (NTRS)

Raitt, W. John; Schunk, Robert W.

1991-01-01

Several studies made on the interaction of active systems with the LEO space environment experienced from orbital or suborbital platforms are covered. The issue of high voltage space interaction is covered by theoretical modeling studies of the interaction of charged solar cell arrays with the ionospheric plasma. The theoretical studies were complemented by experimental measurements made in a vacuum chamber. The other active system studied was the emission of effluent from a space platform. In one study the emission of plasma into the LEO environment was studied by using initially a 2-D model, and then extending this model to 3-D to correctly take account of plasma motion parallel to the geomagnetic field. The other effluent studies related to the releases of neutral gas from an orbiting platform. One model which was extended and used determined the density, velocity, and energy of both an effluent gas and the ambient upper atmospheric gases over a large volume around the platform. This model was adapted to study both ambient and contaminant distributions around smaller objects in the orbital frame of reference with scale sizes of 1 m. The other effluent studies related to the interaction of the released neutral gas with the ambient ionospheric plasma. An electrostatic model was used to help understand anomalously high plasma densities measured at times in the vicinity of the space shuttle orbiter.

Lunar Reconnaissance Orbiter Orbit Determination Accuracy Analysis

NASA Technical Reports Server (NTRS)

Slojkowski, Steven E.

2014-01-01

Results from operational OD produced by the NASA Goddard Flight Dynamics Facility for the LRO nominal and extended mission are presented. During the LRO nominal mission, when LRO flew in a low circular orbit, orbit determination requirements were met nearly 100% of the time. When the extended mission began, LRO returned to a more elliptical frozen orbit where gravity and other modeling errors caused numerous violations of mission accuracy requirements. Prediction accuracy is particularly challenged during periods when LRO is in full-Sun. A series of improvements to LRO orbit determination are presented, including implementation of new lunar gravity models, improved spacecraft solar radiation pressure modeling using a dynamic multi-plate area model, a shorter orbit determination arc length, and a constrained plane method for estimation. The analysis presented in this paper shows that updated lunar gravity models improved accuracy in the frozen orbit, and a multiplate dynamic area model improves prediction accuracy during full-Sun orbit periods. Implementation of a 36-hour tracking data arc and plane constraints during edge-on orbit geometry also provide benefits. A comparison of the operational solutions to precision orbit determination solutions shows agreement on a 100- to 250-meter level in definitive accuracy.

Formation Flying in Earth, Libration, and Distant Retrograde Orbits

NASA Technical Reports Server (NTRS)

Folta, David C.

2004-01-01

This slide presentation examines the current and future state of formation flying, LEO formations, control strategies for flight in the vicinity of the libration points, and distant retrograde orbit formations. This discussion of LEO formations includes background on perturbation theory/accelerations and LEO formation flying. The discussion of strategies for formation flight in the vicinity of the libration points includes libration missions and natural and controlled libration orbit formations. A reference list is included.

A new concept for high-cycle-life LEO: Rechargeable MnO2-hydrogen

NASA Technical Reports Server (NTRS)

Appleby, A. J.; Dhar, H. P.; Kim, Y. J.; Murphy, O. J.

1989-01-01

The nickel-hydrogen secondary battery system, developed in the early 1970s, has become the system of choice for geostationary earth orbit (GEO) applications. However, for low earth orbit (LEO) satellites with long expected lifetimes the nickel positive limits performance. This requires derating of the cell to achieve very long cycle life. A new system, rechargeable MnO2-Hydrogen, which does not require derating, is described here. For LEO applications, it promises to have longer cycle life, high rate capability, a higher effective energy density, and much lower self-discharge behavior than those of the nickel-hydrogen system.

Space-based laser-powered orbital transfer vehicle (Project SLICK)

NASA Technical Reports Server (NTRS)

1988-01-01

A conceptual design study of a laser-powered orbital transfer vehicle (LOTV) is presented. The LOTV, nicknamed SLICK (Space Laser Interorbital Cargo Kite), will be utilized for the transfer of 16000 kg of cargo between Low Earth Orbit (LEO) and either Geosynchronous Earth Orbit (GEO) or Low Lunar Orbit (LLO). This design concentrates primarily on the LEO/GEO scenario, which will have typical LEO-to-GEO trip time of 6 days and two return versions. One version uses an all propulsive return while the other utilizes a ballute aerobrake for the return trip. Furthermore, three return cargo options of 16000 kg, 5000 kg (standard option), and 1600 kg are considered for this scenario. The LEO/LLO scenario uses only a standard, aerobraked version. The basic concept behind the LOTV is that the power for the propulsion system is supplied by a source separate from the LOTV itself. For the LEO/GEO scenario the LOTV utilizes a direct solar-pumped iodide laser and possibly two relay stations, all orbiting at an altitude of one Earth radius and zero inclination. An additional nuclear-powered laser is placed on the Moon for the LEO/LLO scenario. The propulsion system of the LOTV consists of a single engine fueled with liquid hydrogen. The laser beam is captured and directed by a four mirror optical system through a window in the thrust chamber of the engine. There, seven plasmas are created to convert the laser beam energy into thermal energy at an efficiency of at least 50 percent. For the LEO/LLO scenario the laser propulsion is supplemented by LH2/LOX chemical thrusters.

Establishing a Robotic, LEO-to-GEO Satellite Servicing Infrastructure as an Economic Foundation for Exploration

NASA Technical Reports Server (NTRS)

Horsham, Gary A. P.; Schmidt, George R.; Gilland, James H.

2010-01-01

The strategy for accomplishing civilian exploration goals and objectives is in the process of a fundamental shift towards a potential new approach called Flexible Path. This paper suggests that a government-industry or public-private partnership in the commercial development of low Earth orbit to geostationary orbit (LEO-to-GEO (LTG)) space, following or in parallel with the commercialization of Earth-to-LEO and International Space Station (ISS) operations, could serve as a necessary, logical step that can be incorporated into the flexible path approach. A LTG satellite-servicing infrastructure and architecture concept is discussed within this new strategic context. The concept consists of a space harbor that serves as a transport facility for a fleet of specialized, fully- or semi-autonomous robotic servicing spacecraft. The baseline, conceptual system architecture is composed of a space harbor equipped with specialized servicer spacecraft; a satellite command, communication, and control system; a parts station; a fuel station or depot; and a fuel/parts replenishment transport. The commercial servicer fleet would consist of several types of spacecraft, each designed with specialized robotic manipulation subsystems to provide services such as refueling, upgrade, repair, inspection, relocation, and removal. The space harbor is conceptualized as an ISS-type, octagonal truss structure equipped with radiation tolerant subsystems. This space harbor would be primarily capable of serving as an operational platform for various commercially owned and operated servicer spacecraft positioned and docked symmetrically on four of the eight sides. Several aspects of this concept are discussed, such as: system-level feasibility in terms of ISS-truss-type infrastructure and subsystems emplacement and maintenance between LEO and GEO; infrastructure components assembly in LEO, derived from ISS assembly experience, and transfer to various higher orbital locations; the evolving Earth-to-orbit

Study of LEO-SAT microwave link for broad-band mobile satellite communication system

NASA Technical Reports Server (NTRS)

Fujise, Masayuki; Chujo, Wataru; Chiba, Isamu; Furuhama, Yoji; Kawabata, Kazuaki; Konishi, Yoshihiko

1993-01-01

In the field of mobile satellite communications, a system based on low-earth-orbit satellites (LEO-SAT's) such as the Iridium system has been proposed. The LEO-SAT system is able to offer mobile telecommunication services in high-latitude areas. Rain degradation, fading and shadowing are also expected to be decreased when the system is operated at a high elevation angle. Furthermore, the propagation delay generated in the LEO-SAT system is less pronounced than that in the geostationary orbit satellite (GEO-SAT) system and, in voice services, the effect of the delay is almost negligible. We proposed a concept of a broad-band mobile satellite communication system with LEO-SAT's and Optical ISL. In that system, a fixed L-band (1.6/1.5 GHz) multibeam is used to offer narrow band service to the mobile terminals in the entire area covered by a LEO-SAT and steerable Ka-band (30/20 GHz) spot beams are used for the wide band service. In this paper, we present results of a study of LEO-SAT microwave link between a satellite and a mobile terminal for a broad-band mobile satellite communication system. First, the results of link budget calculations are presented and the antennas mounted on satellites are shown. For a future mobile antenna technology, we also show digital beamforming (DBF) techniques. DBF, together with modulation and/or demodulation, is becoming a key technique for mobile antennas with advanced functions such as antenna pattern calibration, correction, and radio interference suppression. In this paper, efficient DBF techniques for transmitting and receiving are presented. Furthermore, an adaptive array antenna system suitable for this LEO-SAT is presented.

Advanced propulsion for LEO and GEO platforms

NASA Technical Reports Server (NTRS)

Sovey, James S.; Pidgeon, David J.

1990-01-01

Mission requirements and mass savings applicable to specific low earth orbit and geostationary earth orbit platforms using three highly developed propulsion systems are described. Advanced hypergolic bipropellant thrusters and hydrazine arcjets can provide about 11 percent additional instrument payload to 14,000 kg LEO platforms. By using electric propulsion on a 8,000 kg class GEO platform, mass savings in excess of 15 percent of the beginning-of-life platform mass are obtained. Effects of large, advanced technology solar arrays and antennas on platform propulsion requirements are also discussed.

LEO-to-GEO low thrust chemical propulsion

NASA Technical Reports Server (NTRS)

Shoji, J. M.

1980-01-01

One approach being considered for transporting large space structures from low Earth orbit (LEO) to geosynchronous equatorial orbit (GEO) is the use of low thrust chemical propulsion systems. A variety of chemical rocket engine cycles evaluated for this application for oxygen/hydrogen and oxygen/hydrocarbon propellants (oxygen/methane and oxygen/RF-1) are discussed. These cycles include conventional propellant turbine drives, turboalternator/electric motor pump drive, and fuel cell/electric motor pump drive as well as pressure fed engines. Thrust chamber cooling analysis results are presented for regenerative/radiation and film/radiation cooling.

Human Exploration Missions Study Launch Window from Earth Orbit

NASA Technical Reports Server (NTRS)

Young, Archie

2001-01-01

The determination of orbital launch window characteristics is of major importance in the analysis of human interplanetary missions and systems. The orbital launch window characteristics are directly involved in the selection of mission trajectories, the development of orbit operational concepts, and the design of orbital launch systems. The orbital launch window problem arises because of the dynamic nature of the relative geometry between outgoing (departure) asymptote of the hyperbolic escape trajectory and the earth parking orbit. The orientation of the escape hyperbola asymptotic relative to earth is a function of time. The required hyperbola energy level also varies with time. In addition, the inertial orientation of the parking orbit is a function of time because of the perturbations caused by the Earth's oblateness. Thus, a coplanar injection onto the escape hyperbola can be made only at a point in time when the outgoing escape asymptote is contained by the plane of parking orbit. Even though this condition may be planned as a nominal situation, it will not generally represent the more probable injection geometry. The general case of an escape injection maneuver performed at a time other than the coplanar time will involve both a path angle and plane change and, therefore, a Delta(V) penalty. Usually, because of the Delta(V) penalty the actual departure injection window is smaller in duration than that determined by energy requirement alone. This report contains the formulation, characteristics, and test cases for five different launch window modes for Earth orbit. These modes are: (1) One impulsive maneuver from a Low Earth Orbit (LEO), (2) Two impulsive maneuvers from LEO, (3) Three impulsive maneuvers from LEO, (4) One impulsive maneuvers from a Highly Elliptical Orbit (HEO), (5) Two impulsive maneuvers from a Highly Elliptical Orbit (HEO) The formulation of these five different launch window modes provides a rapid means of generating realistic parametric

Arcing in Leo and Geo Simulated Environments: Comparative Analysis

NASA Technical Reports Server (NTRS)

Vayner, Boris V.; Ferguson, Dale C.; Galofaro, Joel TY.

2006-01-01

Comprehensive tests of two solar array samples in simulated Low Earth Orbit (LEO) and Geosynchronous Orbit (GEO) environments have demonstrated that the arc inception voltage was 2-3 times lower in the LEO plasma than in the GEO vacuum. Arc current pulse wave forms are also essentially different in these environments. Moreover, the wide variations of pulse forms do not allow introducing the definition of a "standard arc wave form" even in GEO conditions. Visual inspection of the samples after testing in a GEO environment revealed considerable damage on coverglass surfaces and interconnects. These harmful consequences can be explained by the discharge energy being one order of magnitude higher in vacuum than in background plasma. The tests also revealed a potential danger of powerful electrostatic discharges that could be initiated on the solar array surface of a satellite in GEO during the ignition of an arcjet thruster.

Dual RF Astrodynamic GPS Orbital Navigator Satellite

NASA Technical Reports Server (NTRS)

Kanipe, David B.; Provence, Robert Steve; Straube, Timothy M.; Reed, Helen; Bishop, Robert; Lightsey, Glenn

2009-01-01

Dual RF Astrodynamic GPS Orbital Navigator Satellite (DRAGONSat) will demonstrate autonomous rendezvous and docking (ARD) in low Earth orbit (LEO) and gather flight data with a global positioning system (GPS) receiver strictly designed for space applications. ARD is the capability of two independent spacecraft to rendezvous in orbit and dock without crew intervention. DRAGONSat consists of two picosatellites (one built by the University of Texas and one built by Texas A and M University) and the Space Shuttle Payload Launcher (SSPL); this project will ultimately demonstrate ARD in LEO.

Optical Tracking Data Validation and Orbit Estimation for Sparse Observations of Satellites by the OWL-Net.

PubMed

Choi, Jin; Jo, Jung Hyun; Yim, Hong-Suh; Choi, Eun-Jung; Cho, Sungki; Park, Jang-Hyun

2018-06-07

An Optical Wide-field patroL-Network (OWL-Net) has been developed for maintaining Korean low Earth orbit (LEO) satellites' orbital ephemeris. The OWL-Net consists of five optical tracking stations. Brightness signals of reflected sunlight of the targets were detected by a charged coupled device (CCD). A chopper system was adopted for fast astrometric data sampling, maximum 50 Hz, within a short observation time. The astrometric accuracy of the optical observation data was validated with precise orbital ephemeris such as Consolidated Prediction File (CPF) data and precise orbit determination result with onboard Global Positioning System (GPS) data from the target satellite. In the optical observation simulation of the OWL-Net for 2017, an average observation span for a single arc of 11 LEO observation targets was about 5 min, while an average optical observation separation time was 5 h. We estimated the position and velocity with an atmospheric drag coefficient of LEO observation targets using a sequential-batch orbit estimation technique after multi-arc batch orbit estimation. Post-fit residuals for the multi-arc batch orbit estimation and sequential-batch orbit estimation were analyzed for the optical measurements and reference orbit (CPF and GPS data). The post-fit residuals with reference show few tens-of-meters errors for in-track direction for multi-arc batch and sequential-batch orbit estimation results.

An atomic gravitational wave interferometric sensor in low earth orbit (AGIS-LEO)

NASA Astrophysics Data System (ADS)

Hogan, Jason M.; Johnson, David M. S.; Dickerson, Susannah; Kovachy, Tim; Sugarbaker, Alex; Chiow, Sheng-Wey; Graham, Peter W.; Kasevich, Mark A.; Saif, Babak; Rajendran, Surjeet; Bouyer, Philippe; Seery, Bernard D.; Feinberg, Lee; Keski-Kuha, Ritva

2011-07-01

We propose an atom interferometer gravitational wave detector in low Earth orbit (AGIS-LEO). Gravitational waves can be observed by comparing a pair of atom interferometers separated by a 30 km baseline. In the proposed configuration, one or three of these interferometer pairs are simultaneously operated through the use of two or three satellites in formation flight. The three satellite configuration allows for the increased suppression of multiple noise sources and for the detection of stochastic gravitational wave signals. The mission will offer a strain sensitivity of {<10^{-18}/sqrt{Hz}} in the 50mHz-10Hz frequency range, providing access to a rich scientific region with substantial discovery potential. This band is not currently addressed with the LIGO, VIRGO, or LISA instruments. We analyze systematic backgrounds that are relevant to the mission and discuss how they can be mitigated at the required levels. Some of these effects do not appear to have been considered previously in the context of atom interferometry, and we therefore expect that our analysis will be broadly relevant to atom interferometric precision measurements. Finally, we present a brief conceptual overview of shorter-baseline ({lesssim100 m}) atom interferometer configurations that could be deployed as proof-of-principle instruments on the International Space Station (AGIS-ISS) or an independent satellite.

Performance of Duplex Communication Between a LEO Satellite and Terrestrial Location Using a GEO Constellation

NASA Technical Reports Server (NTRS)

Robinson, Daryl C.; Konangi, Vijay K.; Wallett, Thomas M.

1998-01-01

A network comprised of a terrestrial site, a constellation of three GEO satellites and a LEO satellite is modeled and simulated. Continuous communication between the terrestrial site and the LEO satellite is facilitated by the GEO satellites. The LEO satellite has the orbital characteristics of the International Space Station. Communication in the network is based on TCP/IP over ATM, with the ABR service category providing the QoS, at OC-3 data rate. The OSPF protocol is used for routing. We simulate FTP file transfers, with the terrestrial site serving as the client and the LEO satellite being the server. The performance characteristics are presented.

Human Mars Mission: Launch Window from Earth Orbit. Pt. 1

NASA Technical Reports Server (NTRS)

Young, Archie

1999-01-01

The determination of orbital window characteristics is of major importance in the analysis of human interplanetary missions and systems. The orbital launch window characteristics are directly involved in the selection of mission trajectories, the development of orbit operational concepts, and the design of orbital launch systems. The orbital launch window problem arises because of the dynamic nature of the relative geometry between outgoing (departure) asymptote of the hyperbolic escape trajectory and the earth parking orbit. The orientation of the escape hyperbola asymptotic relative to the earth is a function of time. The required hyperbola energy level also varies with time. In addition, the inertial orientation of the parking orbit is a function of time because of the perturbations caused by the Earth's oblateness. Thus, a coplanar injection onto the escape hyperbola can be made only at a point in time when the outgoing escape asymptote is contained by the plane of parking orbit. Even though this condition may be planned as a nominal situation, it will not generally represent the more probable injection geometry. The general case of an escape injection maneuver performed at a time other than the coplanar time will involve both a path angle and plane change and, therefore, a delta V penalty. Usually, because of the delta V penalty the actual departure injection window is smaller in duration than that determined by energy requirement alone. This report contains the formulation, characteristics, and test cases for five different launch window modes for Earth orbit. These modes are: 1) One impulsive maneuver from a Highly Elliptical Orbit (HEO); 2) Two impulsive maneuvers from a Highly Elliptical Orbit (HEO); 3) One impulsive maneuver from a Low Earth Orbit (LEO); 4) Two impulsive maneuvers form LEO; and 5) Three impulsive maneuvers form LEO. The formulation of these five different launch window modes provides a rapid means of generating realistic parametric data

LEO Spacecraft Charging Guidelines

NASA Technical Reports Server (NTRS)

Hillard, G. B.; Ferguson, D. C.

2002-01-01

Over the past decade, Low Earth Orbiting (LEO) spacecraft have gradually required ever-increasing power levels. As a rule, this has been accomplished through the use of high voltage systems. Recent failures and anomalies on such spacecraft have been traced to various design practices and materials choices related to the high voltage solar arrays. NASA Glenn has studied these anomalies including plasma chamber testing on arrays similar to those that experienced difficulties on orbit. Many others in the community have been involved in a comprehensive effort to understand the problems and to develop practices to avoid them. The NASA Space Environments and Effects program, recognizing the timeliness of this effort, has commissioned and funded a design guidelines document intended to capture the current state of understanding. We present here an overview of this document, which is now nearing completion.

Study of mobile satellite network based on GEO/LEO satellite constellation

NASA Astrophysics Data System (ADS)

Hu, Xiulin; Zeng, Yujiang; Wang, Ying; Wang, Xianhui

2005-11-01

Mobile satellite network with Inter Satellite Links (ISLs), which consists of non-geostationary satellites, has the characteristic of network topology's variability. This is a great challenge to the design and management of mobile satellite network. This paper analyzes the characteristics of mobile satellite network, takes multimedia Quality of Service (QoS) as the chief object and presents a reference model based on Geostationary Earth Orbit (GEO)/ Low Earth Orbit (LEO) satellite constellation which adapts to the design and management of mobile satellite network. In the reference model, LEO satellites constitute service subnet with responsibility for the access, transmission and switch of the multimedia services for mobile users, while GEO satellites constitute management subnet taking on the centralized management to service subnet. Additionally ground control centre realizes the whole monitoring and control via management subnet. Comparing with terrestrial network, the above reference model physically separates management subnet from service subnet, which not only enhances the advantage of centralized management but also overcomes the shortcoming of low reliability in terrestrial network. Routing of mobile satellite network based on GEO/LEO satellite constellation is also discussed in this paper.

Ionospheric refraction effects on orbit determination using the orbit determination error analysis system

NASA Technical Reports Server (NTRS)

Yee, C. P.; Kelbel, D. A.; Lee, T.; Dunham, J. B.; Mistretta, G. D.

1990-01-01

The influence of ionospheric refraction on orbit determination was studied through the use of the Orbit Determination Error Analysis System (ODEAS). The results of a study of the orbital state estimate errors due to the ionospheric refraction corrections, particularly for measurements involving spacecraft-to-spacecraft tracking links, are presented. In current operational practice at the Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF), the ionospheric refraction effects on the tracking measurements are modeled in the Goddard Trajectory Determination System (GTDS) using the Bent ionospheric model. While GTDS has the capability of incorporating the ionospheric refraction effects for measurements involving ground-to-spacecraft tracking links, such as those generated by the Ground Spaceflight Tracking and Data Network (GSTDN), it does not have the capability to incorporate the refraction effects for spacecraft-to-spacecraft tracking links for measurements generated by the Tracking and Data Relay Satellite System (TDRSS). The lack of this particular capability in GTDS raised some concern about the achievable accuracy of the estimated orbit for certain classes of spacecraft missions that require high-precision orbits. Using an enhanced research version of GTDS, some efforts have already been made to assess the importance of the spacecraft-to-spacecraft ionospheric refraction corrections in an orbit determination process. While these studies were performed using simulated data or real tracking data in definitive orbit determination modes, the study results presented here were obtained by means of covariance analysis simulating the weighted least-squares method used in orbit determination.

Nickel hydrogen low Earth orbit life testing

NASA Technical Reports Server (NTRS)

Badcock, C. C.; Haag, R. L.

1986-01-01

A program to demonstrate the long term reliability of NiH2 cells in low Earth orbits (LEO) and support use in mid-altitude orbits (MAO) was initiated. Both 3.5 and 4.5 inch diameter nickel hydrogen cells are included in the test plan. Cells from all U.S. vendors are to be tested. The tests will be performed at -5 and 10 C at 40 and 60% DOD for LEO orbit and 10 C and 80% DOD for MAO orbit simulations. The goals of the testing are 20,000 cycles at 60% DOD and 30,000 cycles at 40% DOD. Cells are presently undergoing acceptance and characterization testing at Naval Weapons Systems Center, Crane.

Earth-Mars transfers through Moon Distant Retrograde Orbits

NASA Astrophysics Data System (ADS)

Conte, Davide; Di Carlo, Marilena; Ho, Koki; Spencer, David B.; Vasile, Massimiliano

2018-02-01

This paper focuses on the trajectory design which is relevant for missions that would exploit the use of asteroid mining in stable cis-lunar orbits to facilitate deep space missions, specifically human Mars exploration. Assuming that a refueling "gas station" is present at a given lunar Distant Retrograde Orbit (DRO), ways of departing from the Earth to Mars via that DRO are analyzed. Thus, the analysis and results presented in this paper add a new cis-lunar departure orbit for Earth-Mars missions. Porkchop plots depicting the required C3 at launch, v∞ at arrival, Time of Flight (TOF), and total Δ V for various DRO departure and Mars arrival dates are created and compared with results obtained for low Δ V Low Earth Orbit (LEO) to Mars trajectories. The results show that propellant-optimal trajectories from LEO to Mars through a DRO have higher overall mission Δ V due to the additional stop at the DRO. However, they have lower Initial Mass in LEO (IMLEO) and thus lower gear ratio as well as lower TOF than direct LEO to Mars transfers. This results in a lower overall spacecraft dry mass that needs to be launched into space from Earth's surface.

Galileo Jupiter approach orbit determination

NASA Technical Reports Server (NTRS)

Miller, J. K.; Nicholson, F. T.

1984-01-01

Orbit determination characteristics of the Jupiter approach phase of the Galileo mission are described. Predicted orbit determination performance is given for the various mission events that occur during Jupiter approach. These mission events include delivery of an atmospheric entry probe, acquisition of probe science data by the Galileo orbiter for relay to earth, delivery of an orbiter to a close encounter of the Galilean satellite Io, and insertion of the orbiter into orbit about Jupiter. The orbit determination strategy and resulting accuracies are discussed for the data types which include Doppler, range, optical imaging of Io, and a new Very Long Baseline Interferometry (VLBI) data type called Differential One-Way Range (DOR).

Precise Orbit Determination for ALOS

NASA Technical Reports Server (NTRS)

Nakamura, Ryo; Nakamura, Shinichi; Kudo, Nobuo; Katagiri, Seiji

2007-01-01

The Advanced Land Observing Satellite (ALOS) has been developed to contribute to the fields of mapping, precise regional land coverage observation, disaster monitoring, and resource surveying. Because the mounted sensors need high geometrical accuracy, precise orbit determination for ALOS is essential for satisfying the mission objectives. So ALOS mounts a GPS receiver and a Laser Reflector (LR) for Satellite Laser Ranging (SLR). This paper deals with the precise orbit determination experiments for ALOS using Global and High Accuracy Trajectory determination System (GUTS) and the evaluation of the orbit determination accuracy by SLR data. The results show that, even though the GPS receiver loses lock of GPS signals more frequently than expected, GPS-based orbit is consistent with SLR-based orbit. And considering the 1 sigma error, orbit determination accuracy of a few decimeters (peak-to-peak) was achieved.

Effect of KOH concentration on LEO cycle life of IPV nickel-hydrogen flight battery cells

NASA Technical Reports Server (NTRS)

Smithrick, John J.; Hall, Stephen W.

1990-01-01

A breakthrough in low earth orbit (LEO) cycle life of individual pressure vessel (IPV) nickel hydrogen battery cells was reported. The cycle life of boiler plate cells containing 26 percent potassium hydroxide (KOH) electrolyte was about 40,000 LEO cycles compared to 3500 cycles for cells containing 31 percent KOH. The effect of KOH concentration on cycle life was studied. The cycle regime was a stressful accelerated LEO, which consisted of a 27.5 min charge followed by a 17.5 min charge (2 x normal rate). The depth of discharge (DOD) was 80 percent. The cell temperature was maintained at 23 C. The next step is to validate these results using flight hardware and a real time LEO test. NASA Lewis has a contract with the Naval Weapons Support Center (NWSC), Crane, Indiana, to validate the boiler plate test results. Six 48 A-hr Hughes recirculation design IPV nickel-hydrogen flight battery cells are being evaluated. Three of the cells contain 26 percent KOH (test cells) and three contain 31 percent KOH (control cells). They are undergoing real time LEO cycle life testing. The cycle regime is a 90-min LEO orbit consisting of a 54-min charge followed by a 36-min discharge. The depth-of-discharge is 80 percent. The cell temperature is maintained at 10 C. The cells were cycled for over 8000 cycles in the continuing test. There were no failures for the cells containing 26 percent KOH. There was two failures, however, for the cells containing 31 percent KOH.

Effect of KOH concentration on LEO cycle life of IPV nickel-hydrogen flight battery cells

NASA Technical Reports Server (NTRS)

Smithrick, John J.; Hall, Stephen W.

1990-01-01

A breakthrough in the low-earth-orbit (LEO) cycle life of individual pressure vessel (IPV) nickel hydrogen battery cells is reported. The cycle life of boiler plate cells containing 26 percent potassium hydroxide (KOH) electrolyte was about 40,000 LEO cycles compared to 3500 cycles for cells containing 31 percent KOH. The effect of KOH concentration on cycle life was studied. The cycle regime was a stressful accelerated LEO, which consisted of a 27.5 min charge followed by a 17.5 min charge (2 x normal rate). The depth of discharge (DOD) was 80 percent. The cell temperature was maintained at 23 C. The next step is to validate these results using flight hardware and real time LEO test. NASA Lewis has a contract with the Naval Weapons Support Center (NWSC), Crane, Indiana to validate the boiler plate test results. Six 48 A-hr Hughes recirculation design IPV nickel-hydrogen flight battery cells are being evaluated. Three of the cells contain 26 percent KOH (test cells) and three contain 31 percent KOH (control cells). They are undergoing real time LEO cycle life testing. The cycle regime is a 90-min LEO orbit consisting of a 54-min charge followed by a 36-min discharge. The depth-of-discharge is 80 percent. The cell temperature is maintained at 10 C. The cells were cycled for over 8000 cycles in the continuing test. There were no failures for the cells containing 26 percent KOH. There were two failures, however, for the cells containing 31 percent KOH.

Optimization of Return Trajectories for Orbital Transfer Vehicle between Earth and Moon

NASA Technical Reports Server (NTRS)

Funase, Ryu; Tsuda, Yuichi; Kawaguchi, Jun'ichiro

2007-01-01

In this paper, optimum trajectories in Earth Transfer Orbit (ETO) for a lunar transportation system are proposed. This paper aims at improving the payload ratio of the reusable orbital transfer vehicle (OTV), which transports the payload from Low Earth Orbit (LEO) to Lunar Low Orbit (LLO) and returns to LEO. In ETO, we discuss ballistic flight using chemical propulsion, multi-impulse flight using electrical propulsion, and aero-assisted flight using aero-brake. The feasibility of the OTV is considered.

Comparison of precise orbit determination methods of zero-difference kinematic, dynamic and reduced-dynamic of GRACE-A satellite using SHORDE software

NASA Astrophysics Data System (ADS)

Li, Kai; Zhou, Xuhua; Guo, Nannan; Zhao, Gang; Xu, Kexin; Lei, Weiwei

2017-09-01

Zero-difference kinematic, dynamic and reduced-dynamic precise orbit determination (POD) are three methods to obtain the precise orbits of Low Earth Orbit satellites (LEOs) by using the on-board GPS observations. Comparing the differences between those methods have great significance to establish the mathematical model and is usefull for us to select a suitable method to determine the orbit of the satellite. Based on the zero-difference GPS carrier-phase measurements, Shanghai Astronomical Observatory (SHAO) has improved the early version of SHORDE and then developed it as an integrated software system, which can perform the POD of LEOs by using the above three methods. In order to introduce the function of the software, we take the Gravity Recovery And Climate Experiment (GRACE) on-board GPS observations in January 2008 as example, then we compute the corresponding orbits of GRACE by using the SHORDE software. In order to evaluate the accuracy, we compare the orbits with the precise orbits provided by Jet Propulsion Laboratory (JPL). The results show that: (1) If we use the dynamic POD method, and the force models are used to represent the non-conservative forces, the average accuracy of the GRACE orbit is 2.40cm, 3.91cm, 2.34cm and 5.17cm in radial (R), along-track (T), cross-track (N) and 3D directions respectively; If we use the accelerometer observation instead of non-conservative perturbation model, the average accuracy of the orbit is 1.82cm, 2.51cm, 3.48cm and 4.68cm in R, T, N and 3D directions respectively. The result shows that if we use accelerometer observation instead of the non-conservative perturbation model, the accuracy of orbit is better. (2) When we use the reduced-dynamic POD method to get the orbits, the average accuracy of the orbit is 0.80cm, 1.36cm, 2.38cm and 2.87cm in R, T, N and 3D directions respectively. This method is carried out by setting up the pseudo-stochastic pulses to absorb the errors of atmospheric drag and other

Orbital transfer and release of tethered payloads. Continuation of investigation of electrodynamic stabilization and control of long orbiting tethers Martinez-Sanchez, Manuel

NASA Technical Reports Server (NTRS)

Colombo, G.; Grossi, M. D.; Arnold, D.

1983-01-01

The effect of reeling operations on the orbital altitude of the tether system and the development of control laws to minimize tether rebound upon payload release were studied. The use of the tether for LEO/GEO payload orbital transfer was also investigated. It was concluded that (1) reeling operations can contribute a significant amount of energy to the orbit of the system and should be considered in orbit calculations and predictions, (2) deployment of payloads, even very large payloads, using tethers is a practical and fully stable operation, (3) tether augmented LEO/GEO transfer operations yield useful payload gains under the practical constraint of fixed size OTV's, and (4) orbit to orbit satellite retrieval is limited by useful revisit times to orbital inclinations of less than forty-five degrees.

The quick acquisition technique for laser communication between LEO and GEO

NASA Astrophysics Data System (ADS)

Zhang, Li-zhong; Zhang, Rui-qin; Li, Yong-hao; Meng, Li-xin; Li, Xiao-ming

2013-08-01

The sight-axis alignment can be accomplished by the quick acquisition operation between two laser communication terminals, which is the premise of establishing a free-space optical communication link. Especially for the laser communication links of LEO (Low Earth Orbit)-Ground and LEO-GEO (Geostationary Earth Orbit), since the earth would break the transmission of laser and break the communication as well, so the effective time for each communication is very shot (several minutes~ dozens of minutes), as a result the communication terminals have to capture each other to rebuild the laser communication link. In the paper, on the basis of the analysis of the traditional methods, it presents a new idea that using the long beacon light instead of the circular beacon light; thereby the original of two-dimensional raster spiral scanning is replaced by one-dimensional scanning. This method will reduce the setup time and decrease the failure probability of acquisition for the LEO-GEO laser communication link. Firstly, the analysis of the external constraint conditions in the acquisition phase has been presented in this paper. Furthermore, the acquisition algorithm models have been established. The optimization analysis for the parameters of the acquisition unit has been carried out, and the ground validation experiments of the acquisition strategy have also been performed. The experiments and analysis show that compared with traditional capturing methods, the method presented in this article can make the capturing time be shortened by about 40%, and the failure probability of capturing be reduced by about 30%. So, the method is significant for the LEO-GEO laser communication link.

HERRO: A Science-Oriented Strategy for Crewed Missions Beyond LEO

NASA Technical Reports Server (NTRS)

Schmidt, George R.

2011-01-01

This paper presents an exploration strategy for human missions beyond Low Earth Orbit (LEO) and the Moon that combines the best features of human and robotic spaceflight. This "Human Exploration using Real-time Robotic Operations" (HERRO) strategy refrains from placing humans on the surfaces of the Moon and Mars in the near-term. Rather, it focuses on sending piloted spacecraft and crews into orbit around exploration targets of interest, such as Mars, and conducting astronaut exploration of the surfaces using telerobots and remotely controlled systems. By eliminating the significant communications delay with Earth due to the speed of light limit, teleoperation provides scientists real-time control of rovers and other sophisticated instruments, in effect giving them a "virtual presence" on planetary surfaces, and thus expanding the scientific return at these destinations. It also eliminates development of the numerous man-rated landers, ascent vehicles and surface systems that are required to land humans on planetary surfaces. The propulsive requirements to travel from LEO to many destinations with shallow gravity-wells in the inner solar system are quite similar. Thus, a single spacecraft design could perform a variety of missions, including orbit-based surface exploration of the Moon, Mars and Venus, and rendezvous with Near Earth Asteroids (NEAs), as well as Phobos and Deimos. Although HERRO bypasses many of the initial steps that have been historically associated with human space exploration, it opens the door to many new destinations that are candidates for future resource utilization and settlement. HERRO is a first step that takes humans to exciting destinations beyond LEO, while expanding the ability to conduct science within the inner solar system.

Nickel metal hydride LEO cycle testing

NASA Technical Reports Server (NTRS)

Lowery, Eric

1995-01-01

The George C. Marshall Space Flight Center is working to characterize aerospace AB5 Nickel Metal Hydride (NiMH) cells. The cells are being evaluated in terms of storage, low earth orbit (LEO) cycling, and response to parametric testing (high rate charge and discharge, charge retention, pulse current ability, etc.). Cells manufactured by Eagle Picher are the subjects of the evaluation. There is speculation that NiMH cells may become direct replacements for current Nickel Cadmium cells in the near future.

Orbit determination singularities in the Doppler tracking of a planetary orbiter

NASA Technical Reports Server (NTRS)

Wood, L. J.

1985-01-01

On a number of occasions, spacecraft launched by the U.S. have been placed into orbit about the moon, Venus, or Mars. It is pointed out that, in particular, in planetary orbiter missions two-way coherent Doppler data have provided the principal data type for orbit determination applications. The present investigation is concerned with the problem of orbit determination on the basis of Doppler tracking data in the case of a spacecraft in orbit about a natural body other than the earth or the sun. Attention is given to Doppler shift associated with a planetary orbiter, orbit determination using a zeroth-order model for the Doppler shift, and orbit determination using a first-order model for the Doppler shift.

Controlling the Growth of Future LEO Debris Populations with Active Debris Removal

NASA Technical Reports Server (NTRS)

Liou, J.-C.; Johnson, N. L.; Hill, N. M.

2008-01-01

Active debris removal (ADR) was suggested as a potential means to remediate the low Earth orbit (LEO) debris environment as early as the 1980s. The reasons ADR has not become practical are due to its technical difficulties and the high cost associated with the approach. However, as the LEO debris populations continue to increase, ADR may be the only option to preserve the near-Earth environment for future generations. An initial study was completed in 2007 to demonstrate that a simple ADR target selection criterion could be developed to reduce the future debris population growth. The present paper summarizes a comprehensive study based on more realistic simulation scenarios, including fragments generated from the 2007 Fengyun-1C event, mitigation measures, and other target selection options. The simulations were based on the NASA long-term orbital debris projection model, LEGEND. A scenario, where at the end of mission lifetimes, spacecraft and upper stages were moved to 25-year decay orbits, was adopted as the baseline environment for comparison. Different annual removal rates and different ADR target selection criteria were tested, and the resulting 200-year future environment projections were compared with the baseline scenario. Results of this parametric study indicate that (1) an effective removal strategy can be developed based on the mass and collision probability of each object as the selection criterion, and (2) the LEO environment can be stabilized in the next 200 years with an ADR removal rate of five objects per year.

Re-Entry Point Targeting for LEO Spacecraft using Aerodynamic Drag

NASA Technical Reports Server (NTRS)

Omar, Sanny; Bevilacqua, Riccardo; Fineberg, Laurence; Treptow, Justin; Johnson, Yusef; Clark, Scott

2016-01-01

Most Low Earth Orbit (LEO) spacecraft do not have thrusters and re-enter atmosphere in random locations at uncertain times. Objects pose a risk to persons, property, or other satellites. Has become a larger concern with the recent increase in small satellites. Working on a NASA funded project to design a retractable drag device to expedite de-orbit and target a re-entry location through modulation of the drag area. Will be discussing the re-entry point targeting algorithm here.

A Parametric Study on Using Active Debris Removal for LEO Environment Remediation

NASA Technical Reports Server (NTRS)

2010-01-01

Recent analyses on the instability of the orbital debris population in the low Earth orbit (LEO) region and the collision between Iridium 33 and Cosmos 2251 have reignited the interest in using active debris removal (ADR) to remediate the environment. There are; however, monumental technical, resource, operational, legal, and political challenges in making economically viable ADR a reality. Before a consensus on the need for ADR can be reached, a careful analysis of its effectiveness must be conducted. The goal is to demonstrate the need and feasibility of using ADR to better preserve the future environment and to guide its implementation to maximize the benefit-to-cost ratio. This paper describes a new sensitivity study on using ADR to stabilize the future LEO debris environment. The NASA long-term orbital debris evolutionary model, LEGEND, is used to quantify the effects of several key parameters, including target selection criteria/constraints and the starting epoch of ADR implementation. Additional analyses on potential ADR targets among the currently existing satellites and the benefits of collision avoidance maneuvers are also included.

Comparison of currents predicted by NASCAP/LEO model simulations with elementary Langmuir-type bare probe models for an insulated cable containing a single pinhole

NASA Technical Reports Server (NTRS)

Galofaro, Joel T.

1990-01-01

The behavior of a defect in the insulation of a short biased section of cable in a Low Earth Orbit (LEO) space environment was examined. Such studies are of the utmost importance for large space power systems where great quantities of cabling will be deployed. An insulated probe containing a pinhole was placed into a hypothetical high speed LEO plasma. The NASA Charging Analyzer Program (NASCAP/LEO) was used to explore sheath growth about the probe as a function of applied voltage and to predict I-V behavior. A set of independent current calculations using Langmuir's formulations for concentric spheres and coaxial cylinders were also performed. The case of concentric spheres was here extended to include the case of concentric hemispheres. Several simple Langmuir-type models were then constructed to bracket the current collected by the cable. The space-charge sheath radius and impact parameters were used to determine the proper current regime. I-V curves were plotted for the models and comparisons were made with NASCAP/LEO results. Finally, NASCAP/LEO potential contours and surface cell potential plots were examined to explain interesting features in the NASCAP/LEO I-V curve.

An Assessment of the Current LEO Debris Environment and the Need for Active Debris Removal

NASA Technical Reports Server (NTRS)

Liou, Jer-Chyi

2010-01-01

The anti-satellite test on the Fengun-1 C weather satellite in early 2007 and the collision between Iridium 33 and Cosmos 2251 in 2009 dramatically altered the landscape of the human-made orbital debris environment in the low Earth orbit (LEO). The two events generated approximately 5500 fragments large enough to be tracked by the U.S. Space Surveillance Network. Those fragments account for more than 60% increase to the debris population in LEO. However, even before the ASAT test, model analyses already indicated that the debris population (for those larger than 10 cm) in LEO had reached a point where the population would continue to increase, due to collisions among existing objects, even without any future launches. The conclusion implies that as satellites continue to be launched and unexpected breakup events continue to occur, commonly-adopted mitigation measures will not be able to stop the collision-driven population growth. To remediate the debris environment in LEO, active debris removal must be considered. This presentation will provide an updated assessment of the debris environment after the Iridium 33/Cosmos 2251 collision, an analysis of several future environment projections based on different scenarios, and a projection of collision activities in LEO in the near future. The need to use active debris removal to stabilize future debris environment will be demonstrated and the effectiveness of various active debris removal strategies will be quantified.

Low Earth Orbit satellite traffic simulator

NASA Technical Reports Server (NTRS)

Hoelzel, John

1995-01-01

This paper describes a significant tool for Low Earth Orbit (LEO) capacity analysis, needed to support marketing, economic, and design analysis, known as a Satellite Traffic Simulator (STS). LEO satellites typically use multiple beams to help achieve the desired communication capacity, but the traffic demand in these beams in usually not uniform. Simulations of dynamic, average, and peak expected demand per beam is a very critical part of the marketing, economic, and design analysis necessary to field a viable LEO system. An STS is described in this paper which can simulate voice, data and FAX traffic carried by LEO satellite beams and Earth Station Gateways. It is applicable world-wide for any LEO satellite constellations operating over any regions. For aeronautical applications to LEO satellites. the anticipates aeronautical traffic (Erlangs for each hour of the day to be simulated) is prepared for geographically defined 'area targets' (each major operational region for the respective aircraft), and used as input to the STS. The STS was designed by Constellations Communications Inc. (CCI) and E-Systems for usage in Brazil in accordance with an ESCA/INPE Statement Of Work, and developed by Analytical Graphics Inc. (AGI) to execute on top of its Satellite Tool Kit (STK) commercial software. The STS simulates constellations of LEO satellite orbits, with input of traffic intensity (Erlangs) for each hour of the day generated from area targets (such as Brazilian States). accumulated in custom LEO satellite beams, and then accumulated in Earth Station Gateways. The STS is a very general simulator which can accommodate: many forms of orbital element and Walker Constellation input; simple beams or any user defined custom beams; and any location of Gateways. The paper describes some of these features, including Manual Mode dynamic graphical display of communication links, to illustrate which Gateway links are accessible and which links are not, at each 'step' of the

Demonstration of Uncued Optical Surveillance of LEO

NASA Astrophysics Data System (ADS)

Zimmer, P.; Ackermann, M.; McGraw, J.

2014-09-01

J.T. McGraw and Associates, LLC, in collaboration with the University of New Mexico (UNM), has built and is operating two proof-of-concept wide-field imaging systems to test novel techniques for uncued surveillance of LEO. The imaging systems are built from off-the-shelf optics and detectors resulting in a 350mm aperture and a 6 square degree field of view. For streak detection, field of view is of critical importance because the maximum exposure time on the object is limited by its crossing time and measurements of apparent angular motion are better constrained with longer streaks. The current match of the detector to the optical system is optimized for detection of objects at altitudes above 450km, which for a circular orbit, corresponds to apparent motions of approximately 1 deg./sec. Using our GPU-accelerated detection scheme, the proof-of-concept systems have detected objects fainter than V=12.3, which approximately corresponds to a 24 cm object at 1000 km altitude at better than 6 sigma significance, from sites near and within Albuquerque, NM. This work demonstrates scalable optical systems designed for near real time detection of fast moving objects, which can be then handed off to other instruments capable of tracking and characterizing them. The two proof-of-concept systems, separated by ~30km, work together by taking simultaneous images of the same orbital volume to constrain the orbits of detected objects using parallax measurements. These detections are followed-up by photometric observations taken at UNM to independently assess the objects and the quality of the derived orbits. We believe this demonstrates the potential of small telescope arrays for detecting and cataloguing heretofore unknown LEO objects.

The Space Motion of Leo I: The Mass of the Milky Way's Dark Matter Halo

NASA Astrophysics Data System (ADS)

Boylan-Kolchin, Michael; Bullock, James S.; Sohn, Sangmo Tony; Besla, Gurtina; van der Marel, Roeland P.

2013-05-01

We combine our Hubble Space Telescope measurement of the proper motion of the Leo I dwarf spheroidal galaxy (presented in a companion paper) with the highest resolution numerical simulations of Galaxy-size dark matter halos in existence to constrain the mass of the Milky Way's dark matter halo (M vir, MW). Despite Leo I's large Galactocentric space velocity (200 km s-1) and distance (261 kpc), we show that it is extremely unlikely to be unbound if Galactic satellites are associated with dark matter substructure, as 99.9% of subhalos in the simulations are bound to their host. The observed position and velocity of Leo I strongly disfavor a low-mass Milky Way: if we assume that Leo I is the least bound of the Milky Way's classical satellites, then we find that M vir, MW > 1012 M ⊙ at 95% confidence for a variety of Bayesian priors on M vir, MW. In lower mass halos, it is vanishingly rare to find subhalos at 261 kpc moving as fast as Leo I. Should an additional classical satellite be found to be less bound than Leo I, this lower limit on M vir, MW would increase by 30%. Imposing a mass-weighted ΛCDM prior, we find a median Milky Way virial mass of M vir, MW = 1.6 × 1012 M ⊙, with a 90% confidence interval of [1.0-2.4] × 1012 M ⊙. We also confirm a strong correlation between subhalo infall time and orbital energy in the simulations and show that proper motions can aid significantly in interpreting the infall times and orbital histories of satellites.

Can Telescopes Help Leo Satellites Dodge Most Lethal Impacts?

NASA Astrophysics Data System (ADS)

GUDIEL, ANDREA; Carroll, Joseph; Rowe, David

2018-01-01

Authors: Joseph Carroll and David RoweABSTRACT LEO objects are tracked by radar because it works day and night, in all weather. This fits military interest in potentially hostile objects. There is less interest in objects too small to be credible active threats. But accidental hypervelocity impact by even 5-10 mm objects can disable most LEO satellites. Such “cm-class” objects greatly outnumber objects of military interest, and will cause most accidental impact losses.Under good viewing conditions, a sunlit 5mm sphere with 0.15 albedo at 800 km altitude is a 19th magnitude object. A ground-based 0.5m telescope tracking it against a 20 mag/arcsec2 sky can see it in seconds, and provide <1 arc-second alt/az fixes against the stars. A key question is how much of the “usually lethal” LEO debris population can be tracked frequently, accurately, and affordably enough to be avoided. The value of a conjunction warning service should scale with the number of lethal objects in its catalog. This should motivate a commercial service to find and catalog most lethal objects. There may already be >1 million such objects in LEO, nearly all debris fragments, mostly cm-class and at 600-1200 km altitude.Maintaining a ~million-item catalog requires a world-wide network of several dozen telescope sites with several telescopes at each site. Each telescope needs a mount capable of ~1,000,000 fast slews/year without wearing out.The paper discusses recent advances that make such a service far more feasible:1. Automated tasking and remote control of distributed telescope networks,2. Direct-drive mounts that can make millions of fast slews without wearing out,3. Telescope optics with low focal curvature that are in focus across large imagers,4. CMOS imagers with 95% peak QE and 1.5e- noise at 2E8 pix/sec readout rates,5. Methods for uncued detection of most lethal LEO debris (eg., >5 mm at 800 km),6. Initial orbit determination using 3 alt-az fixes made during the discovery pass,7

Optimal aeroassisted orbital transfer with plane change using collocation and nonlinear programming

NASA Technical Reports Server (NTRS)

Shi, Yun. Y.; Nelson, R. L.; Young, D. H.

1990-01-01

The fuel optimal control problem arising in the non-planar orbital transfer employing aeroassisted technology is addressed. The mission involves the transfer from high energy orbit (HEO) to low energy orbit (LEO) with orbital plane change. The basic strategy here is to employ a combination of propulsive maneuvers in space and aerodynamic maneuvers in the atmosphere. The basic sequence of events for the aeroassisted HEO to LEO transfer consists of three phases. In the first phase, the orbital transfer begins with a deorbit impulse at HEO which injects the vehicle into an elliptic transfer orbit with perigee inside the atmosphere. In the second phase, the vehicle is optimally controlled by lift and bank angle modulations to perform the desired orbital plane change and to satisfy heating constraints. Because of the energy loss during the turn, an impulse is required to initiate the third phase to boost the vehicle back to the desired LEO orbital altitude. The third impulse is then used to circularize the orbit at LEO. The problem is solved by a direct optimization technique which uses piecewise polynomial representation for the state and control variables and collocation to satisfy the differential equations. This technique converts the optimal control problem into a nonlinear programming problem which is solved numerically. Solutions were obtained for cases with and without heat constraints and for cases of different orbital inclination changes. The method appears to be more powerful and robust than other optimization methods. In addition, the method can handle complex dynamical constraints.

Proceedings of the NASA Workshop on Atomic Oxygen Effects. [low earth orbital environment

NASA Technical Reports Server (NTRS)

Brinza, David E. (Editor)

1987-01-01

A workshop was held to address the scientific issues concerning the effects of atomic oxygen on materials in the low Earth orbital (LEO) environment. The program included 18 invited speakers plus contributed posters covering topics such as LEO spaceflight experiments, interaction mechanisms, and atomic oxygen source development. Discussion sessions were also held to organize a test program to evaluate atomic oxygen exposure facilities. The key issues raised in the workshop were: (1) the need to develop a reliable predictive model of the effects of long-term exposure of materials to the LEO environment; (2) the ability of ground-based exposure facilities to provide useful data for development of durable materials; and (3) accurate determination of the composition of the LEO environment. These proceedings include the invited papers, the abstracts for the contributed posters, and an account of the test program discussion sessions.

Effect of KOH concentration on LEO cycle life of IPV nickel-hydrogen flight cells-update 2

NASA Technical Reports Server (NTRS)

Smithrick, John J.; Hall, Stephen W.

1991-01-01

An update of validation test results confirming the breakthrough in low earth orbit (LEO) cycle life of nickel-hydrogen cells containing 26 percent KOH electrolyte is presented. A breakthrough in the LEO cycle life of individual pressure vessel (IPV nickel-hydrogen cells has been previously reported. The cycle life of boiler plate cells containing 26 percent potassium hydroxide (KOH) electrolyte was about 40 000 LEO cycles compared to 3500 cycles for cells containing 31 percent KOH. This test was conducted at Hughes Aircraft Company under a NASA Lewis contract. The purpose was to investigate the effect of KOH concentration on cycle life. The cycle regime was a stressful accelerated LEO, which consisted of a 27.5 min charge followed by a 17.5 min discharge (2x normal rate). The depth of discharge (DOD) was 80 percent. The cell temperature was maintained at 23 C. The boiler plate test results are in the process of being validated using flight hardware and real time LEO test at the Naval Weapons Support Center (NWSC), Crane, Indiana under a NASA Lewis Contract. Six 48 Ah Hughes recirculation design IPV nickel-hydrogen flight battery cells are being evaluated. Three of the cells contain 26 percent KOH (test cells), and three contain 31 percent KOH (control cells). They are undergoing real time LEO cycle life testing. The cycle regime is a 90-min LEO orbit consisting of a 54-min charge followed by a 36-min discharge. The depth-of-discharge is 80 percent. The cell temperature is maintained at 10 C. The three 31 percent KOH cells failed (cycles 3729, 4165, and 11355). One of the 26 percent KOH cells failed at cycle 15314. The other two 26 percent KOH cells were cycled for over 16600 cycles during the continuing test.

Effects of Orbital Lifetime Reduction on the Long-Term Earth Satellite Population as Modeled by EVOLVE 4.0

NASA Technical Reports Server (NTRS)

Krisko, Paula H.; Opiela, John N.; Liou, Jer-Chyi; Anz-Meador, Phillip D.; Theall, Jeffrey R.

1999-01-01

The latest update of the NASA orbital debris environment model, EVOLVE 4.0, has been used to study the effect of various proposed debris mitigation measures, including the NASA 25-year guideline. EVOLVE 4.0, which includes updates of the NASA breakup, solar activity, and the orbit propagator models, a GEO analysis option, and non-fragmentation debris source models, allows for the statistical modeling and predicted growth of the particle population >1 mm in characteristic length in LEO and GEO orbits. The initial implementation of this &odel has been to study the sensitivity of the overall LEO debris environment to mitigation measures designed to limit the lifetime of intact objects in LEO orbits. The mitigation measures test matrix for this study included several commonly accepted testing schemes, i.e., the variance of the maximum LEO lifetime from 10 to 50 years, the date of the initial implementation of this policy, the shut off of all explosions at some specified date, and the inclusion of disposal orbits. All are timely studies in that all scenarios have been suggested by researchers and satellite operators as options for the removal of debris from LEO orbits.

Geocenter Motion Derived from GNSS and SLR Tracking Data of LEO

NASA Astrophysics Data System (ADS)

Li, Y. S.; Ning, F. S.; Tseng, K. H.; Tseng, T. P.; Wu, J. M.; Chen, K. L.

2017-12-01

Space geodesy techniques can provide the monitoring data of global variations with high precision and large coverage through the satellites. Geocenter motion (GM) describes the difference of CF (Center of Figure) respect to CM (Center of Mass of the Earth System) due to the re-distribution and deformation of the earth system. Because satellite tracking data between ground stations and satellites orbit around the CM, geocenter motion is related to the realization of the ITRF (International Terrestrial Reference Frame) origin. In this study, GPS (Global Positioning System) observation data of IGS (International GNSS Service) and SLR (Satellite Laser Ranging) tracking data are applied to estimate the coordinates of observing sites on Earth's surface. The GPS observing sites are distributed deliberately and globally by 15° ×15° grids. Meanwhile, two different global ocean tide models are applied here. The model used in ITRF comparison and combination is parameter transformation, which is a mathematical formula allowing to transform the different frames between ITRF and CM system. Following the parameter transformation, the results of geocenter motion can be determined. The FORMOSAT-7/COSMIC-2 (F7C2) mission is a constellation of LEO (Low-Earth-Orbit) satellites, which will be launched in 2018. Besides the observing system for Meteorology, Ionosphere, and Climate, the F7C2 will be equipped with LRR (Laser Ranging Retroreflector). This work is a pilot survey to study the application of LEO SLR data in Taiwan.

Fuzzy attitude control for a nanosatellite in leo orbit

NASA Astrophysics Data System (ADS)

Calvo, Daniel; Laverón-Simavilla, Ana; Lapuerta, Victoria; Aviles, Taisir

Fuzzy logic controllers are flexible and simple, suitable for small satellites Attitude Determination and Control Subsystems (ADCS). In this work, a tailored fuzzy controller is designed for a nanosatellite and is compared with a traditional Proportional Integrative Derivative (PID) controller. Both control methodologies are compared within the same specific mission. The orbit height varies along the mission from injection at around 380 km down to a 200 km height orbit, and the mission requires pointing accuracy over the whole time. Due to both the requirements imposed by such a low orbit, and the limitations in the power available for the attitude control, a robust and efficient ADCS is required. For these reasons a fuzzy logic controller is implemented as the brain of the ADCS and its performance and efficiency are compared to a traditional PID. The fuzzy controller is designed in three separated controllers, each one acting on one of the Euler angles of the satellite in an orbital frame. The fuzzy memberships are constructed taking into account the mission requirements, the physical properties of the satellite and the expected performances. Both methodologies, fuzzy and PID, are fine-tuned using an automated procedure to grant maximum efficiency with fixed performances. Finally both methods are probed in different environments to test their characteristics. The simulations show that the fuzzy controller is much more efficient (up to 65% less power required) in single maneuvers, achieving similar, or even better, precision than the PID. The accuracy and efficiency improvement of the fuzzy controller increase with orbit height because the environmental disturbances decrease, approaching the ideal scenario. A brief mission description is depicted as well as the design process of both ADCS controllers. Finally the validation process and the results obtained during the simulations are described. Those results show that the fuzzy logic methodology is valid for small

LEDsats: LEO CubeSats with LEDs for Optical Tracking

NASA Astrophysics Data System (ADS)

Seitzer, P.; Cutler, J.; Piergentili, F.; Santoni, F.; Arena, L.; Cardona, T.; Cowardin, H.; Lee, C.; Sharma, S.

2016-09-01

We describe a project to launch 1U CubeSats equipped with Light Emitting Diodes (LEDs) into Low Earth Orbit (LEO) for optical tracking with ground-based telescopes. Active illumination on the satellites increases tremendously the number of passes where the LEO satellite is visible when the ground-based telescope is in darkness. The restriction that the satellite is in direct Sun is removed, and so tracking can take place all night long rather than just in twilight. The inspiration for this project came from the Japanese CubeSat FITSAT-1 that carried red and green high-powered LED arrays, and was clearly visible from the ground with small telescopes. There are two goals: 1) increase the accuracy and precision of LEO orbits by increasing the number and length of passes that satellite is visible, and 2) minimize the confusion between objects in the case of multiple CubeSats being launched at the same time. Technical issues to be discussed include the power level required for detection by small (20 - 40 cm) ground based telescopes, the optimum flash pattern for astrometry against star fields, and the timing of the flash pattern to millisecond or better accuracy and precision. We propose to deploy two such LEDsats simultaneously from the International Space Station: the first to be built at the University of Michigan, and the second to be built at Sapienza University Rome. One experiment is to see how we can distinguish these two CubeSats shortly after deployment solely from optical tracking, and so the CubeSats will have different flash patterns.

Low-Earth-Orbit and Geosynchronous-Earth-Orbit Testing of 80 Ah Batteries under Real-time Profiles

NASA Technical Reports Server (NTRS)

Staniewicz, Robert J.; Willson, John; Briscoe, J. Douglas; Rao, Gopalakrishna M.

2004-01-01

This viewgraph presentation gives an update on test results from two 16 cell batteries, one in a simulated Low Earth Orbit (LEO) environment and the other in simulated Geosynchronous Earth Orbit (GEO) environment. The tests measured how voltage and capacity are affected over time by thermal cycling.

MEMS Reaction Control and Maneuvering for Picosat Beyond LEO

NASA Technical Reports Server (NTRS)

Alexeenko, Alina

2016-01-01

The MEMS Reaction Control and Maneuvering for Picosat Beyond LEO project will further develop a multi-functional small satellite technology for low-power attitude control, or orientation, of picosatellites beyond low Earth orbit (LEO). The Film-Evaporation MEMS Tunable Array (FEMTA) concept initially developed in 2013, is a thermal valving system which utilizes capillary forces in a microchannel to offset internal pressures in a bulk fluid. The local vapor pressure is increased by resistive film heating until it exceeds meniscus strength in a nozzle which induces vacuum boiling and provides a stagnation pressure equal to vapor pressure at that point which is used for propulsion. Interplanetary CubeSats can utilize FEMTA for high slew rate attitude corrections in addition to desaturating reaction wheels. The FEMTA in cooling mode can be used for thermal control during high-power communication events, which are likely to accompany the attitude correction. Current small satellite propulsion options are limited to orbit correction whereas picosatellites are lacking attitude control thrusters. The available attitude control systems are either quickly saturated reaction wheels or movable high drag surfaces with long response times.

A Laser Optical System to Remove Low Earth Orbit Space Debris

NASA Astrophysics Data System (ADS)

Phipps, Claude R.; Baker, Kevin L.; Libby, Stephen B.; Liedahl, Duane A.; Olivier, Scot S.; Pleasance, Lyn D.; Rubenchik, Alexander; Nikolaev, Sergey; Trebes, James E.; George, Victor E.; Marrcovici, Bogdan; Valley, Michael T.

2013-08-01

Collisions between existing Low Earth Orbit (LEO) debris are now a main source of new debris, threatening future use of LEO space. As solutions, flying up and interacting with each object is inefficient due to the energy cost of orbit plane changes, while debris removal systems using blocks of aerogel or gas-filled balloons are prohibitively expensive. Furthermore, these solutions to the debris problem address only large debris, but it is also imperative to remove 10-cm-class debris. In Laser-Orbital-Debris-Removal (LODR), a ground-based pulsed laser makes plasma jets on LEO debris objects, slowing them slightly, and causing them to re-enter the atmosphere and burn up. LODR takes advantage of recent advances in pulsed lasers, large mirrors, nonlinear optics and acquisition systems. LODR is the only solution that can address both large and small debris. International cooperation is essential for building and operating such a system. We also briefly discuss the orbiting laser debris removal alternative.

New NASA SEE LEO Spacecraft Charging Design Guidelines: How to Survive in LEO Rather than GEO

NASA Technical Reports Server (NTRS)

Ferguson, Dale C.; Hillard, G. Barry

2004-01-01

It has been almost two solar cycles since the GEO Guidelines of Purvis et al (1984) were published. In that time, interest in high voltage LEO systems has increased. The correct and conventional wisdom has been that LEO conditions are sufficiently different from GEO that the GEO Guidelines (and other GEO and POLAR documents produced since then) should not be used for LEO spacecraft. Because of significant recent GEO spacecraft failures that have been shown in ground testing to be likely to also occur on LEO spacecraft, the SEE program commissioned the production of the new LEO Spacecraft Charging Design Guidelines (hereafter referred to as the LEO Guidelines). Now available in CD-ROM form, the LEO Guidelines highlight mitigation techniques to prevent spacecraft arcing on LEO solar arrays and other systems. We compare and contrast the mitigation techniques for LEO and GEO in this paper. We also discuss the extensive bibliography included in the LEO Guidelines, so results can be found in their primary sources.

Orbit Determination and Maneuver Detection Using Event Representation with Thrust-Fourier-Coefficients

NASA Astrophysics Data System (ADS)

Lubey, D.; Ko, H.; Scheeres, D.

The classical orbit determination (OD) method of dealing with unknown maneuvers is to restart the OD process with post-maneuver observations. However, it is also possible to continue the OD process through such unknown maneuvers by representing those unknown maneuvers with an appropriate event representation. It has been shown in previous work (Ko & Scheeres, JGCD 2014) that any maneuver performed by a satellite transitioning between two arbitrary orbital states can be represented as an equivalent maneuver connecting those two states using Thrust-Fourier-Coefficients (TFCs). Event representation using TFCs rigorously provides a unique control law that can generate the desired secular behavior for a given unknown maneuver. This paper presents applications of this representation approach to orbit prediction and maneuver detection problem across unknown maneuvers. The TFCs are appended to a sequential filter as an adjoint state to compensate unknown perturbing accelerations and the modified filter estimates the satellite state and thrust coefficients by processing OD across the time of an unknown maneuver. This modified sequential filter with TFCs is capable of fitting tracking data and maintaining an OD solution in the presence of unknown maneuvers. Also, the modified filter is found effective in detecting a sudden change in TFC values which indicates a maneuver. In order to illustrate that the event representation approach with TFCs is robust and sufficiently general to be easily adjustable, different types of measurement data are processed with the filter in a realistic LEO setting. Further, cases with mis-modeling of non-gravitational force are included in our study to verify the versatility and efficiency of our presented algorithm. Simulation results show that the modified sequential filter with TFCs can detect and estimate the orbit and thrust parameters in the presence of unknown maneuvers with or without measurement data during maneuvers. With no measurement

Need for expanded environmental measurement capabilities in geosynchronous Earth orbit

NASA Technical Reports Server (NTRS)

Mercanti, Enrico P.

1991-01-01

The proliferation of environmental satellites in low altitude earth orbit (LEO) has demonstrated the usefulness of earth remote sensing from space. As use of the technology grows, the limitations of LEO missions become more apparent. Many inadequacies can be met by remote sensing from geosynchronous earth orbits (GEO) that can provide high temporal resolution, consistent viewing of specific earth targets, long sensing dwell times with varying sun angles, stereoscopic coverage, and correlative measurements with ground and LEO observations. An environmental platform in GEO is being studied by NASA. Small research satellite missions in GEO were studied (1990) at GSFC. Some recent independent assessments of NASA Earth Science Programs recommend accelerating the earlier deployment of smaller missions.

Orbit Determination of LEO Satellites for a Single Pass through a Radar: Comparison of Methods

NASA Technical Reports Server (NTRS)

Khutorovsky, Z.; Kamensky, S.; Sbytov, N.; Alfriend, K. T.

2007-01-01

The problem of determining the orbit of a space object from measurements based on one pass through the field of view of a radar is not a new one. Extensive research in this area has been carried out in the USA and Russia since the late 50s when these countries started the development of ballistic missile defense (BMD) and Early Warning systems. In Russia these investigations got additional stimulation in the early 60s after the decision to create a Space Surveillance System, whose primary task would be the maintenance of the satellite catalog. These problems were the focus of research interest until the middle 70s when the appropriate techniques and software were implemented for all radars. Then for more than 20 years no new research papers appeared on this subject. This produced an impression that all the problems of track determination based on one pass had been solved and there was no need for further research. In the late 90s interest in this problem arose again in relation to the following. It was estimated that there would be greater than 100,000 objects with size greater than 1-2 cm and collision of an operational spacecraft with any of these objects could have catastrophic results. Thus, for prevention of hazardous approaches and collisions with valuable spacecraft the existing satellite catalog should be extended by at least an order of magnitude This is a very difficult scientific and engineering task. One of the issues is the development of data fusion procedures and the software capable of maintaining such a huge catalog in near real time. The number of daily processed measurements (of all types, radar and optical) for such a system may constitute millions, thus increasing the number of measurements by at least an order of magnitude. Since we will have ten times more satellites and measurements the computer effort required for the correlation of measurements will be two orders of magnitude greater. This could create significant problems for processing

Outgassing products from orbiter TPS materials

NASA Technical Reports Server (NTRS)

Harvey, Gale A.; Lash, Tom J.; Rawls, J. Richard

1995-01-01

The Space Transportation System (STS) orbiters are known to be significant sources of outgassing in low earth orbit (LEO). Infrared and mass spectra of residues and outgassing from orbiter thermal protection tile and an external blanket are presented. Several sources of methyl and phenyl methyl silicones are identified. About fifty pounds of silicones are estimated to be outgassed during an STS mission.

LEO to ground optical communications from a small satellite platform

NASA Astrophysics Data System (ADS)

Rose, T. S.; Janson, S. W.; LaLumondiere, S.; Werner, N.; Hinkley, D. H.; Rowen, D. W.; Fields, R. A.; Welle, R. P.

2015-03-01

A pair of 2.2 kg CubeSats using COTS hardware is being developed for a proof-of-principle optical communications demo from a 450-600 km LEO orbit to ground. The 10x10x15 cm platform incorporates a 25% wall-plug efficient 10-W Yb fiber transmitter emitting at 1.06 μm. Since there are no gimbals on board, the entire spacecraft is body-steered toward the ground station. The pointing accuracy of the LEO craft, which governs the data rate capability, is expected to be ~ 0.1-0.2 deg. Two optical ground stations, located at the Mt. Wilson observatory, have receiver apertures of 30 and 80 cm. Launch of the CubeSat pair is anticipated to be mid to late 2015.

The LEO Archipelago: A system of earth-rings for communications, mass-transport to space, solar power, and control of global warming

NASA Astrophysics Data System (ADS)

Meulenberg, Andrew; Karthik Balaji, P. S.

2011-06-01

Man's quest to get into space is hindered by major problems (e.g., system-development and capital costs, expense of putting mass into orbit, trapped-radiation belts, and environmental impact of a large increase in rocket launches). A multi-purpose low-earth-orbit system of rings circling the earth - the "LEO ARCHIPELAGO TM" - is proposed as a means of solving or bypassing many of them. A fiber-optic ring about the earth would be an initial testing and developmental stage for the Ring Systems, while providing cash-flow through a LEO-based, high-band-width, world-wide communication system. A low-earth-orbit-based space-elevator system, "Sling-on-a-Ring TM", is proposed as the crucial developmental stage of the LEO Archipelago. Being a LEO-based heavy-mass lifter, rather than earth- or GEO-based, it is much less massive and therefore less costly than other proposed space-elevators. With the advent of lower-cost, higher-mass transport to orbit, the options for further space development (e.g., space solar power, radiation, and space-debris dampers, sun shades, and permanent LEO habitation) are greatly expanded. This paper provides an update of the Sling-on-a-Ring concept in terms of new materials, potential applications, and trade-offs associated with an earlier model. The impact of Colossal Carbon Tubes, CCT, a new material with high tensile strength, extremely-low density, and other favorable properties, and other new technologies (e.g., solar-powered lasers, power beaming to near-space and earth, and thermal-control systems) on the development of associated LEO-Ring systems is also explored. The material's effect on the timeline for the system development indicates the feasibility of near-term implementation of the system (possibly within the decade). The Sling-on-a-Ring can provide a less-expensive, environment-friendly mode of access to space. This would pave the way (via eventual operation at >1000 t per day by 2050) for large scale development of space

A novel data association scheme for LEO space debris surveillance based on a double fence radar system

NASA Astrophysics Data System (ADS)

Huang, Jian; Hu, Weidong; Xin, Qin; Guo, Weiwei

2012-12-01

The increasing amount of space debris threatens to seriously deteriorate and damage space-based instruments in Low Earth Orbit (LEO) environments. Therefore, LEO space debris surveillance systems must be developed to provide situational awareness in space and issue warnings of collisions with LEO space debris. In this paper, a double fence radar system is proposed as an emerging paradigm for LEO space debris surveillance. This system exhibits several unique and promising characteristics compared with existing surveillance systems. In this paper, we also investigate the data association scheme for LEO space debris surveillance based on a double fence radar system. We also perform a theoretical analysis of the performance of our proposed scheme. The superiority and the effectiveness of our novel data association scheme is demonstrated by experimental results. The data used in our experiments is the LEO space debris catalog produced by the North American Air Defense Command (NORAD) up to 2009, especially for scenarios with high densities of LEO space debris, which were primarily produced by the collisions between Iridium 33 and Cosmos 2251. We hope that our work will stimulate and benefit future work on LEO space debris surveillance approaches and enable construction of the double fence radar system.

Reinforcement learning for resource allocation in LEO satellite networks.

PubMed

Usaha, Wipawee; Barria, Javier A

2007-06-01

In this paper, we develop and assess online decision-making algorithms for call admission and routing for low Earth orbit (LEO) satellite networks. It has been shown in a recent paper that, in a LEO satellite system, a semi-Markov decision process formulation of the call admission and routing problem can achieve better performance in terms of an average revenue function than existing routing methods. However, the conventional dynamic programming (DP) numerical solution becomes prohibited as the problem size increases. In this paper, two solution methods based on reinforcement learning (RL) are proposed in order to circumvent the computational burden of DP. The first method is based on an actor-critic method with temporal-difference (TD) learning. The second method is based on a critic-only method, called optimistic TD learning. The algorithms enhance performance in terms of requirements in storage, computational complexity and computational time, and in terms of an overall long-term average revenue function that penalizes blocked calls. Numerical studies are carried out, and the results obtained show that the RL framework can achieve up to 56% higher average revenue over existing routing methods used in LEO satellite networks with reasonable storage and computational requirements.

Satellite Power Systems (SPS) concept definition study. Volume 5: Transportation and operations analysis. [heavy lift launch and orbit transfer vehicles for orbital assembly

NASA Technical Reports Server (NTRS)

Hanley, G.

1978-01-01

The development of transportation systems to support the operations required for the orbital assembly of a 5-gigawatt satellite is discussed as well as the construction of a ground receiving antenna (rectenna). Topics covered include heavy lift launch vehicle configurations for Earth-to LEO transport; the use of chemical, nuclear, and electric orbit transfer vehicles for LEO to GEO operations; personnel transport systems; ground operations; end-to-end analysis of the construction, operation, and maintenance of the satellite and rectenna; propellant production and storage; and payload packaging.

New NASA SEE LEO Spacecraft Charging Design Guidelines: How to Survive in LEO Rather Than GEO

NASA Technical Reports Server (NTRS)

Ferguson, Dale C.; Hillard, G. Barry

2003-01-01

It has been almost two solar cycles since the 1984 GEO Guidelines of Purvis, Garrett, Whittlesey, and Stevens were published. In that time, interest in high voltage LEO systems has increased. Correct and conventional wisdom has been that LEO conditions are sufficiently different from GEO that the GEO Guidelines (and other GEO and POLAR documents produced since then) should not be used for LEO spacecraft. Because of significant recent GEO spacecraft failures that have been shown in ground testing to be likely to also occur on LEO spacecraft, the SEE program commissioned the production of the new LEO Spacecraft Charging Design Guidelines. Now available in CD-ROM form, the LEO Guidelines highlight mitigation techniques to prevent spacecraft arcing on LEO solar arrays and other systems. We compare and contrast the mitigation techniques for LEO and GEO in this paper. We also discuss the extensive bibliography included in the LEO Guidelines, so results can be found in their primary sources.

Particle flows to shape and voltage surface discontinuities in the electron sheath surrounding a high voltage solar array in LEO

NASA Technical Reports Server (NTRS)

Metz, Roger N.

1991-01-01

This paper discusses the numerical modeling of electron flows from the sheath surrounding high positively biased objects in LEO (Low Earth Orbit) to regions of voltage or shape discontinuity on the biased surfaces. The sheath equations are derived from the Two-fluid, Warm Plasma Model. An equipotential corner and a plane containing strips of alternating voltage bias are treated in two dimensions. A self-consistent field solution of the sheath equations is outlined and is pursued through one cycle. The electron density field is determined by numerical solution of Poisson's equation for the electrostatic potential in the sheath using the NASCAP-LEO relation between electrostatic potential and charge density. Electron flows are calculated numerically from the electron continuity equation. Magnetic field effects are not treated.

A Novel Spacecraft Charge Monitor for LEO

NASA Technical Reports Server (NTRS)

Goembel, Luke

2004-01-01

Five years ago we introduced a new method for measuring spacecraft chassis floating potential relative to the space plasma (absolute spacecraft potential) in low Earth orbit. The method, based on a straightforward interpretation of photoelectron spectra, shows promise for numerous applications, but has not yet been tried. In the interest of testing the method, and ultimately supplying another tool for measuring absolute spacecraft charge, we are producing a flight prototype Spacecraft Charge Monitor (SCM) with support from NASA's Small Business Innovation Research (SBIR) program. Although insight into the technique came from data collected in space over two decades ago, very little data are available. The data indicate that it may be possible to determine spacecraft floating potential to within 0.1 volt each with the SCM second under certain conditions. It is debatable that spacecraft floating potential has ever been measured with such accuracy. The compact, easily deployed SCM also offers the advantage of long-term stability in calibration. Accurate floating potential determinations from the SCM could be used to correct biases in space plasma measurements and evaluate charge mitigation and/or sensing devices. Although this paper focuses on the device's use in low Earth orbit (LEO), the device may also be able to measure spacecraft charge at higher altitudes, in the solar wind, and in orbits around other planets. The flight prototype SCM we are producing for delivery to NASA in the third quarter of 2004 will measure floating potential from 0 to -150 volts with 0.1 volt precision, weigh approximately 600-700 grams, consume approximately 2 watts, and will measure approximately 8 x 10 x 17 cm.

Impact of GPS antenna phase center and code residual variation maps on orbit and baseline determination of GRACE

NASA Astrophysics Data System (ADS)

Mao, X.; Visser, P. N. A. M.; van den IJssel, J.

2017-06-01

Precision Orbit Determination (POD) is a prerequisite for the success of many Low Earth Orbiting (LEO) satellite missions. With high-quality, dual-frequency Global Positioning System (GPS) receivers, typically precisions of the order of a few cm are possible for single-satellite POD, and of a few mm for relative POD of formation flying spacecraft with baselines up to hundreds of km. To achieve the best precision, the use of Phase Center Variation (PCV) maps is indispensable. For LEO GPS receivers, often a-priori PCV maps are obtained by a pre-launch ground campaign, which is not able to represent the real space-borne environment of satellites. Therefore, in-flight calibration of the GPS antenna is more widely conducted. This paper shows that a further improvement is possible by including the so-called Code Residual Variation (CRV) maps in absolute/undifferenced and relative/Double-differenced (DD) POD schemes. Orbit solutions are produced for the GRACE satellite formation for a four months test period (August-November, 2014), demonstrating enhanced orbit precision after first using the in-flight PCV maps and a further improvement after including the CRV maps. The application of antenna maps leads to a better consistency with independent Satellite Laser Ranging (SLR) and K-band Ranging (KBR) low-low Satellite-to-Satellite Tracking (ll-SST) observations. The inclusion of the CRV maps results also in a much better consistency between reduced-dynamic and kinematic orbit solutions for especially the cross-track direction. The improvements are largest for GRACE-B, where a cross-talk between the GPS main antenna and the occultation antenna yields higher systematic observation residuals. For high-precision relative POD which necessitates DD carrier-phase ambiguity fixing, in principle frequency-dependent PCV maps would be required. To this aim, use is made of an Extended Kalman Filter (EKF) that is capable of optimizing relative spacecraft dynamics and iteratively fixing

Minimum fuel coplanar aeroassisted orbital transfer using collocation and nonlinear programming

NASA Technical Reports Server (NTRS)

Shi, Yun Yuan; Young, D. H.

1991-01-01

The fuel optimal control problem arising in coplanar orbital transfer employing aeroassisted technology is addressed. The mission involves the transfer from high energy orbit (HEO) to low energy orbit (LEO) without plane change. The basic approach here is to employ a combination of propulsive maneuvers in space and aerodynamic maneuvers in the atmosphere. The basic sequence of events for the coplanar aeroassisted HEO to LEO orbit transfer consists of three phases. In the first phase, the transfer begins with a deorbit impulse at HEO which injects the vehicle into a elliptic transfer orbit with perigee inside the atmosphere. In the second phase, the vehicle is optimally controlled by lift and drag modulation to satisfy heating constraints and to exit the atmosphere with the desired flight path angle and velocity so that the apogee of the exit orbit is the altitude of the desired LEO. Finally, the second impulse is required to circularize the orbit at LEO. The performance index is maximum final mass. Simulation results show that the coplanar aerocapture is quite different from the case where orbital plane changes are made inside the atmosphere. In the latter case, the vehicle has to penetrate deeper into the atmosphere to perform the desired orbital plane change. For the coplanar case, the vehicle needs only to penetrate the atmosphere deep enough to reduce the exit velocity so the vehicle can be captured at the desired LEO. The peak heating rates are lower and the entry corridor is wider. From the thermal protection point of view, the coplanar transfer may be desirable. Parametric studies also show the maximum peak heating rates and the entry corridor width are functions of maximum lift coefficient. The problem is solved using a direct optimization technique which uses piecewise polynomial representation for the states and controls and collocation to represent the differential equations. This converts the optimal control problem into a nonlinear programming problem

Transfer Orbit Plasma Interaction Experiment (TROPIX)

NASA Astrophysics Data System (ADS)

Hickman, Mark

Viewgraphs on the Transfer Orbit Plasma Interaction Experiment (TROPIX) are presented. Objectives of this research are (1) to map the charged particles in Earth's magnetosphere from LEO to GEO at high inclinations; (2) to measure plasma current collection and resulting shifts in vehicle electrical potential from LEO to GEO across range of orbital inclinations; (3) to study spacecraft interaction with plasma environment using solar electric propulsion (SEP) thrusters as plasma contactors; (4) to measure array degradation over mission duration; (5) to evaluate the potential of various microelectronics, spacecraft components, and instruments for future space missions; and (6) to demonstrate SEP technology applied to a flight vehicle. An overview of TROPIX is presented.

What Happened to Leo P's Metals?

NASA Astrophysics Data System (ADS)

Kohler, Susanna

2015-12-01

Measurements of metal abundances in galaxies present a conundrum: compared to expectations, there are not nearly enough metals observed within galaxies. New observations of a nearby dwarf galaxy may help us understand where this enriched material went.Removal ProcessesStar formation is responsible for the build-up of metals (elements heavier than helium) in a galaxy. But when we use a galaxys star-formation history to estimate the amount of enriched material it should contain, our predictions are inconsistent with measured abundances: large galaxies contain only about 2025% of the expected metals, and small dwarf galaxies contain as little as 1%!So what happens to galaxies metals after they have been formed? The favored explanation is that metals are removed from galaxies via stellar feedback: stars that explode in violent supernovae can drive high-speed winds, expelling the enriched material from a galaxy. This process should be more efficient in low-mass galaxies due to their smaller gravitational wells, which would explain why low-mass galaxies have especially low metallicities.But external processes may also contribute to the removal of metals, such as tidal stripping during interactions between galaxies. To determine the role of stellar feedback alone, an ideal test would be to observe an isolated low-mass, star-forming galaxy i.e., one that is not affected by external processes.Luckily, such an isolated, low-mass galaxy has recently been discovered just outside of the Local Group: Leo P, a gas-rich dwarf galaxy with a total stellar mass of 5.6 x 105 solar masses.Isolated ResultsPercentage of oxygen lost in Leo P compared to the percentage of metals lost in three other, similar-size dwarfs that are not isolated. If the gas-phase oxygen in Leo P were removed, Leo Ps measurements would be consistent with those of the other dwarfs. [McQuinn et al. 2015]Led by Kristen McQuinn (University of Minnesota, University of Texas at Austin), a team of researchers has used

Mitigating Aviation Communication and Satellite Orbit Operations Surprises from Adverse Space Weather

NASA Technical Reports Server (NTRS)

Tobiska, W. Kent

2008-01-01

Adverse space weather affects operational activities in aviation and satellite systems. For example, large solar flares create highly variable enhanced neutral atmosphere and ionosphere electron density regions. These regions impact aviation communication frequencies as well as precision orbit determination. The natural space environment, with its dynamic space weather variability, is additionally changed by human activity. The increase in orbital debris in low Earth orbit (LEO), combined with lower atmosphere CO2 that rises into the lower thermosphere and causes increased cooling that results in increased debris lifetime, adds to the environmental hazards of navigating in near-Earth space. This is at a time when commercial space endeavors are posed to begin more missions to LEO during the rise of the solar activity cycle toward the next maximum (2012). For satellite and aviation operators, adverse space weather results in greater expenses for orbit management, more communication outages or aviation and ground-based high frequency radio used, and an inability to effectively plan missions or service customers with space-based communication, imagery, and data transferal during time-critical activities. Examples of some revenue-impacting conditions and solutions for mitigating adverse space weather are offered.

Orbit determination for ISRO satellite missions

NASA Astrophysics Data System (ADS)

Rao, Ch. Sreehari; Sinha, S. K.

Indian Space Research Organisation (ISRO) has been successful in using the in-house developed orbit determination and prediction software for satellite missions of Bhaskara, Rohini and APPLE. Considering the requirements of satellite missions, software packages are developed, tested and their accuracies are assessed. Orbit determination packages developed are SOIP, for low earth orbits of Bhaskara and Rohini missions, ORIGIN and ODPM, for orbits related to all phases of geo-stationary missions and SEGNIP, for drift and geo-stationary orbits. Software is tested and qualified using tracking data of SIGNE-3, D5-B, OTS, SYMPHONIE satellites with the help of software available with CNES, ESA and DFVLR. The results match well with those available from these agencies. These packages have supported orbit determination successfully throughout the mission life for all ISRO satellite missions. Member-Secretary

Supportability for Beyond Low Earth Orbit Missions

NASA Technical Reports Server (NTRS)

Crillo, William M.; Goodliff, Kandyce E.; Aaseng, Gordon; Stromgren, Chel; Maxwell, Andrew J.

2011-01-01

Exploration beyond Low Earth Orbit (LEO) presents many unique challenges that will require changes from current Supportability approaches. Currently, the International Space Station (ISS) is supported and maintained through a series of preplanned resupply flights, on which spare parts, including some large, heavy Orbital Replacement Units (ORUs), are delivered to the ISS. The Space Shuttle system provided for a robust capability to return failed components to Earth for detailed examination and potential repair. Additionally, as components fail and spares are not already on-orbit, there is flexibility in the transportation system to deliver those required replacement parts to ISS on a near term basis. A similar concept of operation will not be feasible for beyond LEO exploration. The mass and volume constraints of the transportation system and long envisioned mission durations could make it difficult to manifest necessary spares. The supply of on-demand spare parts for missions beyond LEO will be very limited or even non-existent. In addition, the remote nature of the mission, the design of the spacecraft, and the limitations on crew capabilities will all make it more difficult to maintain the spacecraft. Alternate concepts of operation must be explored in which required spare parts, materials, and tools are made available to make repairs; the locations of the failures are accessible; and the information needed to conduct repairs is available to the crew. In this paper, ISS heritage information is presented along with a summary of the challenges of beyond LEO missions. A number of Supportability issues are discussed in relation to human exploration beyond LEO. In addition, the impacts of various Supportability strategies will be discussed. Any measure that can be incorporated to reduce risk and improve mission success should be evaluated to understand the advantages and disadvantages of implementing those measures. Finally, an effort to model and evaluate

Risk Analysis of On-Orbit Spacecraft Refueling Concepts

NASA Technical Reports Server (NTRS)

Cirillo, William M.; Stromgren, Chel; Cates, Grant R.

2010-01-01

On-orbit refueling of spacecraft has been proposed as an alternative to the exclusive use of Heavy-lift Launch Vehicles to enable human exploration beyond Low Earth Orbit (LEO). In these scenarios, beyond LEO spacecraft are launched dry (without propellant) or partially dry into orbit, using smaller or fewer element launch vehicles. Propellant is then launched into LEO on separate launch vehicles and transferred to the spacecraft. Refueling concepts are potentially attractive because they reduce the maximum individual payload that must be placed in Earth orbit. However, these types of approaches add significant complexity to mission operations and introduce more uncertainty and opportunities for failure to the mission. In order to evaluate these complex scenarios, the authors developed a Monte Carlo based discrete-event model that simulates the operational risks involved with such strategies, including launch processing delays, transportation system failures, and onorbit element lifetimes. This paper describes the methodology used to simulate the mission risks for refueling concepts, the strategies that were evaluated, and the results of the investigation. The results of the investigation show that scenarios that employ refueling concepts will likely have to include long launch and assembly timelines, as well as the use of spare tanker launch vehicles, in order to achieve high levels of mission success through Trans Lunar Injection.

Information Measures for Statistical Orbit Determination

ERIC Educational Resources Information Center

Mashiku, Alinda K.

2013-01-01

The current Situational Space Awareness (SSA) is faced with a huge task of tracking the increasing number of space objects. The tracking of space objects requires frequent and accurate monitoring for orbit maintenance and collision avoidance using methods for statistical orbit determination. Statistical orbit determination enables us to obtain…

Spaceport aurora: An orbiting transportation node

NASA Technical Reports Server (NTRS)

1990-01-01

With recent announcements of the development of permanently staffed facilities on the Moon and Mars, the national space plan is in need of an infrastructure system for transportation and maintenance. A project team at the University of Houston College of Architecture and the Sasakawa International Center for Space Architecture, recently examined components for a low Earth orbit (LEO) transportation node that supports a lunar build-up scenario. Areas of investigation included identifying transportation node functions, identifying existing space systems and subsystems, analyzing variable orbits, determining logistics strategies for maintenance, and investigating assured crew return systems. The information resulted in a requirements definition document, from which the team then addressed conceptual designs for a LEO transportation node. The primary design drivers included: orbital stability, maximizing human performance and safety, vehicle maintainability, and modularity within existing space infrastructure. For orbital stability, the power tower configuration provides a gravity gradient stabilized facility and serves as the backbone for the various facility components. To maximize human performance, human comfort is stressed through zoning of living and working activities, maintaining a consistent local vertical orientation, providing crew interaction and viewing areas and providing crew return vehicles. Vehicle maintainability is accomplished through dual hangars, dual work cupolas, work modules, telerobotics and a fuel depot. Modularity is incorporated using Space Station Freedom module diameter, Space Station Freedom standard racks, and interchangeable interior partitions. It is intended that the final design be flexible and adaptable to provide a facility prototype that can service multiple mission profiles using modular space systems.

Low Earth Orbit satellite/terrestrial mobile service compatibility

NASA Technical Reports Server (NTRS)

Sheriff, Ray E.; Gardiner, John G.

1993-01-01

Currently the geostationary type of satellite is the only one used to provide commercial mobile-satellite communication services. Low earth orbit (LEO) satellite systems are now being proposed as a future alternative. By the implementation of LEO satellite systems, predicted at between 5 and 8 years time, mobile space/terrestrial technology will have progressed to the third generation stage of development. This paper considers the system issues that will need to be addressed when developing a dual mode terminal, enabling access to both terrestrial and LEO satellite systems.

High-voltage plasma interactions calculations using NASCAP/LEO

NASA Technical Reports Server (NTRS)

Mandell, M. J.; Katz, I.

1990-01-01

This paper reviews four previous simulations (two laboratory and two space-flight) of interactions of a high-voltage spacecraft with a plasma under low-earth orbit conditions, performed using a three-dimensional computer code NASCAP/LEO. Results show that NASCAP/LEO can perform meaningful simulations of high-voltage plasma interactions taking into account three-dimensional effects of geometry, spacecraft motion, and magnetic field. Two new calculations are presented: (1) for current collection by 1-mm pinholes in wires (showing that a pinhole in a wire can collect far more current than a similar pinhole in a flat plate); and (2) current collection by Charge-2 mother vehicle launched in December 1985. It is shown that the Charge-2 calculations predicted successfully ion collection at negative bias, the floating potential of a probe outside or inside the sheath under negative bias conditions, and magnetically limited electron collection under electron beam operation at high altitude.

Implementation of National Space Policy on US Air Force End of Life Operations and Orbital Debris Mitigation

DTIC Science & Technology

2012-06-01

procedures have been driven by the rising significance of the orbital debris problem in Low Earth Orbit (LEO). Therefore current EOL plans are...does not display a currently valid OMB control number. 1. REPORT DATE JUN 2012 2. REPORT TYPE 3. DATES COVERED 00-00-2012 to 00-00-2012 4...by the rising significance of the orbital debris problem in Low Earth Orbit (LEO). Therefore current EOL plans are written largely with the aim of

The Komplast Experiment: Space Environmental Effects after 12 Years in LEO (and Counting)

NASA Technical Reports Server (NTRS)

Shaevich, Sergei K.; Aleksandrov, Nicholai G.; Shumov, Andrei E.; Novikov, L. S.; Alred, John A.; Shindo, David J.; Kravchenko, Michael; Golden, Johnny L.

2014-01-01

The Komplast materials experiment was designed by the Khrunichev Space Center, together with other Russian scientific institutes, and has been carried out by Mission Control Moscow since 1998. The purpose is to study the effect of the low earth orbit (LEO) environment on exposed samples of various spacecraft materials. The Komplast experiment began with the launch of the first International Space Station (ISS) module on November 20, 1998. Two of eight experiment panels were retrieved during Russian extravehicular activity in February 2011 after 12 years of LEO exposure, and were subsequently returned to Earth by Space Shuttle "Discovery" on the STS-133/ULF-5 mission. The retrieved panels contained an experiment to detect micrometeoroid and orbital debris (MMOD) impacts, radiation sensors, a temperature sensor, several pieces of electrical cable, both carbon composite and adhesive-bonded samples, and many samples made from elastomeric and fluoroplastic materials. Our investigation is complete and a summary of the results obtained from this uniquely long-duration exposure experiment will be presented.

Simultaneous orbit determination

NASA Technical Reports Server (NTRS)

Wright, J. R.

1988-01-01

Simultaneous orbit determination is demonstrated using live range and Doppler data for the NASA/Goddard tracking configuration defined by the White Sands Ground Terminal (WSGT), the Tracking and Data Relay Satellite (TDRS), and the Earth Radiation Budget Satellite (ERBS). A physically connected sequential filter-smoother was developed for this demonstration. Rigorous necessary conditions are used to show that the state error covariance functions are realistic; and this enables the assessment of orbit estimation accuracies for both TDRS and ERBS.

Design and Fabrication of DebriSat - A Representative LEO Satellite for Improvements to Standard Satellite Breakup Models

NASA Technical Reports Server (NTRS)

Clark, S.; Dietrich, A.; Fitz-Coy, N.; Weremeyer, M.; Liou, J.-C.

2012-01-01

This paper discusses the design and fabrication of DebriSat, a 50 kg satellite developed to be representative of a modern low Earth orbit satellite in terms of its components, materials used, and fabrication procedures. DebriSat will be the target of a future hypervelocity impact experiment to determine the physical characteristics of debris generated after an on-orbit collision of a modern LEO satellite. The major ground-based satellite impact experiment used by DoD and NASA in their development of satellite breakup models was SOCIT, conducted in 1992. The target used for that experiment was a Navy transit satellite (40 cm, 35 kg) fabricated in the 1960's. Modern satellites are very different in materials and construction techniques than those built 40 years ago. Therefore, there is a need to conduct a similar experiment using a modern target satellite to improve the fidelity of the satellite breakup models. To ensure that DebriSat is truly representative of typical LEO missions, a comprehensive study of historical LEO satellite designs and missions within the past 15 years for satellites ranging from 1 kg to 5000 kg was conducted. This study identified modern trends in hardware, material, and construction practices utilized in recent LEO missions. Although DebriSat is an engineering model, specific attention is placed on the quality, type, and quantity of the materials used in its fabrication to ensure the integrity of the outcome. With the exception of software, all other aspects of the satellite s design, fabrication, and assembly integration and testing will be as rigorous as that of an actual flight vehicle. For example, to simulate survivability of launch loads, DebriSat will be subjected to a vibration test. As well, the satellite will undergo thermal vacuum tests to verify that the components and overall systems meet typical environmental standards. Proper assembly and integration techniques will involve comprehensive joint analysis, including the precise

TT and C - First TDRSS, Then Commercial GEO and Big LEO and Now through LEO

NASA Technical Reports Server (NTRS)

Morgan, Dwayne; Bull, Barton; Grant, Charles; Streich, Ronald; Powers, Edward I. (Technical Monitor)

2001-01-01

The advent of low earth orbit (LEO) commercial communications satellites provides an opportunity to dramatically reduce Telemetry Tracking and Control (TT&C) costs of launch vehicles and Unpiloted Aerial Vehicles (UAVs) by reducing or eliminating ground infrastructure. Personnel from the Goddard Space Flight Center Wallops Flight Facility (GSFC/WFF) in Virginia have successfully used commercial GEO & Big LEO communications satellites for Long Duration Balloon flight TT&C. In addition, TDRSS capability for these balloons has been developed by WFF for the Ultra Long Duration Balloons with the first test flight launch in January 2001 for one global circumnavigation at 120,000 feet altitude launched from Alice Springs. Australia. Numerous other low cost applications can new utilize the commercial LEO satellites for TT&C. The Flight Modern became a GSFC/WFF Advanced Range Technology Initiative (ARTI) in an effort to streamline TT&C capability to the user community at low cost. Phase I ground tests of The Flight Modem verified downlink communications quality of service and measured transmission latencies. These tests were completed last year, Phase II consisting of aircraft flight tests provide much of the data presented in this paper. Phase III of the Flight Modern baseline test program is a demonstration of the ruggedized version of the WFF Flight Modem flown on one sounding rocket launched from Sweden. Flights of opportunity have been and are being actively pursued with other centers, ranges and users at universities. The WFF goal is to reduce TT&C costs by providing a low cost COTS Flight Modem with a User Handbook containing system capability and limitation descriptions. Additionally, since data transmission is by packetized Internet Protocol (IP), data can be received and commands initialed from practically any location with no infrastructure. The WFF, like most ranges, has been using GPS receivers on sounding rockets and long duration balloons for several years

An analysis of the effect of aeroassist maneuvers on orbital transfer vehicle performance

NASA Technical Reports Server (NTRS)

Murphy, Gregory O.; Suit, William T.

1987-01-01

This paper summarizes a Langley Research Summer Scholars (LARSS) research project (Summer 1986) dealing with the topic of the effectiveness of aeroassist maneuvers to accomplish a change in the orbital inclination of an Orbital Transfer Vehicle (OTV). This task was subject to OTV design constraints, chief of which were the axial acceleration and the aerodynamic heating rate limits of the OTV. The use of vehicle thrust to replace lost kinetic energy and, thereby, to increase the maximum possible change in orbital inclination was investigated. A relation between time in the hover orbit and payload to LEO was established. The amount of plane change possible during this type of maneuver was checked for several runs and a possible thrusting procedure to increase the plane change and still get to LEO was suggested. Finally, the sensitivity of various target parameters to controllable independent variables was established, trades between the amount of control allowed, and payload to LEO suggested.

A Sensitivity Study on the Effectiveness of Active Debris Removal in LEO

NASA Technical Reports Server (NTRS)

Liou, J. C.; Johnson, Nicholas L.

2007-01-01

The near-Earth orbital debris population will continue to increase in the future due to ongoing space activities, on-orbit explosions, and accidental collisions among resident space objects. Commonly adopted mitigation measures, such as limiting postmission orbital lifetimes of satellites to less than 25 years, will slow down the population growth, but may be insufficient to stabilize the environment. The nature of the growth, in the low Earth orbit (LEO) region, is further demonstrated by a recent study where no future space launches were conducted in the environment projection simulations. The results indicate that, even with no new launches, the LEO debris population would remain relatively constant for only the next 50 years. Beyond that, the debris population would begin to increase noticeably, due to the production of collisional debris. Therefore, to better limit the growth of future debris population to protect the environment, remediation option, i.e., removing existing large and massive objects from orbit, needs to be considered. This paper does not intend to address the technical or economical issues for active debris removal. Rather, the objective is to provide a sensitivity study to quantify the effectiveness of various remediation options. A removal criterion based upon mass and collision probability is developed to rank objects at the beginning of each projection year. This study includes simulations with removal rates ranging from 2 to 20 objects per year, starting in the year 2020. The outcome of each simulation is analyzed, and compared with others. The summary of the study serves as a general guideline for future debris removal consideration.

Materials selection for long life in low earth orbit - A critical evaluation of atomic oxygen testing with thermal atom systems

NASA Technical Reports Server (NTRS)

Koontz, S. L.; Albyn, K.; Leger, L.

1990-01-01

The use of thermal atom test methods as a materials selection and screening technique for low-earth orbit (LEO) spacecraft is critically evaluated. The chemistry and physics of thermal atom environments are compared with the LEO environment. The relative reactivities of a number of materials determined in thermal atom environments are compared with those observed in LEO and in high-quality LEO simulations. Reaction efficiencies (cu cm/atom) measured in a new type of thermal atom apparatus are one-thousandth to one ten-thousandth those observed in LEO, and many materials showing nearly identical reactivities in LEO show relative reactivities differing by as much as a factor of eight in thermal atom systems. A simple phenomenological kinetic model for the reaction of oxygen atoms with organic materials can be used to explain the differences in reactivity in different environments. Certain speciic thermal atom test environments can be used as reliable materials screening tools.

Modification of earth-satellite orbits using medium-energy pulsed lasers

DOE Office of Scientific and Technical Information (OSTI.GOV)

Phipps, C.R.

1992-01-01

Laser Impulse Space Propulsion (LISP) has become an attractive concept, due to recent advances in gas laser technology, high-speed segmented mirrors and improved coeffici-ents for momentum coupling to targets in pulsed laser ablation. There are numerous specialized applications of the basic concept to space science-ranging from far-future and high capital cost to the immediate and inexpensive, such as: LEO-LISP (launch of massive objects into low-Earth-Orbit at dramatically improved cost-per-kg relative to present practice); LEGO-LISP (LEO to geosynchronous transfers); LO-LISP) (periodic re-boost of decaying LEO orbits); and LISK (geosynchronous satellite station-keeping). It is unlikely that one type of laser will bemore » best for all scenarios. In this paper, we will focus on the last two applications.« less

Modification of earth-satellite orbits using medium-energy pulsed lasers

DOE Office of Scientific and Technical Information (OSTI.GOV)

Phipps, C.R.

1992-10-01

Laser Impulse Space Propulsion (LISP) has become an attractive concept, due to recent advances in gas laser technology, high-speed segmented mirrors and improved coeffici-ents for momentum coupling to targets in pulsed laser ablation. There are numerous specialized applications of the basic concept to space science-ranging from far-future and high capital cost to the immediate and inexpensive, such as: LEO-LISP (launch of massive objects into low-Earth-Orbit at dramatically improved cost-per-kg relative to present practice); LEGO-LISP (LEO to geosynchronous transfers); LO-LISP) (periodic re-boost of decaying LEO orbits); and LISK (geosynchronous satellite station-keeping). It is unlikely that one type of laser will bemore » best for all scenarios. In this paper, we will focus on the last two applications.« less

In-flight performance analysis of MEMS GPS receiver and its application to precise orbit determination of APOD-A satellite

NASA Astrophysics Data System (ADS)

Gu, Defeng; Liu, Ye; Yi, Bin; Cao, Jianfeng; Li, Xie

2017-12-01

An experimental satellite mission termed atmospheric density detection and precise orbit determination (APOD) was developed by China and launched on 20 September 2015. The micro-electro-mechanical system (MEMS) GPS receiver provides the basis for precise orbit determination (POD) within the range of a few decimetres. The in-flight performance of the MEMS GPS receiver was assessed. The average number of tracked GPS satellites is 10.7. However, only 5.1 GPS satellites are available for dual-frequency navigation because of the loss of many L2 observations at low elevations. The variations in the multipath error for C1 and P2 were estimated, and the maximum multipath error could reach up to 0.8 m. The average code noises are 0.28 m (C1) and 0.69 m (P2). Using the MEMS GPS receiver, the orbit of the APOD nanosatellite (APOD-A) was precisely determined. Two types of orbit solutions are proposed: a dual-frequency solution and a single-frequency solution. The antenna phase center variations (PCVs) and code residual variations (CRVs) were estimated, and the maximum value of the PCVs is 4.0 cm. After correcting the antenna PCVs and CRVs, the final orbit precision for the dual-frequency and single-frequency solutions were 7.71 cm and 12.91 cm, respectively, validated using the satellite laser ranging (SLR) data, which were significantly improved by 3.35 cm and 25.25 cm. The average RMS of the 6-h overlap differences in the dual-frequency solution between two consecutive days in three dimensions (3D) is 4.59 cm. The MEMS GPS receiver is the Chinese indigenous onboard receiver, which was successfully used in the POD of a nanosatellite. This study has important reference value for improving the MEMS GPS receiver and its application in other low Earth orbit (LEO) nanosatellites.

TDRS orbit determination by radio interferometry

NASA Technical Reports Server (NTRS)

Pavloff, Michael S.

1994-01-01

In support of a NASA study on the application of radio interferometry to satellite orbit determination, MITRE developed a simulation tool for assessing interferometry tracking accuracy. The Orbit Determination Accuracy Estimator (ODAE) models the general batch maximum likelihood orbit determination algorithms of the Goddard Trajectory Determination System (GTDS) with the group and phase delay measurements from radio interferometry. ODAE models the statistical properties of tracking error sources, including inherent observable imprecision, atmospheric delays, clock offsets, station location uncertainty, and measurement biases, and through Monte Carlo simulation, ODAE calculates the statistical properties of errors in the predicted satellites state vector. This paper presents results from ODAE application to orbit determination of the Tracking and Data Relay Satellite (TDRS) by radio interferometry. Conclusions about optimal ground station locations for interferometric tracking of TDRS are presented, along with a discussion of operational advantages of radio interferometry.

Multi-Satellite Orbit Determination Using Interferometric Observables with RF Localization Applications

NASA Astrophysics Data System (ADS)

Geeraert, Jeroen L.

Very long baseline interferometry (VLBI) specifically same-beam interferometry (SBI), and dual-satellite geolocation are two fields of research not previously connected. This is due to the different application of each field, SBI is used for relative interplanetary navigation of two satellites while dual-satellite geolocation is used to locate the source of a radio frequency (RF) signal. In this dissertation however, we leverage both fields to create a novel method for multi-satellite orbit determination (OD) using time difference of arrival (TDOA) and frequency difference of arrival (FDOA) measurements. The measurements are double differenced between the satellites and the stations, in so doing, many of the common errors are canceled which can significantly improve measurement precision. Provided with this novel OD technique, the observability is first analyzed to determine the benefits and limitations of this method. In all but a few scenarios the measurements successfully reduce the covariance when examining the Cramer-Rao Lower Bound (CRLB). Reduced observability is encountered with geostationary satellites as their motion with respect to the stations is limited, especially when only one baseline is used. However, when using satellite pairs with greater relative motion with respect to the stations, even satellites that are close to, but not exactly in a geostationary orbit can be estimated accurately. We find that in a strong majority of cases the OD technique provides lower uncertainties and solutions far more accurate than using conventional OD observables such as range and range-rate while also not being affected by common errors and biases. We specifically examine GEO-GEO, GEO-MEO, and GEO-LEO dual-satellite estimation cases. The work is further extended by developing a relative navigation scenario where the chief satellite is assumed to have perfect knowledge, or some small amount of uncertainty considered but not estimated, while estimating the deputy

NASA Orbital Debris Large-Object Baseline Population in ORDEM 3.0

NASA Technical Reports Server (NTRS)

Krisco, Paula H.; Vavrin, A. B.; Anz-Meador, P. D.

2013-01-01

The NASA Orbital Debris Program Office (ODPO) has created and validated high fidelity populations of the debris environment for the latest Orbital Debris Engineering Model (ORDEM 3.0). Though the model includes fluxes of objects 10 um and larger, this paper considers particle fluxes for 1 cm and larger debris objects from low Earth orbit (LEO) through Geosynchronous Transfer Orbit (GTO). These are validated by several reliable radar observations through the Space Surveillance Network (SSN), Haystack, and HAX radars. ORDEM 3.0 populations were designed for the purpose of assisting, debris researchers and sensor developers in planning and testing. This environment includes a background derived from the LEO-to-GEO ENvironment Debris evolutionary model (LEGEND) with a Bayesian rescaling as well as specific events such as the FY-1C anti-satellite test, the Iridium 33/Cosmos 2251 accidental collision, and the Soviet/Russian Radar Ocean Reconnaissance Satellite (RORSAT) sodium-potassium droplet releases. The environment described in this paper is the most realistic orbital debris population larger than 1 cm, to date. We describe derivations of the background population and added specific populations. We present sample validation charts of our 1 cm and larger LEO population against Space Surveillance Network (SSN), Haystack, and HAX radar measurements.

Medium Earth Orbits: Is There a Need for a Third Protected Region?

NASA Technical Reports Server (NTRS)

Johnson, Nicholas L.

2010-01-01

The Inter-Agency Space Debris Coordination Committee (IADC) and the United Nations have adopted the concept of near-Earth regions which should be afforded protection from the accumulation of orbital debris. These regions are low Earth orbit (LEO), which extends up to 2000 km altitude, and geosynchronous orbit (GEO), which includes the volume of space encompassed by 35,786 km +/- 200 km in altitude and +/- 15 degrees in inclination. The region between LEO and GEO is commonly referred to as Medium Earth Orbit (MEO). Although historically a small minority of spacecraft have operated in MEO, the number of such satellites residing in or routinely transiting the zone is increasing. The question thus arises: should MEO be considered an orbital debris protected region? This paper first reviews the characteristics of space systems now utilizing MEO, as well as those anticipated to join them in the near future. MEO is then contrasted with LEO and GEO, both physically and pragmatically. Recommended orbital debris mitigation guidelines for MEO space vehicles are highlighted, and the challenges of spacecraft and launch vehicle stage disposal are recognized. Note is also made of the principal tenets of the United Nations Outer Space Treaty and of recent trends toward de facto partitioning of MEO. Finally, the efficacy and practicality of establishing MEO as a new protected region with regard to orbital debris is addressed.

The GEOS-3 orbit determination investigation

NASA Technical Reports Server (NTRS)

Pisacane, V. L.; Eisner, A.; Yionoulis, S. M.; Mcconahy, R. J.; Black, H. D.; Pryor, L. L.

1978-01-01

The nature and improvement in satellite orbit determination when precise altimetric height data are used in combination with conventional tracking data was determined. A digital orbit determination program was developed that could singly or jointly use laser ranging, C-band ranging, Doppler range difference, and altimetric height data. Two intervals were selected and used in a preliminary evaluation of the altimeter data. With the data available, it was possible to determine the semimajor axis and eccentricity to within several kilometers, in addition to determining an altimeter height bias. When used jointly with a limited amount of either C-band or laser range data, it was shown that altimeter data can improve the orbit solution.

Mass driver reaction engine characteristics and performance in earth orbital transfer missions

NASA Technical Reports Server (NTRS)

Snow, W. R.; Dunbar, R. S.

1982-01-01

Configurations of a typical mass driver reaction engine (MDRE) are presented and its use for delivery of payloads to geosynchronous orbit (GEO) from low earth orbit (LEO) is discussed. Basic rocket equations are developed for LEO to GEO round-trip missions using a single exhaust velocity. It is shown that exhaust velocities in the 5-10 km/sec range (specific impulse of 500-1000 sec) are well suited for mass drivers, minimizing the overall cost of missions. Payload delivery rate fractions show that there is little to be gained by stretching out LEO to GEO transfer times from 90 to 180 days. It therefore pays to use the shorter trip time, approximately doubling the amount of delivered payload during any fixed time of use of the MDRE.

Low Earth Orbiter: Terminal

NASA Technical Reports Server (NTRS)

Kremer, Steven E.; Bundick, Steven N.

1999-01-01

In response to the current government budgetary environment that requires the National Aeronautics and Space Administration (NASA) to do more with less, NASA/Goddard Space Flight Center's Wallops Flight Facility has developed and implemented a class of ground stations known as a Low Earth Orbiter-Terminal (LEO-T). This development thus provides a low-cost autonomous ground tracking service for NASA's customers. More importantly, this accomplishment provides a commercial source to spacecraft customers around the world to purchase directly from the company awarded the NASA contract to build these systems. A few years ago, NASA was driven to provide more ground station capacity for spacecraft telemetry, tracking, and command (TT&C) services with a decreasing budget. NASA also made a decision to develop many smaller, cheaper satellites rather than a few large spacecraft as done in the past. In addition, university class missions were being driven to provide their own TT&C services due to the increasing load on the NASA ground-tracking network. NASA's solution for this ever increasing load was to use the existing large aperture systems to support those missions requiring that level of performance and to support the remainder of the missions with the autonomous LEO-T systems. The LEO-T antenna system is a smaller, cheaper, and fully autonomous unstaffed system that can operate without the existing NASA support infrastructure. The LEO-T provides a low-cost, reliable space communications service to the expanding number of low-earth orbiting missions around the world. The system is also fostering developments that improve cost-effectiveness of autonomous-class capabilities for NASA and commercial space use. NASA has installed three LEO-T systems. One station is at the University of Puerto Rico, the second system is installed at the Poker Flat Research Range near Fairbanks, Alaska, and the third system is installed at NASA's Wallops Flight Facility in Virginia. This paper

Angles-only relative orbit determination in low earth orbit

NASA Astrophysics Data System (ADS)

Ardaens, Jean-Sébastien; Gaias, Gabriella

2018-06-01

The paper provides an overview of the angles-only relative orbit determination activities conducted to support the Autonomous Vision Approach Navigation and Target Identification (AVANTI) experiment. This in-orbit endeavor was carried out by the German Space Operations Center (DLR/GSOC) in autumn 2016 to demonstrate the capability to perform spaceborne autonomous close-proximity operations using solely line-of-sight measurements. The images collected onboard have been reprocessed by an independent on-ground facility for precise relative orbit determination, which served as ultimate instance to monitor the formation safety and to characterize the onboard navigation and control performances. During two months, several rendezvous have been executed, generating a valuable collection of images taken at distances ranging from 50 km to only 50 m. Despite challenging experimental conditions characterized by a poor visibility and strong orbit perturbations, angles-only relative positioning products could be continuously derived throughout the whole experiment timeline, promising accuracy at the meter level during the close approaches. The results presented in the paper are complemented with former angles-only experience gained with the PRISMA satellites to better highlight the specificities induced by different orbits and satellite designs.

Charged aerodynamics of a Low Earth Orbit cylinder

NASA Astrophysics Data System (ADS)

Capon, C. J.; Brown, M.; Boyce, R. R.

2016-11-01

This work investigates the charged aerodynamic interaction of a Low Earth Orbiting (LEO) cylinder with the ionosphere. The ratio of charge to neutral drag force on a 2D LEO cylinder with diffusely reflecting cool walls is derived analytically and compared against self-consistent electrostatic Particle-in-Cell (PIC) simulations. Analytical calculations predict that neglecting charged drag in an O+ dominated LEO plasma with a neutral to ion number density ratio of 102 will cause a 10% over-prediction of O density based on body accelerations when body potential (ɸB) is ≤ -390 V. Above 900 km altitude in LEO, where H+ becomes the dominant ion species, analytical predictions suggest charge drag becomes equivalent to neutral drag for ɸB ≤ -0.75 V. Comparing analytical predictions against PIC simulations in the range of 0 < - ɸB < 50 V found that analytical charged drag was under-estimated for all body potentials; the degree of under-estimation increasing with ɸB. Based on the -50 V PIC simulations, our in-house 6 degree of freedom orbital propagator saw a reduction in the semi-major axis of a 10 kg satellite at 700 km of 6.9 m/day and 0.98 m/day at 900 km compared that caused purely by neutral drag - 0.67 m/day and 0.056 m/day respectively. Hence, this work provides initial evidence that charged aerodynamics may become significant compared to neutral aerodynamics for high voltage LEO bodies.

Effects of plasma drag on low Earth orbiting satellites due to solar forcing induced perturbations and heating

NASA Astrophysics Data System (ADS)

Nwankwo, Victor U. J.; Chakrabarti, Sandip K.; Weigel, Robert S.

2015-07-01

The upper atmosphere changes significantly in temperature, density and composition as a result of solar cycle variations, which causes severe storms and flares, and increases in the amount of absorbed solar radiation from solar energetic events. Satellite orbits are consequently affected by this process, especially those in low Earth orbit (LEO). In this paper, we present a model of atmospheric drag effects on the trajectory of two hypothetical LEO satellites of different ballistic coefficients, initially injected at h = 450 km. We investigate long-term trends of atmospheric drag on LEO satellites due to solar forcing induced atmospheric perturbations and heating at different phases of the solar cycle, and during short intervals of strong geomagnetic disturbances or magnetic storms. We show dependence of orbital decay on the severity of both solar cycle and phase and the extent of geomagnetic perturbations. The result of the model compares well with observed decay profile of some existing LEO satellites and provide a justification of the theoretical considerations used here.

Simulation platform of LEO satellite communication system based on OPNET

NASA Astrophysics Data System (ADS)

Zhang, Yu; Zhang, Yong; Li, Xiaozhuo; Wang, Chuqiao; Li, Haihao

2018-02-01

For the purpose of verifying communication protocol in the low earth orbit (LEO) satellite communication system, an Optimized Network Engineering Tool (OPNET) based simulation platform is built. Using the three-layer modeling mechanism, the network model, the node model and the process model of the satellite communication system are built respectively from top to bottom, and the protocol will be implemented by finite state machine and Proto-C language. According to satellite orbit parameters, orbit files are generated via Satellite Tool Kit (STK) and imported into OPNET, and the satellite nodes move along their orbits. The simulation platform adopts time-slot-driven mode, divides simulation time into continuous time slots, and allocates slot number for each time slot. A resource allocation strategy is simulated on this platform, and the simulation results such as resource utilization rate, system throughput and packet delay are analyzed, which indicate that this simulation platform has outstanding versatility.

Effect of KOH concentration on LEO cycle life of IPV nickel-hydrogen flight cell - Update II

NASA Technical Reports Server (NTRS)

Smithrick, John J.; Hall, Stephen W.

1992-01-01

An update of validation test results confirming the breakthrough in LEO cycle life of nickel-hydrogen cells containing 26 percent KOH electrolyte is presented. A breakthrough in the LEO cycle life of individual pressure vessel (IPV) nickel-hydrogen cells has been previously reported. The cycle life of boiler plate cells containing 26 percent potassium hydroxide (KOH) electrolyte was about 40,000 LEO cycles, compared to 3500 cycles for cells containing 31 percent KOH. The cycle regime was a stressful accelerated LEO, which consisted of a 27.5 min charge followed by a 17.5 min discharge (2X normal rate). The depth-of-discharge was 80 percent. Six 48-Ah Hughes recirculation design IPV nickel-hydrogen flight battery cells are being evaluated. Three of the cells contain 26 percent KOH (test cells), and three contain 31 percent KOH (control cells). They are undergoing real time LEO cycle life testing. The cycle regime is a 90-min LEO orbit consisting of a 54-min charge followed by a 36-min discharge. The depth-of-discharge is 80 percent. The cell temperature is maintained at 10 C. The three 31 percent KOH cells failed (cycles 3729, 4165, and 11355). One of the 26 percent KOH cells failed at cycle 15314. The other two 26 percent KOH cells were cycled for over 16,000 cycles during the continuing test.

Space-based laser-powered orbital transfer vehicle (Project SLICK)

NASA Technical Reports Server (NTRS)

1988-01-01

The project SLICK (Space Laser Interorbital Cargo Kite) involves conceptual designs of reusable space-based laser-powered orbital transfer vehicle (LOTV) for ferrying 16,000 kg cargo primarily between low Earth orbit (LEO) and geosynchronous earth orbit (GEO). The power of LOTV is beamed by a single 32-MW solar-pumped iodide laser orbiting the Earth at an altitude of one Earth radius. The laser engine selected for the LOTV is based on a continuous-wave, steady-state propulsion scheme and uses an array of seven discrete plasmas in a flow of hydrogen propellant. Both all-propulsive and aerobraked LOTV configurations were analyzed and developed. The all-propulsive vehicle uses a rigid 11.5-m aperture primary mirror and its engine produces a thrust of 2000 N at a specific impulse of 1500 sec. For the LEO-to-GEO trip, the payload ratio, m(sub payload/m(sub propellant)+m(sub dry vehicle) = 1.19 and the trip time is about 6 days. The aerobraked version uses a lightweight, retractable wrapped-rib primary mirror which is folded for aerobraking and a 20-m-diameter inflatable-ballute aeroshield which is jettisoned after aeromaneuver. Lifecycle cost analysis shows that the aerobraked configuration may have an economic advantage over the all-propulsive configuration as long as the cost of launching the propellant to LEO is higher than about $500/kg in current dollars.

Low Earth Orbit Rendezvous Strategy for Lunar Missions

NASA Technical Reports Server (NTRS)

Cates, Grant R.; Cirillo, William M.; Stromgren, Chel

2006-01-01

On January 14, 2004 President George W. Bush announced a new Vision for Space Exploration calling for NASA to return humans to the moon. In 2005 NASA decided to use a Low Earth Orbit (LEO) rendezvous strategy for the lunar missions. A Discrete Event Simulation (DES) based model of this strategy was constructed. Results of the model were then used for subsequent analysis to explore the ramifications of the LEO rendezvous strategy.

Properties of the optimal trajectories for coplanar, aeroassisted orbital transfer

NASA Technical Reports Server (NTRS)

Miele, A.; Wang, T.; Deaton, A. W.

1990-01-01

The optimization of trajectories for coplaner, aeroassisted orbital transfer (AOT) from a high Earth orbit (HEO) to a low Earth orbit (LEO) is examined. In particular, HEO can be a geosynchronous Earth orbit (GEO). It is assumed that the initial and final orbits are circular, that the gravitational field is central and is governed by the inverse square law, and that two impulses are employed, one at HEO exit and one at LEO entry. During the atmospheric pass, the trajectory is controlled via the lift coefficient in such a way that the total characteristic velocity is minimized. First, an ideal optimal trajectory is determined analytically for lift coefficient unbounded. This trajectory is called grazing trajectory, because the atmospheric pass is made by flying at constant altitude along the edge of the atmosphere until the excess velocity is depleted. For the grazing trajectory, the lift coefficient varies in such a way that the lift, the centrifugal force due to the Earth's curvature, the weight, and the Coriolis force due to the Earth's rotation are in static balance. Also, the grazing trajectory minimizes the total characteristic velocity and simultaneously nearly minimizes the peak values of the altitude drop, dynamic pressure, and heating rate. Next, starting from the grazing trajectory results, a real optimal trajectory is determined numerically for the lift coefficient bounded from both below and above. This trajectory is characterized by atmospheric penetration with the smallest possible entry angle, followed by flight at the lift coefficient lower bound. Consistently with the grazing trajectory behavior, the real optimal trajectory minimizes the total characteristic velocity and simultaneously nearly minimizes the peak values of the altitude drop, the dynamic pressure, and the heating rate.

On the Determination of the Orbits of Comets

NASA Astrophysics Data System (ADS)

Englefield, Henry

2013-06-01

Preface; 1. General view of the method; 2. On the motion of the point of intersection of the radius vector and cord; 3. On the comparison of the parabolic cord with the space which answers to the mean velocity of the earth in the same time; 4. Of the reduction of the second longitude of the comet; 5. On the proportion of the three curtate distances of the comet from the earth; 6. Of the graphical declination of the orbit of the earth; 7. Of the numerical quantities to be prepared for the construction or computation of the comet's orbit; 8. Determination of the distances of the comet from the earth and the sun; 9. Determination of the elements of the orbit from the determined distances; 10. Determination of the place of the comet from the earth and sun; 11. Determination of the distances of the comet from the earth and sun; 12. Determination of the comet's orbit; 13. Determination of the place of the comet; 14. Application of the graphical method to the comet of 1769; 15. Application of the distances found; 16. Determination of the place of the comet, for another given time; 17. Application of the trigonometrical method to the comet of 1769; 18. Determination of the elements of the orbit of the comet of 1769; Example of the graphical operation for the orbit of the comet of 1769; Example of the trigonometrical operation for the orbit of the comet of 1769; Conclusion; La Place's general method for determining the orbits of comets; Determination of the two elements of the orbit; Application of La Place's method of finding the approximate perihelion distance; Application of La Place's method for correcting the orbit of a comet, to the comet of 1769; Explanation and use of the tables; Tables; Appendix; Plates.

On-Orbit Collision Hazard Analysis in Low Earth Orbit Using the Poisson Probability Distribution (Version 1.0)

DOT National Transportation Integrated Search

1992-08-26

This document provides the basic information needed to estimate a general : probability of collision in Low Earth Orbit (LEO). Although the method : described in this primer is a first order approximation, its results are : reasonable. Furthermore, t...

Nickel-hydrogen LEO cycling at 20-50 percent DOD. [depth of discharge

NASA Technical Reports Server (NTRS)

Lowery, John E.; Mai, Jenny

1991-01-01

Two NiH2 two-cell packs made up of engineering cells built according to the Hubble Space Telescope design (EPI RNH 90-3) are currently being low-earth-orbit (LEO) cycled at 20-50 percent depth of discharge (DOD). The cells were manufactured by Eagle-Picher Industries, Inc., and activated with electrolyte (KOH) concentrations of 26 percent (pack No.1) and 31 percent (pack No.2), for use during evaluation of the HST cell design. The cells have been grouped according to electrolyte concentration but follow the same test schedule for comparison. This test was set up to study the behavior of NiH2 cells having differing electrolyte concentrations, when operated at relatively high DOD (20-50 percent) in a LEO cycling program. The test was designed specifically to allow the cells to pick their own recharge ratio for varying DOD and varying EOC (end of charge) voltages. The cells are being cycled in a simulated 96-min orbit with 60-min charge and 36-min discharge where an EOC cutoff voltage controls high-rate charging. EOC cutoff voltages vary between 1.48 V and 1.56 V.

Satellite orbit determination using quantum correlation technology

NASA Astrophysics Data System (ADS)

Zhang, Bo; Sun, Fuping; Zhu, Xinhui; Jia, Xiaolin

2018-03-01

After the presentation of second-order correlation ranging principles with quantum entanglement, the concept of quantum measurement is introduced to dynamic satellite precise orbit determination. Based on the application of traditional orbit determination models for correcting the systematic errors within the satellite, corresponding models for quantum orbit determination (QOD) are established. This paper experiments on QOD with the BeiDou Navigation Satellite System (BDS) by first simulating quantum observations of 1 day arc-length. Then the satellite orbits are resolved and compared with the reference precise ephemerides. Subsequently, some related factors influencing the accuracy of QOD are discussed. Furthermore, the accuracy for GEO, IGSO and MEO satellites increase about 20, 30 and 10 times, respectively, compared with the results from the resolution by measured data. Therefore, it can be expected that quantum technology may also bring delightful surprises to satellite orbit determination as have already emerged in other fields.

Determination of GPS orbits to submeter accuracy

NASA Technical Reports Server (NTRS)

Bertiger, W. I.; Lichten, S. M.; Katsigris, E. C.

1988-01-01

Orbits for satellites of the Global Positioning System (GPS) were determined with submeter accuracy. Tests used to assess orbital accuracy include orbit comparisons from independent data sets, orbit prediction, ground baseline determination, and formal errors. One satellite tracked 8 hours each day shows rms error below 1 m even when predicted more than 3 days outside of a 1-week data arc. Differential tracking of the GPS satellites in high Earth orbit provides a powerful relative positioning capability, even when a relatively small continental U.S. fiducial tracking network is used with less than one-third of the full GPS constellation. To demonstrate this capability, baselines of up to 2000 km in North America were also determined with the GPS orbits. The 2000 km baselines show rms daily repeatability of 0.3 to 2 parts in 10 to the 8th power and agree with very long base interferometry (VLBI) solutions at the level of 1.5 parts in 10 to the 8th power. This GPS demonstration provides an opportunity to test different techniques for high-accuracy orbit determination for high Earth orbiters. The best GPS orbit strategies included data arcs of at least 1 week, process noise models for tropospheric fluctuations, estimation of GPS solar pressure coefficients, and combine processing of GPS carrier phase and pseudorange data. For data arc of 2 weeks, constrained process noise models for GPS dynamic parameters significantly improved the situation.

Semi-Major Axis Knowledge and GPS Orbit Determination

NASA Technical Reports Server (NTRS)

Carpenter, J. Russell; Schiesser, Emil R.; Bauer, F. (Technical Monitor)

2000-01-01

In recent years spacecraft designers have increasingly sought to use onboard Global Positioning System receivers for orbit determination. The superb positioning accuracy of GPS has tended to focus more attention on the system's capability to determine the spacecraft's location at a particular epoch than on accurate orbit determination, per se. The determination of orbit plane orientation and orbit shape to acceptable levels is less challenging than the determination of orbital period or semi-major axis. It is necessary to address semi-major axis mission requirements and the GPS receiver capability for orbital maneuver targeting and other operations that require trajectory prediction. Failure to determine semi-major axis accurately can result in a solution that may not be usable for targeting the execution of orbit adjustment and rendezvous maneuvers. Simple formulas, charts, and rules of thumb relating position, velocity, and semi-major axis are useful in design and analysis of GPS receivers for near circular orbit operations, including rendezvous and formation flying missions. Space Shuttle flights of a number of different GPS receivers, including a mix of unfiltered and filtered solution data and Standard and Precise Positioning Service modes, have been accomplished. These results indicate that semi-major axis is often not determined very accurately, due to a poor velocity solution and a lack of proper filtering to provide good radial and speed error correlation.

Semi-Major Axis Knowledge and GPS Orbit Determination

NASA Technical Reports Server (NTRS)

Carpenter, J. Russell; Schiesser, Emil R.; Bauer, F. (Technical Monitor)

2000-01-01

In recent years spacecraft designers have increasingly sought to use onboard Global Positioning System receivers for orbit determination. The superb positioning accuracy of GPS has tended to focus more attention on the system's capability to determine the spacecraft's location at a particular epoch than on accurate orbit determination, per se. The determination of orbit plane orientation and orbit shape to acceptable levels is less challenging than the determination of orbital period or semi-major axis. It is necessary to address semi-major axis mission requirements and the GPS receiver capability for orbital maneuver targeting and other operations that require trajectory prediction. Failure to determine semi-major axis accurately can result in a solution that may not be usable for targeting the execution of orbit adjustment and rendezvous maneuvers. Simple formulas, charts, and rules of thumb relating position, velocity, and semi-major axis are useful in design and analysis of GPS receivers for near circular orbit operations, including rendezvous and formation flying missions. Space Shuttle flights of a number of different GPS receivers, including a mix of unfiltered and filtered solution data and Standard and Precise Positioning, Service modes, have been accomplished. These results indicate that semi-major axis is often not determined very accurately, due to a poor velocity solution and a lack of proper filtering to provide good radial and speed error correlation.

Earth Orbiting Support Systems for commercial low Earth orbit data relay: Assessing architectures through tradespace exploration

NASA Astrophysics Data System (ADS)

Palermo, Gianluca; Golkar, Alessandro; Gaudenzi, Paolo

2015-06-01

As small satellites and Sun Synchronous Earth Observation systems are assuming an increased role in nowadays space activities, including commercial investments, it is of interest to assess how infrastructures could be developed to support the development of such systems and other spacecraft that could benefit from having a data relay service in Low Earth Orbit (LEO), as opposed to traditional Geostationary relays. This paper presents a tradespace exploration study of the architecture of such LEO commercial satellite data relay systems, here defined as Earth Orbiting Support Systems (EOSS). The paper proposes a methodology to formulate architectural decisions for EOSS constellations, and enumerate the corresponding tradespace of feasible architectures. Evaluation metrics are proposed to measure benefits and costs of architectures; lastly, a multicriteria Pareto criterion is used to downselect optimal architectures for subsequent analysis. The methodology is applied to two case studies for a set of 30 and 100 customer-spacecraft respectively, representing potential markets for LEO services in Exploration, Earth Observation, Science, and CubeSats. Pareto analysis shows how increased performance of the constellation is always achieved by an increased node size, as measured by the gain of the communications antenna mounted on EOSS spacecraft. On the other hand, nonlinear trends in optimal orbital altitude, number of satellites per plane, and number of orbital planes, are found in both cases. An upward trend in individual node memory capacity is found, although never exceeding 256 Gbits of onboard memory for both cases that have been considered, assuming the availability of a polar ground station for EOSS data downlink. System architects can use the proposed methodology to identify optimal EOSS constellations for a given service pricing strategy and customer target, thus identifying alternatives for selection by decision makers.

The Velocity Dispersion Profile of the Remote Dwarf Spheroidal Galaxy Leo I: A Tidal Hit and Run?

NASA Astrophysics Data System (ADS)

Mateo, Mario; Olszewski, Edward W.; Walker, Matthew G.

2008-03-01

We present new kinematic results for 387 stars near the Milky Way satellite dwarf spheroidal galaxy Leo I. Spectra were obtained with the Hectochelle multiobject echelle spectrograph on the MMT, centered in the optical near 5200 Å. From 297 repeat measurements of 108 stars, we estimate the mean velocity error (1 σ) of our sample to be 2.4 km s-1, with a systematic precision of <=1 km s-1. The final sample of 328 Leo I members gives a mean heliocentric velocity of 282.9 +/- 0.5 km s-1 and a dispersion of 9.2 +/- 0.4 km s-1. The dispersion profile of Leo I is flat to beyond its classical "tidal" radius. We fit the profile to various equilibrium dynamical models. We strongly rule out all models where mass follows light. Anisotropic Sérsic+NFW models fit the dispersion profile well, but isotropic models are ruled out at a 95% confidence level. Inside a projected radius of ~1040 pc, the mass and V-band mass-to-light ratio of Leo I from equilibrium models are in the ranges (5-7) × 107 M⊙ and 9-14 (solar units), respectively. Leo I members outside a "break radius" of Rb ~ 400'' (500 pc) exhibit significant velocity anisotropy, whereas stars interior to this radius are consistent with an isotropic velocity distribution. We interpret the break radius as the tidal radius of Leo I at perigalacticon some 1-2 Gyr ago. This interpretation accounts for the complex star formation history of Leo I, population segregation within the galaxy, and Leo I's large outward galactocentric velocity. The lack of evident tidal arms in Leo I suggests that the galaxy may have been injected into its present highly elliptical orbit by a third body a few Gyr before its last perigalacticon. This scenario is plausible within current hierarchical structure formation models.

Orbit determination support of the Ocean Topography Experiment (TOPEX)/Poseidon operational orbit

NASA Technical Reports Server (NTRS)

Schanzle, A. F.; Rovnak, J. E.; Bolvin, D. T.; Doll, C. E.

1993-01-01

The Ocean Topography Experiment (TOPEX/Poseidon) mission is designed to determine the topography of the Earth's sea surface over a 3-year period, beginning shortly after launch in July 1992. TOPEX/Poseidon is a joint venture between the United States National Aeronautics and Space Administration (NASA) and the French Centre Nationale d'Etudes Spatiales. The Jet Propulsion Laboratory is NASA's TOPEX/Poseidon project center. The Tracking and Data Relay Satellite System (TDRSS) will nominally be used to support the day-to-day orbit determination aspects of the mission. Due to its extensive experience with TDRSS tracking data, the NASA Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) will receive and process TDRSS observational data. To fulfill the scientific goals of the mission, it is necessary to achieve and maintain a very precise orbit. The most stringent accuracy requirements are associated with planning and evaluating orbit maneuvers, which will place the spacecraft in its mission orbit and maintain the required ground track. To determine if the FDF can meet the TOPEX/Poseidon maneuver accuracy requirements, covariance analysis was undertaken with the Orbit Determination Error Analysis System (ODEAS). The covariance analysis addressed many aspects of TOPEX/Poseidon orbit determination, including arc length, force models, and other processing options. The most recent analysis has focused on determining the size of the geopotential field necessary to meet the maneuver support requirements. Analysis was undertaken with the full 50 x 50 Goddard Earth Model (GEM) T3 field as well as smaller representations of this model.

The K-1 Active Dispenser for Orbit Transfer

NASA Astrophysics Data System (ADS)

Lai, G.; Cochran, D.; Curtis, R.

2002-01-01

Kistler Aerospace Corporation is building the K-1, the world's first fully reusable launch vehicle. The two-stage K- 1 is designed primarily to service the market for low-earth orbit (LEO) missions, due to Kistler's need to recover both stages. For customers requiring payload delivery to high-energy orbits, Kistler can outfit the payload with a K- 1 Active Dispenser (an expendable third stage). The K-1 second stage will deploy the Active Dispenser mated with its payload into a 200 km circular LEO parking orbit. From this orbit, the Active Dispenser would use its own propulsion to place its payload into the final desired drop-off orbit or earth-escape trajectory. This approach allows Kistler to combine the low-cost launch services offered by the reusable two-stage K-1 with the versatility of a restartable, expendable upper stage. Enhanced with an Active Dispenser, the K-1 will be capable of delivering 1,500 kg to a geosynchronous transfer orbit or up to approximately 1,000 kg into a Mars rendezvous trajectory. The list price of a K-1 Active Dispenser launch is 25 million (plus the price of mission unique integration services) significantly less than the price of any launch vehicle service in the world with comparable capability.

Constrained orbital intercept-evasion

NASA Astrophysics Data System (ADS)

Zatezalo, Aleksandar; Stipanovic, Dusan M.; Mehra, Raman K.; Pham, Khanh

2014-06-01

An effective characterization of intercept-evasion confrontations in various space environments and a derivation of corresponding solutions considering a variety of real-world constraints are daunting theoretical and practical challenges. Current and future space-based platforms have to simultaneously operate as components of satellite formations and/or systems and at the same time, have a capability to evade potential collisions with other maneuver constrained space objects. In this article, we formulate and numerically approximate solutions of a Low Earth Orbit (LEO) intercept-maneuver problem in terms of game-theoretic capture-evasion guaranteed strategies. The space intercept-evasion approach is based on Liapunov methodology that has been successfully implemented in a number of air and ground based multi-player multi-goal game/control applications. The corresponding numerical algorithms are derived using computationally efficient and orbital propagator independent methods that are previously developed for Space Situational Awareness (SSA). This game theoretical but at the same time robust and practical approach is demonstrated on a realistic LEO scenario using existing Two Line Element (TLE) sets and Simplified General Perturbation-4 (SGP-4) propagator.

Backgrounds, radiation damage, and spacecraft orbits

NASA Astrophysics Data System (ADS)

Grant, Catherine E.; Miller, Eric D.; Bautz, Mark W.

2017-08-01

The scientific utility of any space-based observatory can be limited by the on-orbit charged particle background and the radiation-induced damage. All existing and proposed missions have had to make choices about orbit selection, trading off the radiation environment against other factors. We present simulations from ESA’s SPace ENVironment Information System (SPENVIS) of the radiation environment for spacecraft in a variety of orbits, from Low Earth Orbit (LEO) at multiple inclinations to High Earth Orbit (HEO) to Earth-Sun L2 orbit. We summarize how different orbits change the charged particle background and the radiation damage to the instrument. We also discuss the limitations of SPENVIS simulations, particularly outside the Earth’s trapped radiation and point to new resources attempting to address those limitations.

A new release of the mean orbital motion theory, and a new tool provided by CNES for long term analysis of disposal orbits and re-entry predictions

NASA Astrophysics Data System (ADS)

Deleflie, Florent; Wailliez, Sébastien; Portmann, Christophe; Gilles, M.; Vienne, Alain; Berthier, J.; Valk, St; Hautesserres, Denis; Martin, Thierry; Fraysse, Hubert

To perform an orbit modelling accurate enough to provide a good estimate of the lifetime of a satellite, or to ensure the stability of a disposal orbit through centuries, we built a new orbit propagator based on the theory of mean orbital motion. It is named SECS-SD2 , for Simplified and Extended CODIOR Software -Space Debris Dedicated . The CODIOR software propagates numerically averaged equations of motion, with a typical integration step size on the order of a few hours, and was originally written in classical orbital elements. The so-called Space Debris -dedicated version is written in orbital elements suitable for orbits with small eccentricities and inclinations, so as to characterize the main dynamic properties of the motion within the LEO, MEO, and GEO regions. The orbital modelling accounts for the very first terms of the geopotential, the perturbations induced by the luni-solar attraction, the solar radiation pressure, and the atmospheric drag (using classical models). The new software was designed so as to ensure short computation times, even over periods of decades or centuries. This paper aims first at describing and validating the main functionalities of the software: we explain how the simplified averaged equations of motion were built, we show how we get sim-plified luni-solar ephemerides without using any huge file for orbit propagations over centuries, and we show how we averaged and simulated the solar flux. We show as well how we expressed short periodic terms to be added to the mean equations of motion, in order to get orbital ele-ments comparable to those deduced from the classical numerical integration of the oscultating equations of motion. The second part of the paper sheds light on some dynamical properties of space debris flying in the LEO and GEO regions, which were obtained from the new software. Knowing that each satellite in the LEO region is now supposed to re-enter the atmosphere within a period of 25 years, we estimated in various

A Parametric Study on Using Active Debris Removal to Stabilize the Future LEO Debris Environment

NASA Technical Reports Server (NTRS)

Liou, J.C.

2010-01-01

Recent analyses of the instability of the orbital debris population in the low Earth orbit (LEO) region and the collision between Iridium 33 and Cosmos 2251 have reignited the interest in using active debris removal (ADR) to remediate the environment. There are; however, monumental technical, resources, operational, legal, and political challenges in making economically viable ADR a reality. Before a consensus on the need for ADR can be reached, a careful analysis of the effectiveness of ADR must be conducted. The goal is to demonstrate the feasibility of using ADR to preserve the future environment and to guide its implementation to maximize the benefit-cost ratio. This paper describes a comprehensive sensitivity study on using ADR to stabilize the future LEO debris environment. The NASA long-term, orbital debris evolutionary model, LEGEND, is used to quantify the effects of many key parameters. These parameters include (1) the starting epoch of ADR implementation, (2) various target selection criteria, (3) the benefits of collision avoidance maneuvers, (4) the consequence of targeting specific inclination or altitude regimes, (5) the consequence of targeting specific classes of vehicles, and (6) the timescale of removal. Additional analyses on the importance of postmission disposal and how future launches might affect the requirements to stabilize the environment are also included.

Numerical Algorithms for Precise and Efficient Orbit Propagation and Positioning

NASA Astrophysics Data System (ADS)

Bradley, Ben K.

orbit propagation, yielding savings in computation time and memory. Orbit propagation and position transformation simulations are analyzed to generate a complete set of recommendations for performing the ITRS/GCRS transformation for a wide range of needs, encompassing real-time on-board satellite operations and precise post-processing applications. In addition, a complete derivation of the ITRS/GCRS frame transformation time-derivative is detailed for use in velocity transformations between the GCRS and ITRS and is applied to orbit propagation in the rotating ITRS. EOP interpolation methods and ocean tide corrections are shown to impact the ITRS/GCRS transformation accuracy at the level of 5 cm and 20 cm on the surface of the Earth and at the Global Positioning System (GPS) altitude, respectively. The precession-nutation and EOP simplifications yield maximum propagation errors of approximately 2 cm and 1 m after 15 minutes and 6 hours in low-Earth orbit (LEO), respectively, while reducing computation time and memory usage. Finally, for orbit propagation in the ITRS, a simplified scheme is demonstrated that yields propagation errors under 5 cm after 15 minutes in LEO. This approach is beneficial for orbit determination based on GPS measurements. We conclude with a summary of recommendations on EOP usage and bias-precession-nutation implementations for achieving a wide range of transformation and propagation accuracies at several altitudes. This comprehensive set of recommendations allows satellite operators, astrodynamicists, and scientists to make informed decisions when choosing the best implementation for their application, balancing accuracy and computational complexity.

Servicing communication satellites in geostationary orbit

NASA Technical Reports Server (NTRS)

Russell, Paul K.; Price, Kent M.

1990-01-01

The econmic benefits of a LEO space station are quantified by identifying alternative operating scenarios utilizing the space station's transportation facilities and assembly and repair facilities. Particular consideration is given to the analysis of the impact of on-orbit assembly and servicing on a typical communications satellite is analyzed. The results of this study show that on-orbit servicing can increase the internal rate of return by as much as 30 percent.

Analytical model for orbital debris environmental management

NASA Technical Reports Server (NTRS)

Talent, David L.

1990-01-01

A differential equation, also referred to as the PIB (particle-in-a-box) model, expressing the time rate of change of the number of objects in orbit, is developed, and its applicability is illustrated. The model can be used as a tool for the assessment of LEO environment stability, and as a starting point for the development of numerical evolutionary models. Within the context of the model, evolutionary scenarios are examined, and found to be sensitive to the growth rate. It is determined that the present environment is slightly unstable to catastrophic growth, and that the number of particles on orbit will continue to increase until approximately 2250-2350 AD, with a maximum of 2,000,000. The model is expandable to the more realistic (complex) case of multiple species in a multiple-tier system.

An Artificial Gravity Spacecraft Approach which Minimizes Mass, Fuel and Orbital Assembly Reg

NASA Astrophysics Data System (ADS)

Bell, L.

2002-01-01

The Sasakawa International Center for Space Architecture (SICSA) is undertaking a multi-year research and design study that is exploring near and long-term commercial space development opportunities. Space tourism in low-Earth orbit (LEO), and possibly beyond LEO, comprises one business element of this plan. Supported by a financial gift from the owner of a national U.S. hotel chain, SICSA has examined opportunities, requirements and facility concepts to accommodate up to 100 private citizens and crewmembers in LEO, as well as on lunar/planetary rendezvous voyages. SICSA's artificial gravity Science Excursion Vehicle ("AGSEV") design which is featured in this presentation was conceived as an option for consideration to enable round-trip travel to Moon and Mars orbits and back from LEO. During the course of its development, the AGSEV would also serve other important purposes. An early assembly stage would provide an orbital science and technology testbed for artificial gravity demonstration experiments. An ultimate mature stage application would carry crews of up to 12 people on Mars rendezvous missions, consuming approximately the same propellant mass required for lunar excursions. Since artificial gravity spacecraft that rotate to create centripetal accelerations must have long spin radii to limit adverse effects of Coriolis forces upon inhabitants, SICSA's AGSEV design embodies a unique tethered body concept which is highly efficient in terms of structural mass and on-orbit assembly requirements. The design also incorporates "inflatable" as well as "hard" habitat modules to optimize internal volume/mass relationships. Other important considerations and features include: maximizing safety through element and system redundancy; means to avoid destabilizing mass imbalances throughout all construction and operational stages; optimizing ease of on-orbit servicing between missions; and maximizing comfort and performance through careful attention to human needs. A

Evaluation of Gravitational Field Models Based on the Laser Range Observation of Low Earth Orbit Satellites

NASA Astrophysics Data System (ADS)

Hong-bo, Wang; Chang-yin, Zhao; Wei, Zhang; Jin-wei, Zhan; Sheng-xian, Yu

2016-07-01

The Earth gravitational field model is one of the most important dynamic models in satellite orbit computation. Several space gravity missions made great successes in recent years, prompting the publishing of several gravitational filed models. In this paper, two classical (JGM3, EGM96) and four latest (EIGEN-CHAMP05S, GGM03S, GOCE02S, EGM2008) models are evaluated by employing them in the precision orbit determination (POD) and prediction. These calculations are performed based on the laser ranging observation of four Low Earth Orbit (LEO) satellites, including CHAMP, GFZ-1, GRACE-A, and SWARM-A. The residual error of observation in POD is adopted to describe the accuracy of six gravitational field models. The main results we obtained are as follows. (1) For the POD of LEOs, the accuracies of 4 latest models are at the same level, and better than those of 2 classical models; (2) Taking JGM3 as reference, EGM96 model's accuracy is better in most situations, and the accuracies of the 4 latest models are improved by 12%-47% in POD and 63% in prediction, respectively. We also confirm that the model's accuracy in POD is enhanced with the increasing degree and order if they are smaller than 70, and when they exceed 70, the accuracy keeps constant, implying that the model's degree and order truncated to 70 are sufficient to meet the requirement of LEO computation of centimeter precision.

Mitochondrial genome of the African lion Panthera leo leo.

PubMed

Ma, Yue-ping; Wang, Shuo

2015-01-01

In this study, the complete mitochondrial genome sequence of the African lion P. leo leo was reported. The total length of the mitogenome was 17,054 bp. It contained the typical mitochondrial structure, including 13 protein-coding genes, 22 transfer RNA genes, 2 ribosomal RNA genes and 1 control region; 21 of the tRNA genes folded into typical cloverleaf secondary structure except for tRNASe. The overall composition of the mitogenome was A (32.0%), G (14.5%), C (26.5%) and T (27.0%). The new sequence will provide molecular genetic information for conservation genetics study of this important large carnivore.

Multi-technique combination of space geodesy observations: Impact of the Jason-2 satellite on the GPS satellite orbits estimation

NASA Astrophysics Data System (ADS)

Zoulida, Myriam; Pollet, Arnaud; Coulot, David; Perosanz, Félix; Loyer, Sylvain; Biancale, Richard; Rebischung, Paul

2016-10-01

In order to improve the Precise Orbit Determination (POD) of the GPS constellation and the Jason-2 Low Earth Orbiter (LEO), we carry out a simultaneous estimation of GPS satellite orbits along with Jason-2 orbits, using GINS software. Along with GPS station observations, we use Jason-2 GPS, SLR and DORIS observations, over a data span of 6 months (28/05/2011-03/12/2011). We use the Geophysical Data Records-D (GDR-D) orbit estimation standards for the Jason-2 satellite. A GPS-only solution is computed as well, where only the GPS station observations are used. It appears that adding the LEO GPS observations results in an increase of about 0.7% of ambiguities fixed, with respect to the GPS-only solution. The resulting GPS orbits from both solutions are of equivalent quality, agreeing with each other at about 7 mm on Root Mean Square (RMS). Comparisons of the resulting GPS orbits to the International GNSS Service (IGS) final orbits show the same level of agreement for both the GPS-only orbits, at 1.38 cm in RMS, and the GPS + Jason2 orbits at 1.33 cm in RMS. We also compare the resulting Jason-2 orbits with the 3-technique Segment Sol multi-missions d'ALTimétrie, d'orbitographie et de localisation précise (SSALTO) POD products. The orbits show good agreement, with 2.02 cm of orbit differences global RMS, and 0.98 cm of orbit differences RMS on the radial component.

Orbital Propagation of Momentum Exchange Tether Systems

NASA Technical Reports Server (NTRS)

Westerhoff, John

2002-01-01

An advanced concept in in-space transportation currently being studied is the Momentum-Exchange/Electrodynamic Reboost Tether System (MXER). The system acts as a large momentum wheel, imparting a Av to a payload in low earth orbit (LEO) at the expense of its own orbital energy. After throwing a payload, the system reboosts itself using an electrodynamic tether to push against Earth's magnetic field and brings itself back up to an operational orbit to prepare for the next payload. The ability to reboost itself allows for continued reuse of the system without the expenditure of propellants. Considering the cost of lifting propellant from the ,ground to LEO to do the same Av boost at $10000 per pound, the system cuts the launch cost of the payload dramatically, and subsequently, the MXER system pays for itself after a small number of missions.1 One of the technical hurdles to be overcome with the MXER concept is the rendezvous maneuver. The rendezvous window for the capture of the payload is on the order of a few seconds, as opposed to traditional docking maneuvers, which can take as long ets necessary to complete a precise docking. The payload, therefore, must be able to match its orbit to meet up with the capture device on the end of the tether at a specific time and location in the future. In order to be able to determine that location, the MXER system must be numerically propagated forward in time to predict where the capture device will be at that instant. It should be kept in mind that the propagation computation must be done faster than real-time. This study focuses on the efforts to find and/or build the tools necessary to numerically propagate the motion of the MXER system as accurately as possible.

Effects of Low Earth Orbit on Docking Seal Materials

NASA Technical Reports Server (NTRS)

Imka, Emily C.; Asmar, Olivia C.; deGroh, Henry C., III; Banks, Bruce A.

2014-01-01

Spacecraft docking seals are typically made of silicone elastomers. When such seals are exposed to low Earth orbit (LEO) conditions, they can suffer damage from ultraviolet (UV) radiation and atomic oxygen (AO, or monoatomic oxygen, the predominant oxygen species in LEO). An experiment flew on the International Space Station (ISS) to measure the effects of LEO on seal materials S0383-70 and ELA-SA-401 and various mating counterface materials which included anodized aluminum. Samples flown in different orientations received different amounts of UV and AO. The hypotheses were that most of the damage would be from UV, and 10 days or more of exposure in LEO would badly damage the seals. Eighteen seals were exposed for 543 days in ram (windward), zenith (away from Earth), or wake (leeward) orientations, and 15 control samples (not flown) provided undamaged baseline leakage. To determine post-flight leak rates, each of the 33 seals were placed in an O-ring groove of a leak test fixture and pressure tested over time. Resistance temperature detectors (RTDs), pressure transducers, and LabVIEW (National Instruments) programs were used to measure and analyze the temperature and pressure and calculate leakage. Average leakage of control samples was 2.6 x 10(exp -7) lbs/day. LEO exposure did not considerably damage ELA-SA-401. The S0383-70 flight samples leaked at least 10 times more than ELA-SA-401 in all cases except one, demonstrating that ELA-SA-401 may be a more suitable sealing material in LEO. AO caused greater damage than UV; samples in ram orientation (receiving an AO fluence of 4.3 x 10(exp 21) atoms/(sq cm) and in wake (2.9x 10(exp 20) atoms/(sq cm)) leaked more than those in zenith orientation (1.58 x 10(exp 20) atoms/(sq cm)), whereas variations in UV exposure did not seem to affect the samples. Exposure to LEO did less damage to the seals than hypothesized, and the data did not support the conjecture that UV causes more damage than AO.

Evaluation of Gravitational Field Models Based on the Laser Range Observation of Low Earth Orbit Satellites

NASA Astrophysics Data System (ADS)

Wang, H. B.; Zhao, C. Y.; Zhang, W.; Zhan, J. W.; Yu, S. X.

2015-09-01

The Earth gravitational filed model is a kind of important dynamic model in satellite orbit computation. In recent years, several space gravity missions have obtained great success, prompting a lot of gravitational filed models to be published. In this paper, 2 classical models (JGM3, EGM96) and 4 latest models, including EIGEN-CHAMP05S, GGM03S, GOCE02S, and EGM2008 are evaluated by being employed in the precision orbit determination (POD) and prediction, based on the laser range observation of four low earth orbit (LEO) satellites, including CHAMP, GFZ-1, GRACE-A, and SWARM-A. The residual error of observation in POD is adopted to describe the accuracy of six gravitational field models. We show the main results as follows: (1) for LEO POD, the accuracies of 4 latest models (EIGEN-CHAMP05S, GGM03S, GOCE02S, and EGM2008) are at the same level, and better than those of 2 classical models (JGM3, EGM96); (2) If taking JGM3 as reference, EGM96 model's accuracy is better in most situations, and the accuracies of the 4 latest models are improved by 12%-47% in POD and 63% in prediction, respectively. We also confirm that the model's accuracy in POD is enhanced with the increasing degree and order if they are smaller than 70, and when they exceed 70 the accuracy keeps stable, and is unrelated with the increasing degree, meaning that the model's degree and order truncated to 70 are sufficient to meet the requirement of LEO orbit computation with centimeter level precision.

Diagnosing low earth orbit satellite anomalies using NOAA-15 electron data associated with geomagnetic perturbations

NASA Astrophysics Data System (ADS)

Ahmad, Nizam; Herdiwijaya, Dhani; Djamaluddin, Thomas; Usui, Hideyuki; Miyake, Yohei

2018-05-01

A satellite placed in space is constantly affected by the space environment, resulting in various impacts from temporary faults to permanent failures depending on factors such as satellite orbit, solar and geomagnetic activities, satellite local time, and satellite construction material. Anomaly events commonly occur during periods of high geomagnetic activity that also trigger plasma variation in the low Earth orbit (LEO) environment. In this study, we diagnosed anomalies in LEO satellites using electron data from the Medium Energy Proton and Electron Detector onboard the National Oceanic and Atmospheric Administration (NOAA)-15 satellite. In addition, we analyzed the fluctuation of electron flux in association with geomagnetic disturbances 3 days before and after the anomaly day. We selected 20 LEO anomaly cases registered in the Satellite News Digest database for the years 2000-2008. Satellite local time, an important parameter for anomaly diagnosis, was determined using propagated two-line element data in the SGP4 simplified general perturbation model to calculate the longitude of the ascending node of the satellite through the position and velocity vectors. The results showed that the majority of LEO satellite anomalies are linked to low-energy electron fluxes of 30-100 keV and magnetic perturbations that had a higher correlation coefficient ( 90%) on the day of the anomaly. The mean local time calculation for the anomaly day with respect to the nighttime migration of energetic electrons revealed that the majority of anomalies (65%) occurred on the night side of Earth during the dusk-to-dawn sector of magnetic local time.

Orbit Determination of Spacecraft in Earth-Moon L1 and L2 Libration Point Orbits

NASA Technical Reports Server (NTRS)

Woodard, Mark; Cosgrove, Daniel; Morinelli, Patrick; Marchese, Jeff; Owens, Brandon; Folta, David

2011-01-01

The ARTEMIS mission, part of the THEMIS extended mission, is the first to fly spacecraft in the Earth-Moon Lissajous regions. In 2009, two of the five THEMIS spacecraft were redeployed from Earth-centered orbits to arrive in Earth-Moon Lissajous orbits in late 2010. Starting in August 2010, the ARTEMIS P1 spacecraft executed numerous stationkeeping maneuvers, initially maintaining a lunar L2 Lissajous orbit before transitioning into a lunar L1 orbit. The ARTEMIS P2 spacecraft entered a L1 Lissajous orbit in October 2010. In April 2011, both ARTEMIS spacecraft will suspend Lissajous stationkeeping and will be maneuvered into lunar orbits. The success of the ARTEMIS mission has allowed the science team to gather unprecedented magnetospheric measurements in the lunar Lissajous regions. In order to effectively perform lunar Lissajous stationkeeping maneuvers, the ARTEMIS operations team has provided orbit determination solutions with typical accuracies on the order of 0.1 km in position and 0.1 cm/s in velocity. The ARTEMIS team utilizes the Goddard Trajectory Determination System (GTDS), using a batch least squares method, to process range and Doppler tracking measurements from the NASA Deep Space Network (DSN), Berkeley Ground Station (BGS), Merritt Island (MILA) station, and United Space Network (USN). The team has also investigated processing of the same tracking data measurements using the Orbit Determination Tool Kit (ODTK) software, which uses an extended Kalman filter and recursive smoother to estimate the orbit. The orbit determination results from each of these methods will be presented and we will discuss the advantages and disadvantages associated with using each method in the lunar Lissajous regions. Orbit determination accuracy is dependent on both the quality and quantity of tracking measurements, fidelity of the orbit force models, and the estimation techniques used. Prior to Lissajous operations, the team determined the appropriate quantity of tracking

Evaluating the environmental criticality of massive objects in LEO for debris mitigation and remediation

NASA Astrophysics Data System (ADS)

Pardini, Carmen; Anselmo, Luciano

2018-04-01

Approximately 95% of the mass in Earth orbit is currently concentrated in about 6700 intact objects, of which nearly 80% are abandoned and more than 90% cannot be maneuvered. The intact objects abandoned in low Earth orbit (LEO) above 650 km, i.e. with an average residual lifetime of more than 25 years, represent the main potential mass reservoir for the generation of new detrimental orbital debris in case of mutual collisions with the existing debris environment, taking into account that an 800 g impactor may be sufficient, in principle, to shatter a 1000 kg spacecraft or rocket stage. Since the 1980's, several mitigation measures were promoted and agreed at the international level in order to prevent the occurrence of new breakups in space and put under control the accumulation of mass abandoned in orbit, but unfortunately the level of compliance with such guidelines, requirements or standards is still far from satisfactory. Moreover, the appearance on the scene of space activity of new private and government actors from a growing number of countries makes the proper management of the circumterrestrial space a task of increasing complexity, taking also into account the rapid emerging of new potential applications, disrupting technologies and operational approaches quite different from the past. In this rapidly evolving environment, it might be useful to have a simple and flexible instrument for evaluating the potential criticality for the environment of massive objects placed or abandoned in LEO. With this goal, in the last few years, a particular effort was devoted to the development of various "criticality indexes", then applied for evaluating many families of rocket bodies and selected spacecraft. In this paper, with the underlining ambition to be simple, intuitive and relevant, from an environmental point of view, a couple of the most complete indexes were coherently applied in order to assess the potential criticality of the most massive objects abandoned in

LEGEND, a LEO-to-GEO Environment Debris Model

NASA Technical Reports Server (NTRS)

Liou, Jer Chyi; Hall, Doyle T.

2013-01-01

LEGEND (LEO-to-GEO Environment Debris model) is a three-dimensional orbital debris evolutionary model that is capable of simulating the historical and future debris populations in the near-Earth environment. The historical component in LEGEND adopts a deterministic approach to mimic the known historical populations. Launched rocket bodies, spacecraft, and mission-related debris (rings, bolts, etc.) are added to the simulated environment. Known historical breakup events are reproduced, and fragments down to 1 mm in size are created. The LEGEND future projection component adopts a Monte Carlo approach and uses an innovative pair-wise collision probability evaluation algorithm to simulate the future breakups and the growth of the debris populations. This algorithm is based on a new "random sampling in time" approach that preserves characteristics of the traditional approach and captures the rapidly changing nature of the orbital debris environment. LEGEND is a Fortran 90-based numerical simulation program. It operates in a UNIX/Linux environment.

Satellite orbit determination from an airborne platform

NASA Astrophysics Data System (ADS)

Shepard, M. M.; Foshee, J. J.

This paper describes the requirements, approach, and problems associated with autonomous satellite orbit determination from an airborne platform. The ability to perform orbit determination from an airborne platform removes the reliance on ground control facilities. Aircraft orbit determination offers a more robust system in that it is less susceptible to direct attack, sabotage, or nuclear disaster. Ranging on a satellite and the processing of range/range-rate data along with INS inputs to produce a set of orbital parameters to be transmitted to user terminals are discussed. Several algorithms that could be utilized by the user terminal to recover the satellite position/velocity data from the transmitted message are presented. The ability to compress the ephemeris message to a small size while remaining autonomous for a long period of time, as would be needed in future military communication satellites, is discussed.

Atomic oxygen interaction with spacecraft materials: Relationship between orbital and ground-based testing for materials certification

NASA Technical Reports Server (NTRS)

Cross, Jon B.; Koontz, Steven L.; Lan, Esther H.

1993-01-01

The effects of atomic oxygen on boron nitride (BN), silicon nitride (Si3N4), Intelsat 6 solar cell interconnects, organic polymers, and MoS2 and WS2 dry lubricant, were studied in Low Earth Orbit (LEO) flight experiments and in a ground based simulation facility. Both the inflight and ground based experiments employed in situ electrical resistance measurements to detect penetration of atomic oxygen through materials and Electron Spectroscopy for Chemical Analysis (ESCA) analysis to measure chemical composition changes. Results are given. The ground based results on the materials studied to date show good qualitative correlation with the LEO flight results, thus validating the simulation fidelity of the ground based facility in terms of reproducing LEO flight results. In addition it was demonstrated that ground based simulation is capable of performing more detailed experiments than orbital exposures can presently perform. This allows the development of a fundamental understanding of the mechanisms involved in the LEO environment degradation of materials.

Cost-effective and robust mitigation of space debris in low earth orbit

NASA Astrophysics Data System (ADS)

Walker, R.; Martin, C.

It is predicted that the space debris population in low Earth orbit (LEO) will continue to grow and in an exponential manner in the long-term due to an increasing rate of collisions between large objects, unless internationally-accepted space debris mitigation measures are adopted soon. Such measures are aimed at avoiding the future generation of space debris objects and primarily need to be effective in preventing significant long-term growth in the debris population, even in the potential scenario of an increase in future space activity. It is also important that mitigation measures can limit future debris population levels, and therefore the underlying collision risk to space missions, to the lowest extent possible. However, for their wide acceptance, the cost of implementation associated with mitigation measures needs to be minimised as far as possible. Generally, a lower collision risk will cost more to achieve and vice versa, so it is necessary to strike a balance between cost and risk in order to find a cost-effective set of mitigation measures. In this paper, clear criteria are established in order to assess the cost-effectiveness of space debris mitigation measures. A full cost-risk-benefit trade-off analysis of numerous mitigation scenarios is presented. These scenarios consider explosion prevention and post-mission disposal of space systems, including de-orbiting to limited lifetime orbits and re-orbiting above the LEO region. The ESA DELTA model is used to provide long-term debris environment projections for these scenarios as input to the benefit and risk parts of the trade-off analysis. Manoeuvre requirements for the different post-mission disposal scenarios were also calculated in order to define the cost-related element. A 25-year post-mission lifetime de-orbit policy, combined with explosion prevention and mission-related object limitation, was found to be the most cost-effective solution to the space debris problem in LEO. This package would also

Orbit Determination and Navigation Software Testing for the Mars Reconnaissance Orbiter

NASA Technical Reports Server (NTRS)

Pini, Alex

2011-01-01

During the extended science phase of the Mars Reconnaissance Orbiter's lifecycle, the operational duties pertaining to navigation primarily involve orbit determination. The orbit determination process utilizes radiometric tracking data and is used for the prediction and reconstruction of MRO's trajectories. Predictions are done twice per week for ephemeris updates on-board the spacecraft and for planning purposes. Orbit Trim Maneuvers (OTM-s) are also designed using the predicted trajectory. Reconstructions, which incorporate a batch estimator, provide precise information about the spacecraft state to be synchronized with scientific measurements. These tasks were conducted regularly to validate the results obtained by the MRO Navigation Team. Additionally, the team is in the process of converting to newer versions of the navigation software and operating system. The capability to model multiple densities in the Martian atmosphere is also being implemented. However, testing outputs among these different configurations was necessary to ensure compliance to a satisfactory degree.

Active and Passive Sensing from Geosynchronous and Libration Orbits

NASA Technical Reports Server (NTRS)

Schoeberl, Mark; Raymond, Carol; Hildebrand, Peter

2003-01-01

The development of the LEO (EOS) missions has led the way to new technologies and new science discoveries. However, LEO measurements alone cannot cost effectively produce high time resolution measurements needed to move the science to the next level. Both GEO and the Lagrange points, L1 and L2, provide vantage points that will allow higher time resolution measurements. GEO is currently being exploited by weather satellites, but the sensors currently operating at GEO do not provide the spatial or spectral resolution needed for atmospheric trace gas, ocean or land surface measurements. It is also may be possible to place active sensors in geostationary orbit. It seems clear, that the next era in earth observation and discovery will be opened by sensor systems operating beyond near earth orbit.

Precise Tracking of the Magellan and Pioneer Venus Orbiters by Same-Beam Interferometry. Part 2: Orbit Determination Analysis

NASA Technical Reports Server (NTRS)

Folkner, W. M.; Border, J. S.; Nandi, S.; Zukor, K. S.

1993-01-01

A new radio metric positioning technique has demonstrated improved orbit determination accuracy for the Magellan and Pioneer Venus Orbiter orbiters. The new technique, known as Same-Beam Interferometry (SBI), is applicable to the positioning of multiple planetary rovers, landers, and orbiters which may simultaneously be observed in the same beamwidth of Earth-based radio antennas. Measurements of carrier phase are differenced between spacecraft and between receiving stations to determine the plane-of-sky components of the separation vector(s) between the spacecraft. The SBI measurements complement the information contained in line-of-sight Doppler measurements, leading to improved orbit determination accuracy. Orbit determination solutions have been obtained for a number of 48-hour data arcs using combinations of Doppler, differenced-Doppler, and SBI data acquired in the spring of 1991. Orbit determination accuracy is assessed by comparing orbit solutions from adjacent data arcs. The orbit solution differences are shown to agree with expected orbit determination uncertainties. The results from this demonstration show that the orbit determination accuracy for Magellan obtained by using Doppler plus SBI data is better than the accuracy achieved using Doppler plus differenced-Doppler by a factor of four and better than the accuracy achieved using only Doppler by a factor of eighteen. The orbit determination accuracy for Pioneer Venus Orbiter using Doppler plus SBI data is better than the accuracy using only Doppler data by 30 percent.

Material exposure effects in a simulated low-Earth orbit environment

NASA Astrophysics Data System (ADS)

Maldonado, C.; McHarg, G.; Asmolova, O.; Andersen, G.; Rodrigues, S.; Ketsdever, A.

2016-11-01

Spacecraft operating in low-Earth orbit (LEO) are subjected to a number of hazardous environmental constituents that can lead to decreased system performance and reduced operational lifetimes. Due to their thermal, optical, and mechanical properties, polymers are used extensively in space systems; however they are particularly susceptible to material erosion and degradation as a result of exposure to the LEO environment. The focus of this research is to examine the material erosion and mass loss experienced by the Novastrat 500 polyimide due to exposure in a simulated LEO environment. In addition to the polymer samples, chrome, silver and gold specimens will be examined to measure the oxidation rate and act as a control specimen, respectively. A magnetically filtered atomic oxygen plasma source has previously been developed and characterized for the purpose of simulating the low-Earth orbit environment. The plasma source can be operated at a variety of discharge currents and gas flow rates, of which the plasma parameters downstream of the source are dependent. The characteristics of the generated plasma were examined as a function of these operating parameters to optimize the production of O+ ions with energy relevant to LEO applications, where the ram energy of the ions due to the motion of the satellite relative to the LEO plasma is high (e.g. 7800 m/s, which corresponds to approximately 5 eV of kinetic energy for O+ ions). The plasma downstream of the source consists of streaming ions with energy of approximately 5 eV and an ion species fraction that is approximately 90% O+.

Environmental Durability Issues for Solar Power Systems in Low Earth Orbit

NASA Technical Reports Server (NTRS)

Degroh, Kim K.; Banks, Bruce A.; Smith, Daniela C.

1994-01-01

Space solar power systems for use in the low Earth orbit (LEO) environment experience a variety of harsh environmental conditions. Materials used for solar power generation in LEO need to be durable to environmental threats such as atomic oxygen, ultraviolet (UV) radiation, thermal cycling, and micrometeoroid and debris impact. Another threat to LEO solar power performance is due to contamination from other spacecraft components. This paper gives an overview of these LEO environmental issues as they relate to space solar power system materials. Issues addressed include atomic oxygen erosion of organic materials, atomic oxygen undercutting of protective coatings, UV darkening of ceramics, UV embrittlement of Teflon, effects of thermal cycling on organic composites, and contamination due to silicone and organic materials. Specific examples of samples from the Long Duration Exposure Facility (LDEF) and materials returned from the first servicing mission of the Hubble Space Telescope (HST) are presented. Issues concerning ground laboratory facilities which simulate the LEO environment are discussed along with ground-to-space correlation issues.

High intensity 5 eV atomic oxygen source and Low Earth Orbit (LEO) simulation facility

NASA Technical Reports Server (NTRS)

Cross, J. B.; Spangler, L. H.; Hoffbauer, M. A.; Archuleta, F. A.; Leger, Lubert; Visentine, James

1987-01-01

An atomic oxygen exposure facility has been developed for studies of material degradation. The goal of these studies is to provide design criteria and information for the manufacture of long life (20 to 30 years) construction materials for use in LEO. The studies that are being undertaken using the facility will provide: absolute reaction cross sections for use in engineering design problems; formulations of reaction mechanisms; and calibration of flight hardware (mass spectrometers, etc.) in order to directly relate experiments performed in LEO to ground based investigations. The facility consists of: (1) a CW laser sustained discharge source of O atoms having a variable energy up to 5 eV and an intensity between 10(15) and 10(17) O atoms s(-1) cm(-2); (2) an atomic beam formation and diagnostics system consisting of various stages of differential pumping, a mass spectrometer detector, and a time of flight analyzer; (3) a spinning rotor viscometer for absolute O atom flux measurements; and (4) provision for using the system for calibration of actual flight instruments. Surface analysis equipment is available for the characterization of material surfaces before and after exposure to O atoms.

Orbit Determination Issues for Libration Point Orbits

NASA Technical Reports Server (NTRS)

Beckman, Mark; Bauer, Frank (Technical Monitor)

2002-01-01

Libration point mission designers require knowledge of orbital accuracy for a variety of analyses including station keeping control strategies, transfer trajectory design, and formation and constellation control. Past publications have detailed orbit determination (OD) results from individual libration point missions. This paper collects both published and unpublished results from four previous libration point missions (ISEE (International Sun-Earth Explorer) -3, SOHO (Solar and Heliospheric Observatory), ACE (Advanced Composition Explorer) and MAP (Microwave Anisotropy Probe)) supported by Goddard Space Flight Center's Guidance, Navigation & Control Center. The results of those missions are presented along with OD issues specific to each mission. All past missions have been limited to ground based tracking through NASA ground sites using standard range and Doppler measurement types. Advanced technology is enabling other OD options including onboard navigation using seaboard attitude sensors and the use of the Very Long Baseline Interferometry (VLBI) measurement Delta Differenced One-Way Range (DDOR). Both options potentially enable missions to reduce coherent dedicated tracking passes while maintaining orbital accuracy. With the increased projected loading of the DSN (Deep Space Network), missions must find alternatives to the standard OD scenario.

Preliminary Experimental Results for Charge Drag in a Simulated Low Earth Orbit Environment

NASA Astrophysics Data System (ADS)

Azema-Rovira, Monica

Interest in the Low Earth Orbit (LEO) environment is growing in the science community as well as in the private sector. The number of spacecraft launched in these altitudes (150 - 700 km) keeps growing, and this region is accumulating space debris. In this scenario, the precise location of all LEO objects is a key factor to avoid catastrophic collisions and to safely perform station-keeping maneuvers. The detailed study of the atmospheric models in LEO can enhance the disturbances forces calculation of an orbiting object. Recent numerical studies indicate that one of the biggest non-conservative forces on a spacecraft is underestimated, the charge drag phenomenon. Validating these numerical models experimentally, will help to improve the numerical models for future spacecraft mission design. For this reason, the motivation of this thesis is to characterize a plasma source to later be used for charged drag measurements. The characterization has been done at the University of Colorado Colorado Springs in the Chamber for Atmospheric and Orbital Space Simulation. In the characterization process, a nano-Newton Thrust Stand has been characterized as a plasma diagnosis tool and compared with Langmuir Probe data.

Precise orbit determination of the Lunar Reconnaissance Orbiter and first gravity field results

NASA Astrophysics Data System (ADS)

Maier, Andrea; Baur, Oliver

2014-05-01

The Lunar Reconnaissance Orbiter (LRO) was launched in 2009 and is expected to orbit the Moon until the end of 2014. Among other instruments, LRO has a highly precise altimeter on board demanding an orbit accuracy of one meter in the radial component. Precise orbit determination (POD) is achieved with radiometric observations (Doppler range rates, ranges) on the one hand, and optical laser ranges on the other hand. LRO is the first satellite at a distance of approximately 360 000 to 400 000 km from the Earth that is routinely tracked with optical laser ranges. This measurement type was introduced to achieve orbits of higher precision than it would be possible with radiometric observations only. In this contribution we investigate the strength of each measurement type (radiometric range rates, radiometric ranges, optical laser ranges) based on single-technique orbit estimation. In a next step all measurement types are combined in a joined analysis. In addition to POD results, preliminary gravity field coefficients are presented being a subsequent product of the orbit determination process. POD and gravity field estimation was accomplished with the NASA/GSFC software packages GEODYN and SOLVE.

The Komplast Experiment: Space Environmental Effects after 12 Years in LEO (and Counting)

NASA Technical Reports Server (NTRS)

Shaevich, S. K.; Aleksandrov, N. G.; Shumov, A. E.; Novikov, L. S.; Alred, J. A.; Shindo, D. J.; Kravchenko, M.; Golden, J. L.

2013-01-01

The Komplast materials experiment was designed by the Khrunichev State Research and Production Space Center, together with other Russian scientific institutes, and has been carried out by Mission Control Moscow since 18. Komplast panels fitted with material samples and sensors were located on the International Space Station (ISS) Functional Cargo Block (FGB) module exterior surface. Within the framework of this experiment, the purpose was to study the effect of the low earth orbit (LEO) environment on exposed samples of various materials. The panels were sent into orbit with the FGB when it launched on November 20, 1998. Panels #2 and #10 were retrieved during Russian extravehicular activity in February 2011 and sealed within cases to temporarily protect the samples from exposure to air until they could be studied on the ground. Panel #2 contained an experiment to detect micrometeoroid and orbital debris (MMOD) impacts, radiation and UV sensors, several pieces of electrical cable, and samples made from elastomeric and fluoroplastic materials. Panel #10 contained a temperature sensor, and both carbon composite and adhesive-bonded samples. A figure shows the location of panels #2 and #10 on the FGB module aft endcone. The panels were subsequently returned to Earth by Space Shuttle Discovery on the STS-133/ULF-5 mission after 12 years of LEO exposure and opened in an argon chamber at the Institute of Nuclear Physics at Moscow State University in July 2011. Based on the results of analyzing the readings from sensors located on Komplast panels and in studying material samples from the panels, the comprehensive effect of spaceflight factors on the FGB (at the locations of Panels #2 and #10) was evaluated. Total solar exposure was determined to be 960 +/- 200 kJ/square cm or 21,000 equivalent solar hours. Because of location of these two panels and the ISS flight attitude, atomic oxygen (AO) fluence was relatively low for such a long duration exposure, approximately 1.5x

Spacecraft Orbit Anomaly Representation Using Thrust-Fourier-Coefficients with Orbit Determination Toolbox

NASA Astrophysics Data System (ADS)

Ko, H.; Scheeres, D.

2014-09-01

Representing spacecraft orbit anomalies between two separate states is a challenging but an important problem in achieving space situational awareness for an active spacecraft. Incorporation of such a capability could play an essential role in analyzing satellite behaviors as well as trajectory estimation of the space object. A general way to deal with the anomaly problem is to add an estimated perturbing acceleration such as dynamic model compensation (DMC) into an orbit determination process based on pre- and post-anomaly tracking data. It is a time-consuming numerical process to find valid coefficients to compensate for unknown dynamics for the anomaly. Even if the orbit determination filter with DMC can crudely estimate an unknown acceleration, this approach does not consider any fundamental element of the unknown dynamics for a given anomaly. In this paper, a new way of representing a spacecraft anomaly using an interpolation technique with the Thrust-Fourier-Coefficients (TFCs) is introduced and several anomaly cases are studied using this interpolation method. It provides a very efficient way of reconstructing the fundamental elements of the dynamics for a given spacecraft anomaly. Any maneuver performed by a satellite transitioning between two arbitrary orbital states can be represented as an equivalent maneuver using an interpolation technique with the TFCs. Given unconnected orbit states between two epochs due to a spacecraft anomaly, it is possible to obtain a unique control law using the TFCs that is able to generate the desired secular behavior for the given orbital changes. This interpolation technique can capture the fundamental elements of combined unmodeled anomaly events. The interpolated orbit trajectory, using the TFCs compensating for a given anomaly, can be used to improve the quality of orbit fits through the anomaly period and therefore help to obtain a good orbit determination solution after the anomaly. Orbit Determination Toolbox (ODTBX

Orbital Debris Detection and Tracking Strategies for the NASA/AFRL Meter Class Autonomous Telescope (MCAT)

NASA Technical Reports Server (NTRS)

Mulrooney, M.; Hickson, P.; Stansbery, Eugene G.

2010-01-01

MCAT (Meter-Class Autonomous Telescope) is a 1.3m f/4 Ritchey-Chr tien on a double horseshoe equatorial mount that will be deployed in early 2011 to the western pacific island of Legan in the Kwajalein Atoll to perform orbital debris observations. MCAT will be capable of tracking earth orbital objects at all inclinations and at altitudes from 200 km to geosynchronous. MCAT s primary objective is the detection of new orbital debris in both low-inclination low-earth orbits (LEO) and at geosynchronous earth orbit (GEO). MCAT was thus designed with a fast focal ratio and a large unvignetted image circle able to accommodate a detector sized to yield a large field of view. The selected primary detector is a close-cycle cooled 4Kx4K 15um pixel CCD camera that yields a 0.9 degree diagonal field. For orbital debris detection in widely spaced angular rate regimes, the camera must offer low read-noise performance over a wide range of framing rates. MCAT s 4-port camera operates from 100 kHz to 1.5 MHz per port at 2 e- and 10 e- read noise respectively. This enables low-noise multi-second exposures for GEO observations as well as rapid (several frames per second) exposures for LEO. GEO observations will be performed using a counter-sidereal time delay integration (TDI) technique which NASA has used successfully in the past. For MCAT the GEO survey, detection, and follow-up prediction algorithms will be automated. These algorithms will be detailed herein. For LEO observations two methods will be employed. The first, Orbit Survey Mode (OSM), will scan specific orbital inclination and altitude regimes, detect new orbital debris objects against trailed background stars, and adjust the telescope track to follow the detected object. The second, Stare and Chase Mode (SCM), will perform a stare, then detect and track objects that enter the field of view which satisfy specific rate and brightness criteria. As with GEO, the LEO operational modes will be fully automated and will be

Orbit Determination Accuracy for Comets on Earth-Impacting Trajectories

NASA Technical Reports Server (NTRS)

Kay-Bunnell, Linda

2004-01-01

The results presented show the level of orbit determination accuracy obtainable for long-period comets discovered approximately one year before collision with Earth. Preliminary orbits are determined from simulated observations using Gauss' method. Additional measurements are incorporated to improve the solution through the use of a Kalman filter, and include non-gravitational perturbations due to outgassing. Comparisons between observatories in several different circular heliocentric orbits show that observatories in orbits with radii less than 1 AU result in increased orbit determination accuracy for short tracking durations due to increased parallax per unit time. However, an observatory at 1 AU will perform similarly if the tracking duration is increased, and accuracy is significantly improved if additional observatories are positioned at the Sun-Earth Lagrange points L3, L4, or L5. A single observatory at 1 AU capable of both optical and range measurements yields the highest orbit determination accuracy in the shortest amount of time when compared to other systems of observatories.

Leo

NASA Astrophysics Data System (ADS)

Murdin, P.

2000-11-01

(the Lion; abbrev. Leo, gen. Leonis; area 947 sq. deg.) A northern zodiacal constellation which lies between Cancer and Virgo, and culminates at midnight in early March. It represents the Nemean lion that, in Greek mythology, Hercules killed as the first of his 12 labors. Its brightest stars were cataloged by Ptolemy (c. AD 100-175) in the Almagest....

GPS-based precision orbit determination - A Topex flight experiment

NASA Technical Reports Server (NTRS)

Melbourne, William G.; Davis, Edgar S.

1988-01-01

Plans for a Topex/Poseiden flight experiment to test the accuracy of using GPS data for precision orbit determination of earth satellites are presented. It is expected that the GPS-based precision orbit determination will provide subdecimeter accuracies in the radial component of the Topex orbit when the extant gravity model is tuned for wavelengths longer than about 1000 kms. The concept, design, flight receiver, antenna system, ground processing, and data processing of GPS are examined. Also, an accurate quasi-geometric orbit determination approach called nondynamic or reduced dynamic tracking which relies on the use of the pseudorange and the carrier phase measurements to reduce orbit errors arising from mismodeled dynamics is discussed.

An Updated Zero Boil-Off Cryogenic Propellant Storage Analysis Applied to Upper Stages or Depots in a LEO Environment

NASA Technical Reports Server (NTRS)

Plachta, David; Kittel, Peter

2003-01-01

Previous efforts have shown the analytical benefits of zero boil-off (ZBO) cryogenic propellant storage in launch vehicle upper stages of Mars transfer vehicles for conceptual Mars Missions. However, recent NASA mission investigations have looked at a different and broad array of missions, including a variety of orbit transfer vehicle (OTV) propulsion concepts, some requiring cryogenic storage. For many of the missions, this vehicle will remain for long periods (greater than one week) in low earth orbit (LEO), a relatively warm thermal environment. Under this environment, and with an array of tank sizes and propellants, the performance of a ZBO cryogenic storage system is predicted and compared with a traditional, passive-only storage concept. The results show mass savings over traditional, passive-only cryogenic storage when mission durations are less than one week in LEO for oxygen, two weeks for methane, and roughly 2 months for LH2. Cryogenic xenon saves mass over passive storage almost immediately.

Orbit Determination for the Lunar Reconnaissance Orbiter Using an Extended Kalman Filter

NASA Technical Reports Server (NTRS)

Slojkowski, Steven; Lowe, Jonathan; Woodburn, James

2015-01-01

Orbit determination (OD) analysis results are presented for the Lunar Reconnaissance Orbiter (LRO) using a commercially available Extended Kalman Filter, Analytical Graphics' Orbit Determination Tool Kit (ODTK). Process noise models for lunar gravity and solar radiation pressure (SRP) are described and OD results employing the models are presented. Definitive accuracy using ODTK meets mission requirements and is better than that achieved using the operational LRO OD tool, the Goddard Trajectory Determination System (GTDS). Results demonstrate that a Vasicek stochastic model produces better estimates of the coefficient of solar radiation pressure than a Gauss-Markov model, and prediction accuracy using a Vasicek model meets mission requirements over the analysis span. Modeling the effect of antenna motion on range-rate tracking considerably improves residuals and filter-smoother consistency. Inclusion of off-axis SRP process noise and generalized process noise improves filter performance for both definitive and predicted accuracy. Definitive accuracy from the smoother is better than achieved using GTDS and is close to that achieved by precision OD methods used to generate definitive science orbits. Use of a multi-plate dynamic spacecraft area model with ODTK's force model plugin capability provides additional improvements in predicted accuracy.

Leo Spacecraft Charging Design Guidelines: A Proposed NASA Standard

NASA Technical Reports Server (NTRS)

Hillard, G. B.; Ferguson, D. C.

2004-01-01

Over the past decade, Low Earth Orbiting (LEO) spacecraft have gradually required ever-increasing power levels. As a rule, this has been accomplished through the use of high voltage systems. Recent failures and anomalies on such spacecraft have been traced to various design practices and materials choices related to the high voltage solar arrays. NASA Glenn has studied these anomalies including plasma chamber testing on arrays similar to those that experienced difficulties on orbit. Many others in the community have been involved in a comprehensive effort to understand the problems and to develop practices to avoid them. The NASA Space Environments and Effects program, recognizing the timeliness of this effort, commissioned and funded a design guidelines document intended to capture the current state of understanding. This document, which was completed in the spring of 2003, has been submitted as a proposed NASA standard. We present here an overview of this document and discuss the effort to develop it as a NASA standard.

Characterization of the cataloged Fengyun-1C fragments and their long-term effect on the LEO environment

NASA Astrophysics Data System (ADS)

Liou, J.-C.; Johnson, N. L.

2009-05-01

The intentional breakup of Fengyun-1C on 11 January 2007 created the most severe orbital debris cloud in history. The altitude where the event occurred was probably the worst location for a major breakup in the low Earth orbit (LEO) region, since it was already highly populated with operational satellites and debris generated from previous breakups. The addition of so many fragments not only poses a realistic threat to operational satellites in the region, but also increases the instability (i.e., collision cascade effect) of the debris population there. Detailed analysis of the large Fengyun-1C fragments indicates that their size and area-to-mass ratio (A/M) distributions are very different from those of other known events. About half of the fragments appear to be composed of light-weight materials and more than 100 of them have A/M values exceeding 1 m 2/kg, consistent with thermal blanket and solar panel pieces. In addition, the orbital elements of the fragments suggest non-trivial velocity gain by the fragment cloud during the impact. These important characteristics were incorporated into numerical simulations to assess the long-term impact of the Fengyun-1C fragments to the LEO debris environment. The collision probabilities between the Fengyun-1C fragments and the rest of the catalog population and the population growth in the low Earth orbit region in the next 100 years are summarized in the paper.

Space Radiation Environment Prediction for VLSI microelectronics devices onboard a LEO Satellite using OMERE-Trad Software

NASA Astrophysics Data System (ADS)

Sajid, Muhammad

This tutorial/survey paper presents the assessment/determination of level of hazard/threat to emerging microelectronics devices in Low Earth Orbit (LEO) space radiation environment with perigee at 300 Km, apogee at 600Km altitude having different orbital inclinations to predict the reliability of onboard Bulk Built-In Current Sensor (BBICS) fabricated in 350nm technology node at OptMA Lab. UFMG Brazil. In this context, the various parameters for space radiation environment have been analyzed to characterize the ionizing radiation environment effects on proposed BBICS. The Space radiation environment has been modeled in the form of particles trapped in Van-Allen radiation belts(RBs), Energetic Solar Particles Events (ESPE) and Galactic Cosmic Rays (GCR) where as its potential effects on Device- Under-Test (DUT) has been predicted in terms of Total Ionizing Dose (TID), Single-Event Effects (SEE) and Displacement Damage Dose (DDD). Finally, the required mitigation techniques including necessary shielding requirements to avoid undesirable effects of radiation environment at device level has been estimated /determined with assumed standard thickness of Aluminum shielding. In order to evaluate space radiation environment and analyze energetic particles effects on BBICS, OMERE toolkit developed by TRAD was utilized.

LEO cooperative multi-spacecraft refueling mission optimization considering J2 perturbation and target's surplus propellant constraint

NASA Astrophysics Data System (ADS)

Zhao, Zhao; Zhang, Jin; Li, Hai-yang; Zhou, Jian-yong

2017-01-01

The optimization of an LEO cooperative multi-spacecraft refueling mission considering the J2 perturbation and target's surplus propellant constraint is studied in the paper. First, a mission scenario is introduced. One service spacecraft and several target spacecraft run on an LEO near-circular orbit, the service spacecraft rendezvouses with some service positions one by one, and target spacecraft transfer to corresponding service positions respectively. Each target spacecraft returns to its original position after obtaining required propellant and the service spacecraft returns to its original position after refueling all target spacecraft. Next, an optimization model of this mission is built. The service sequence, orbital transfer time, and service position are used as deign variables, whereas the propellant cost is used as the design objective. The J2 perturbation, time constraint and the target spacecraft's surplus propellant capability constraint are taken into account. Then, a hybrid two-level optimization approach is presented to solve the formulated mixed integer nonlinear programming (MINLP) problem. A hybrid-encoding genetic algorithm is adopted to seek the near optimal solution in the up-level optimization, while a linear relative dynamic equation considering the J2 perturbation is used to obtain the impulses of orbital transfer in the low-level optimization. Finally, the effectiveness of the proposed model and method is validated by numerical examples.

Precise orbit determination and rapid orbit recovery supported by time synchronization

NASA Astrophysics Data System (ADS)

Guo, Rui; Zhou, JianHua; Hu, XiaoGong; Liu, Li; Tang, Bo; Li, XiaoJie; Wu, Shan

2015-06-01

In order to maintain optimal signal coverage, GNSS satellites have to experience orbital maneuvers. For China's COMPASS system, precise orbit determination (POD) as well as rapid orbit recovery after maneuvers contribute to the overall Positioning, Navigation and Timing (PNT) service performance in terms of accuracy and availability. However, strong statistical correlations between clock offsets and the radial component of a satellite's positions require long data arcs for POD to converge. We propose here a new strategy which relies on time synchronization between ground tracking stations and in-orbit satellites. By fixing satellite clock offsets measured by the satellite station two-way synchronization (SSTS) systems and receiver clock offsets, POD and orbital recovery performance can be improved significantly. Using the Satellite Laser Ranging (SLR) as orbital accuracy evaluation, we find the 4-hr recovered orbit achieves about 0.71 m residual root mean square (RMS) error of fit SLR data, the recovery time is improved from 24-hr to 4-hr compared with the conventional POD without time synchronization support. In addition, SLR evaluation shows that for 1-hr prediction, about 1.47 m accuracy is achieved with the new proposed POD strategy.

Low Earth Orbital Atomic Oxygen Interactions With Materials

NASA Technical Reports Server (NTRS)

Banks, Bruce A.; Miller, Sharon K.; deGroh, Kim K.

2004-01-01

Atomic oxygen is formed in the low Earth orbital environment (LEO) by photo dissociation of diatomic oxygen by short wavelength (< 243 nm) solar radiation which has sufficient energy to break the 5.12 eV O2 diatomic bond in an environment where the mean free path is sufficiently long ( 108 meters) that the probability of reassociation or the formation of ozone (O3) is small. As a consequence, between the altitudes of 180 and 650 km, atomic oxygen is the most abundant species. Spacecraft impact the atomic oxygen resident in LEO with sufficient energy to break hydrocarbon polymer bonds, causing oxidation and thinning of the polymers due to loss of volatile oxidation products. Mitigation techniques, such as the development of materials with improved durability to atomic oxygen attack, as well as atomic oxygen protective coatings, have been employed with varying degrees of success to improve durability of polymers in the LEO environment. Atomic oxygen can also oxidize silicones and silicone contamination to produce non-volatile silica deposits. Such contaminants are present on most LEO missions and can be a threat to performance of optical surfaces. The LEO atomic oxygen environment, its interactions with materials, results of space testing, computational modeling, mitigation techniques, and ground laboratory simulation procedures and issues are presented.

Sentinel-1A - First precise orbit determination results

NASA Astrophysics Data System (ADS)

Peter, H.; Jäggi, A.; Fernández, J.; Escobar, D.; Ayuga, F.; Arnold, D.; Wermuth, M.; Hackel, S.; Otten, M.; Simons, W.; Visser, P.; Hugentobler, U.; Féménias, P.

2017-09-01

Sentinel-1A is the first satellite of the European Copernicus programme. Equipped with a Synthetic Aperture Radar (SAR) instrument the satellite was launched on April 3, 2014. Operational since October 2014 the satellite delivers valuable data for more than two years. The orbit accuracy requirements are given as 5 cm in 3D. In order to fulfill this stringent requirement the precise orbit determination (POD) is based on the dual-frequency GPS observations delivered by an eight-channel GPS receiver. The Copernicus POD (CPOD) Service is in charge of providing the orbital and auxiliary products required by the PDGS (Payload Data Ground Segment). External orbit validation is regularly performed by comparing the CPOD Service orbits to orbit solutions provided by POD expert members of the Copernicus POD Quality Working Group (QWG). The orbit comparisons revealed systematic orbit offsets mainly in radial direction (approx. 3 cm). Although no independent observation technique (e.g. DORIS, SLR) is available to validate the GPS-derived orbit solutions, comparisons between the different antenna phase center variations and different reduced-dynamic orbit determination approaches used in the various software packages helped to detect the cause of the systematic offset. An error in the given geometry information about the satellite has been found. After correction of the geometry the orbit validation shows a significant reduction of the radial offset to below 5 mm. The 5 cm orbit accuracy requirement in 3D is fulfilled according to the results of the orbit comparisons between the different orbit solutions from the QWG.

Orbit determination of the Next-Generation Beidou satellites with Intersatellite link measurements and a priori orbit constraints

NASA Astrophysics Data System (ADS)

Ren, Xia; Yang, Yuanxi; Zhu, Jun; Xu, Tianhe

2017-11-01

Intersatellite Link (ISL) technology helps to realize the auto update of broadcast ephemeris and clock error parameters for Global Navigation Satellite System (GNSS). ISL constitutes an important approach with which to both improve the observation geometry and extend the tracking coverage of China's Beidou Navigation Satellite System (BDS). However, ISL-only orbit determination might lead to the constellation drift, rotation, and even lead to the divergence in orbit determination. Fortunately, predicted orbits with good precision can be used as a priori information with which to constrain the estimated satellite orbit parameters. Therefore, the precision of satellite autonomous orbit determination can be improved by consideration of a priori orbit information, and vice versa. However, the errors of rotation and translation in a priori orbit will remain in the ultimate result. This paper proposes a constrained precise orbit determination (POD) method for a sub-constellation of the new Beidou satellite constellation with only a few ISLs. The observation model of dual one-way measurements eliminating satellite clock errors is presented, and the orbit determination precision is analyzed with different data processing backgrounds. The conclusions are as follows. (1) With ISLs, the estimated parameters are strongly correlated, especially the positions and velocities of satellites. (2) The performance of determined BDS orbits will be improved by the constraints with more precise priori orbits. The POD precision is better than 45 m with a priori orbit constrain of 100 m precision (e.g., predicted orbits by telemetry tracking and control system), and is better than 6 m with precise priori orbit constraints of 10 m precision (e.g., predicted orbits by international GNSS monitoring & Assessment System (iGMAS)). (3) The POD precision is improved by additional ISLs. Constrained by a priori iGMAS orbits, the POD precision with two, three, and four ISLs is better than 6, 3, and 2

A DEEPER LOOK AT LEO IV: STAR FORMATION HISTORY AND EXTENDED STRUCTURE

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sand, David J.; Seth, Anil; Olszewski, Edward W.

We present MMT/Megacam imaging of the Leo IV dwarf galaxy in order to investigate its structure and star formation history, and to search for signs of association with the recently discovered Leo V satellite. Based on parameterized fits, we find that Leo IV is round, with {epsilon} < 0.23 (at the 68% confidence limit) and a half-light radius of r{sub h} {approx_equal} 130 pc. Additionally, we perform a thorough search for extended structures in the plane of the sky and along the line of sight. We derive our surface brightness detection limit by implanting fake structures into our catalog withmore » stellar populations identical to that of Leo IV. We show that we are sensitive to stream-like structures with surface brightness {mu}{sub r} {approx}< 29.6 mag arcsec{sup -2}, and at this limit we find no stellar bridge between Leo IV (out to a radius of {approx}0.5 kpc) and the recently discovered, nearby satellite Leo V. Using the color-magnitude fitting package StarFISH, we determine that Leo IV is consistent with a single age ({approx}14 Gyr), single metallicity ([Fe/H] {approx} -2.3) stellar population, although we cannot rule out a significant spread in these values. We derive a luminosity of M{sub V} = -5.5 {+-} 0.3. Studying both the spatial distribution and frequency of Leo IV's 'blue plume' stars reveals evidence for a young ({approx}2 Gyr) stellar population which makes up {approx}2% of its stellar mass. This sprinkling of star formation, only detectable in this deep study, highlights the need for further imaging of the new Milky Way satellites along with theoretical work on the expected, detailed properties of these possible 'reionization fossils'.« less

A Deeper Look at Leo IV: Star Formation History and Extended Structure

NASA Astrophysics Data System (ADS)

Sand, David J.; Seth, Anil; Olszewski, Edward W.; Willman, Beth; Zaritsky, Dennis; Kallivayalil, Nitya

2010-07-01

We present MMT/Megacam imaging of the Leo IV dwarf galaxy in order to investigate its structure and star formation history, and to search for signs of association with the recently discovered Leo V satellite. Based on parameterized fits, we find that Leo IV is round, with epsilon < 0.23 (at the 68% confidence limit) and a half-light radius of rh ~= 130 pc. Additionally, we perform a thorough search for extended structures in the plane of the sky and along the line of sight. We derive our surface brightness detection limit by implanting fake structures into our catalog with stellar populations identical to that of Leo IV. We show that we are sensitive to stream-like structures with surface brightness μ r <~ 29.6 mag arcsec-2, and at this limit we find no stellar bridge between Leo IV (out to a radius of ~0.5 kpc) and the recently discovered, nearby satellite Leo V. Using the color-magnitude fitting package StarFISH, we determine that Leo IV is consistent with a single age (~14 Gyr), single metallicity ([Fe/H] ~ -2.3) stellar population, although we cannot rule out a significant spread in these values. We derive a luminosity of MV = -5.5 ± 0.3. Studying both the spatial distribution and frequency of Leo IV's "blue plume" stars reveals evidence for a young (~2 Gyr) stellar population which makes up ~2% of its stellar mass. This sprinkling of star formation, only detectable in this deep study, highlights the need for further imaging of the new Milky Way satellites along with theoretical work on the expected, detailed properties of these possible "reionization fossils." Observations reported here were obtained at the MMT observatory, a joint facility of the Smithsonian Institution and the University of Arizona.

SEL2 servicing: increased science return via on-orbit propellant replenishment

NASA Astrophysics Data System (ADS)

Reed, Benjamin B.; DeWeese, Keith; Kienlen, Michael; Aranyos, Thomas; Pellegrino, Joseph; Bacon, Charles; Qureshi, Atif

2016-07-01

Spacecraft designers are driving observatories to the distant Sun-Earth Lagrange Point 2 (SEL2) to meet ever-increasing science requirements. The mass fraction dedicated to propellant for these observatories to reach and operate at SEL2 will be allocated with the upmost care, as it comes at the expense of optics and instrument masses. As such, these observatories could benefit from on-orbit refueling, allowing greater dry-to-wet mass ratio at launch and/or longer mission life. NASA is developing technologies, capabilities and integrated mission designs for multiple servicing applications in low Earth orbit (LEO), geosynchronous Earth orbit (GEO) and cisluner locations. Restore-L, a mission officially in formulation, will launch a free-flying robotic servicer to refuel a government-owned satellite in LEO by mid 2020. This paper will detail the results of a point design mission study to extend Restore-L servicing technologies from LEO to SEL2. This SEL2 mission would launch an autonomous, robotic servicer spacecraft equipped to extend the life of two space assets through refueling. Two space platforms were chosen to 1) drive the requirements for achieving SEL2 orbit and rendezvous with a spacecraft, and 2) to drive the requirements to translate within SEL2 to conduct a follow-on servicing mission. Two fuels, xenon and hydrazine, were selected to assess a multiple delivery system. This paper will address key mission drivers, such as servicer autonomy (necessitated due to communications latency at L2). Also discussed will be the value of adding cooperative servicing elements to the client observatories to reduce mission risk.

Estimating maneuvers for precise relative orbit determination using GPS

NASA Astrophysics Data System (ADS)

Allende-Alba, Gerardo; Montenbruck, Oliver; Ardaens, Jean-Sébastien; Wermuth, Martin; Hugentobler, Urs

2017-01-01

Precise relative orbit determination is an essential element for the generation of science products from distributed instrumentation of formation flying satellites in low Earth orbit. According to the mission profile, the required formation is typically maintained and/or controlled by executing maneuvers. In order to generate consistent and precise orbit products, a strategy for maneuver handling is mandatory in order to avoid discontinuities or precision degradation before, after and during maneuver execution. Precise orbit determination offers the possibility of maneuver estimation in an adjustment of single-satellite trajectories using GPS measurements. However, a consistent formulation of a precise relative orbit determination scheme requires the implementation of a maneuver estimation strategy which can be used, in addition, to improve the precision of maneuver estimates by drawing upon the use of differential GPS measurements. The present study introduces a method for precise relative orbit determination based on a reduced-dynamic batch processing of differential GPS pseudorange and carrier phase measurements, which includes maneuver estimation as part of the relative orbit adjustment. The proposed method has been validated using flight data from space missions with different rates of maneuvering activity, including the GRACE, TanDEM-X and PRISMA missions. The results show the feasibility of obtaining precise relative orbits without degradation in the vicinity of maneuvers as well as improved maneuver estimates that can be used for better maneuver planning in flight dynamics operations.

A Comparison of Atomic Oxygen Degradation in Low Earth Orbit and in a Plasma Etcher

NASA Technical Reports Server (NTRS)

Townsend, Jacqueline A.; Park, Gloria

1997-01-01

In low Earth orbit (LEO) significant degradation of certain materials occurs from exposure to atomic oxygen (AO). Orbital opportunities to study this degradation for specific materials are limited and expensive. While plasma etchers are commonly used in ground-based studies because of their low cost and convenience, the environment produced in an etcher chamber differs greatly from the LEO environment. Because of the differences in environment, the validity of using etcher data has remained an open question. In this paper, degradation data for 22 materials from the orbital experiment Evaluation of Oxygen Interaction with Materials (EOIM-3) are compared with data from EOIM-3 control specimens exposed in a typical plasma etcher. This comparison indicates that, when carefully considered, plasma etcher results can produce order-of-magnitude estimates of orbital degradation. This allows the etcher to be used to screen unacceptable materials from further, more expensive tests.

TOPEX/Poseidon precision orbit determination production and expert system

NASA Technical Reports Server (NTRS)

Putney, Barbara; Zelensky, Nikita; Klosko, Steven

1993-01-01

TOPEX/Poseidon (T/P) is a joint mission between NASA and the Centre National d'Etudes Spatiales (CNES), the French Space Agency. The TOPEX/Poseidon Precision Orbit Determination Production System (PODPS) was developed at Goddard Space Flight Center (NASA/GSFC) to produce the absolute orbital reference required to support the fundamental ocean science goals of this satellite altimeter mission within NASA. The orbital trajectory for T/P is required to have a RMS accuracy of 13 centimeters in its radial component. This requirement is based on the effective use of the satellite altimetry for the isolation of absolute long-wavelength ocean topography important for monitoring global changes in the ocean circulation system. This orbit modeling requirement is at an unprecedented accuracy level for this type of satellite. In order to routinely produce and evaluate these orbits, GSFC has developed a production and supporting expert system. The PODPS is a menu driven system allowing routine importation and processing of tracking data for orbit determination, and an evaluation of the quality of the orbit so produced through a progressive series of tests. Phase 1 of the expert system grades the orbit and displays test results. Later phases undergoing implementation, will prescribe corrective actions when unsatisfactory results are seen. This paper describes the design and implementation of this orbit determination production system and the basis for its orbit accuracy assessment within the expert system.

Comparison of TOPEX/Poseidon orbit determination solutions obtained by the Goddard Space Flight Center Flight Dynamics Division and Precision Orbit Determination Teams

NASA Technical Reports Server (NTRS)

Doll, C.; Mistretta, G.; Hart, R.; Oza, D.; Cox, C.; Nemesure, M.; Bolvin, D.; Samii, Mina V.

1993-01-01

Orbit determination results are obtained by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) using the Goddard Trajectory Determination System (GTDS) and a real-time extended Kalman filter estimation system to process Tracking Data and Relay Satellite (TDRS) System (TDRSS) measurements in support of the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft navigation and health and safety operations. GTDS is the operational orbit determination system used by the FDD, and the extended Kalman fliter was implemented in an analysis prototype system, the Real-Time Orbit Determination System/Enhanced (RTOD/E). The Precision Orbit Determination (POD) team within the GSFC Space Geodesy Branch generates an independent set of high-accuracy trajectories to support the TOPEX/Poseidon scientific data. These latter solutions use the Geodynamics (GEODYN) orbit determination system with laser ranging tracking data. The TOPEX/Poseidon trajectories were estimated for the October 22 - November 1, 1992, timeframe, for which the latest preliminary POD results were available. Independent assessments were made of the consistencies of solutions produced by the batch and sequential methods. The batch cases were assessed using overlap comparisons, while the sequential cases were assessed with covariances and the first measurement residuals. The batch least-squares and forward-filtered RTOD/E orbit solutions were compared with the definitive POD orbit solutions. The solution differences were generally less than 10 meters (m) for the batch least squares and less than 18 m for the sequential estimation solutions. The differences among the POD, GTDS, and RTOD/E solutions can be traced to differences in modeling and tracking data types, which are being analyzed in detail.

The world state of orbital debris measurements and modeling

NASA Astrophysics Data System (ADS)

Johnson, Nicholas L.

2004-02-01

For more than 20 years orbital debris research around the world has been striving to obtain a sharper, more comprehensive picture of the near-Earth artificial satellite environment. Whereas significant progress has been achieved through better organized and funded programs and with the assistance of advancing technologies in both space surveillance sensors and computational capabilities, the potential of measurements and modeling of orbital debris has yet to be realized. Greater emphasis on a systems-level approach to the characterization and projection of the orbital debris environment would prove beneficial. On-going space surveillance activities, primarily from terrestrial-based facilities, are narrowing the uncertainties of the orbital debris population for objects greater than 2 mm in LEO and offer a better understanding of the GEO regime down to 10 cm diameter objects. In situ data collected in LEO is limited to a narrow range of altitudes and should be employed with great care. Orbital debris modeling efforts should place high priority on improving model fidelity, on clearly and completely delineating assumptions and simplifications, and on more thorough sensitivity studies. Most importantly, however, greater communications and cooperation between the measurements and modeling communities are essential for the efficient advancement of the field. The advent of the Inter-Agency Space Debris Coordination Committee (IADC) in 1993 has facilitated this exchange of data and modeling techniques. A joint goal of these communities should be the identification of new sources of orbital debris.

A Study into the Method of Precise Orbit Determination of a HEO Orbiter by GPS and Accelerometer

NASA Technical Reports Server (NTRS)

Ikenaga, Toshinori; Hashida, Yoshi; Unwin, Martin

2007-01-01

In the present day, orbit determination by Global Positioning System (GPS) is not unusual. Especially for low-cost small satellites, position determination by an on-board GPS receiver provides a cheap, reliable and precise method. However, the original purpose of GPS is for ground users, so the transmissions from all of the GPS satellites are directed toward the Earth s surface. Hence there are some restrictions for users above the GPS constellation to detect those signals. On the other hand, a desire for precise orbit determination for users in orbits higher than GPS constellation exists. For example, the next Japanese Very Long Baseline Interferometry (VLBI) mission "ASTRO-G" is trying to determine its orbit in an accuracy of a few centimeters at apogee. The use of GPS is essential for such ultra accurate orbit determination. This study aims to construct a method for precise orbit determination for such high orbit users, especially in High Elliptical Orbits (HEOs). There are several approaches for this objective. In this study, a hybrid method with GPS and an accelerometer is chosen. Basically, while the position cannot be determined by an on-board GPS receiver or other Range and Range Rate (RARR) method, all we can do to estimate the user satellite s position is to propagate the orbit along with the force model, which is not perfectly correct. However if it has an accelerometer (ACC), the coefficients of the air drag and the solar radiation pressure applied to the user satellite can be updated and then the propagation along with the "updated" force model can improve the fitting accuracy of the user satellite s orbit. In this study, it is assumed to use an accelerometer available in the present market. The effects by a bias error of an accelerometer will also be discussed in this paper.

Time Serial Analysis of the Induced LEO Environment within the ISS 6A

NASA Technical Reports Server (NTRS)

Wilson, John W.; Nealy, John E.; Tomov, B. T.; Cucinotta, Francis A.; Badavi, Frank F.; DeAngelis, Giovanni; Atwell, William; Leutke, N.

2006-01-01

Anisotropies in the low Earth orbit (LEO) radiation environment were found to influence the thermoluminescence detectors (TLD) dose within the (International Space Station) ISS 7A Service Module. Subsequently, anisotropic environmental models with improved dynamic time extrapolation have been developed including westward and northern drifts using AP8 Min & Max as estimates of the historic spatial distribution of trapped protons in the 1965 and 1970 era, respectively. In addition, a directional dependent geomagnetic cutoff model was derived for geomagnetic field configurations from the 1945 to 2020 time frame. A dynamic neutron albedo model based on our atmospheric radiation studies has likewise been required to explain LEO neutron measurements. The simultaneous measurements of dose and dose rate using four Liulin instruments at various locations in the US LAB and Node 1 has experimentally demonstrated anisotropic effects in ISS 6A and are used herein to evaluate the adequacy of these revised environmental models.

Dealing with Uncertainties in Initial Orbit Determination

NASA Technical Reports Server (NTRS)

Armellin, Roberto; Di Lizia, Pierluigi; Zanetti, Renato

2015-01-01

A method to deal with uncertainties in initial orbit determination (IOD) is presented. This is based on the use of Taylor differential algebra (DA) to nonlinearly map the observation uncertainties from the observation space to the state space. When a minimum set of observations is available DA is used to expand the solution of the IOD problem in Taylor series with respect to measurement errors. When more observations are available high order inversion tools are exploited to obtain full state pseudo-observations at a common epoch. The mean and covariance of these pseudo-observations are nonlinearly computed by evaluating the expectation of high order Taylor polynomials. Finally, a linear scheme is employed to update the current knowledge of the orbit. Angles-only observations are considered and simplified Keplerian dynamics adopted to ease the explanation. Three test cases of orbit determination of artificial satellites in different orbital regimes are presented to discuss the feature and performances of the proposed methodology.

Effect of LEO cycling on 125 Ah advanced design IPV nickel-hydrogen battery cells

NASA Technical Reports Server (NTRS)

Smithrick, John J.; Hall, Stephen W.

1990-01-01

An advanced 125 Ah individual pressure vessel (IPV) nickel-hydrogen cell was designed. The primary function of the advanced cell, is to store and deliver energy for long term, low earth-orbit (LEO) spacecraft missions. The new features of this design are: (1) use of 26 percent rather than 31 percent potassium hydroxide (KOH) electrolyte, (2) use of a patented catalyzed wall wick, (3) use of serrated edge separators to facilitate gaseous oxygen and hydrogen flow within the cell, while still maintaining physical contact with the wall wick for electrolyte management, and (4) use of a floating rather than a fixed stack (state-of-the-art) to accommodate nickel electrode expansion. Six 125 Ah flight cells based on this design were fabricated by Eagle-Picher. Three of the cells contain all of the advanced features (test cells) and three are the same as the test cells except they don't have catalyst on the wall wick (control cells). All six cells are in the process of being evaluated in a LEO cycle life test. The cells have accumulated about 4700 LEO cycles (60 percent DOD 10 C). There have been no cell failures, the catalyzed wall wick cells however, are performing better.

Effect of LEO cycling on 125 Ah advanced design IPV nickel-hydrogen battery cells

NASA Technical Reports Server (NTRS)

Smithrick, John J.; Hall, Stephen W.

1990-01-01

An advanced 125 Ah individual pressure vessel (IPV) nickel-hydrogen cell was designed. The primary function of the advanced cell is to store and deliver energy for long-term, low earth-orbit (LEO) spacecraft missions. The new features of this design are: (1) use of 26 percent rather than 31 percent potassium hydroxide (KOH) electrolyte, (2) use of a patented catalyzed wall wick, (3) use of serrated-edge separators to facilitate gaseous oxygen and hydrogen flow within the cell, while still maintaining physical contact with the wall wick for electrolyte management, and (4) use of a floating rather than a fixed stack (state-of-the-art) to accommodate nickel electrode expansion. Six 125-Ah flight cells based on this design were fabricated by Eagle-Picher. Three of the cells contain all of the advanced features (test cells) and three are the same as the test cells except they don't have catalyst on the wall wick (control cells). All six cells are in the process of being evaluated in a LEO cycle life test. The cells have accumulated about 4700 LEO cycles (60 percent DOD 10 C). There have been no cell failures; the catalyzed wall wick cells, however, are performing better.

CASTOR: Cathode/Anode Satellite Thruster for Orbital Repositioning

NASA Technical Reports Server (NTRS)

Mruphy, Gloria A.

2010-01-01

The purpose of CASTOR (Cathode/Anode Satellite Thruster for Orbital Repositioning) satellite is to demonstrate in Low Earth Orbit (LEO) a nanosatellite that uses a Divergent Cusped Field Thruster (DCFT) to perform orbital maneuvers representative of an orbital transfer vehicle. Powered by semi-deployable solar arrays generating 165W of power, CASTOR will achieve nearly 1 km/s of velocity increment over one year. As a technology demonstration mission, success of CASTOR in LEO will pave the way for a low cost, high delta-V orbital transfer capability for small military and civilian payloads in support of Air Force and NASA missions. The educational objective is to engage graduate and undergraduate students in critical roles in the design, development, test, carrier integration and on-orbit operations of CASTOR as a supplement to their curricular activities. This program is laying the foundation for a long-term satellite construction program at MIT. The satellite is being designed as a part of AFRL's University Nanosatellite Program, which provides the funding and a framework in which student satellite teams compete for a launch to orbit. To this end, the satellite must fit within an envelope of 50cmx50cmx60cm, have a mass of less than 50kg, and meet stringent structural and other requirements. In this framework, the CASTOR team successfully completed PDR in August 2009 and CDR in April 2010 and will compete at FCR (Flight Competition Review) in January 2011. The complexity of the project requires implementation of many systems engineering techniques which allow for development of CASTOR from conception through FCR and encompass the full design, fabrication, and testing process.

Variable stars in the Leo A dwarf galaxy (DDO 69)

NASA Technical Reports Server (NTRS)

Hoessel, John G.; Saha, A.; Krist, John; Danielson, G. Edward

1994-01-01

Observations of the Leo A dwarf galaxy, obtained over the period from 1980 to 1991 are reported. Forty two separate Charge Coupled Devices (CCD) frames were searched for variable stars. A total of 14 suspected variables were found, 9 had sufficient coverage for period determination, and 5 had Cepheid light curves. Four of these stars fit well on a P-L relation and yield a distance modulus, after correction for Galactic foreground extinction, of m-M = 26.74. This corresponds to a distance of 2.2 Mpc, placing Leo A near the Local Group zero-velocity surface.

Strategies for high-precision Global Positioning System orbit determination

NASA Technical Reports Server (NTRS)

Lichten, Stephen M.; Border, James S.

1987-01-01

Various strategies for the high-precision orbit determination of the GPS satellites are explored using data from the 1985 GPS field test. Several refinements to the orbit determination strategies were found to be crucial for achieving high levels of repeatability and accuracy. These include the fine tuning of the GPS solar radiation coefficients and the ground station zenith tropospheric delays. Multiday arcs of 3-6 days provided better orbits and baselines than the 8-hr arcs from single-day passes. Highest-quality orbits and baselines were obtained with combined carrier phase and pseudorange solutions.

THE ARECIBO LEGACY FAST ALFA SURVEY. IX. THE LEO REGION H I CATALOG, GROUP MEMBERSHIP, AND THE H I MASS FUNCTION FOR THE LEO I GROUP

DOE Office of Scientific and Technical Information (OSTI.GOV)

Stierwalt, Sabrina; Haynes, Martha P.; Giovanelli, Riccardo

We present the catalog of H I sources extracted from the ongoing Arecibo Legacy Fast ALFA (ALFALFA) extragalactic H I line survey, found within the sky region bounded by 9{sup h}36{sup m} < {alpha} < 11{sup h}36{sup m} and +08{sup 0} < {delta} < +12{sup 0}. The H I catalog presented here for this 118 deg{sup 2} region is combined with the ones derived from surrounding regions also covered by the ALFALFA survey to examine the large-scale structure in the complex Leo region. Because of the combination of wide sky coverage and superior sensitivity, spatial and spectral resolution, the ALFALFAmore » H I catalog of the Leo region improves significantly on the numbers of low H I mass sources as compared with those found in previous H I surveys. The H I mass function of the Leo I group presented here is dominated by low-mass objects: 45 of the 65 Leo I members have M{sub H{sub l}}<10{sup 8} M-odot, yielding tight constraints on the low-mass slope of the Leo I H I mass function. The best-fit slope is {alpha} {approx_equal} -1.41 + 0.2 - 0.1. A direct comparison between the ALFALFA H I line detections and an optical search of the Leo I region proves the advantage of the ALFALFA strategy in finding low-mass, gas-rich dwarfs. These results suggest the existence of a significant population of low surface brightness, gas-rich, yet still very low H I mass galaxies, and may reflect the same type of morphological segregation as is seen in the Local Group. While the low-mass end slope of the Leo I H I mass function is steeper than that determined for luminosity functions of the group, the slope still falls short of the values predicted by simulations of structure formation in the lambda cold dark matter paradigm.« less

LEO P: AN UNQUENCHED VERY LOW-MASS GALAXY

DOE Office of Scientific and Technical Information (OSTI.GOV)

McQuinn, Kristen B. W.; Skillman, Evan D.; Berg, Danielle

Leo P is a low-luminosity dwarf galaxy discovered through the blind H i Arecibo Legacy Fast ALFA survey. The H i and follow-up optical observations have shown that Leo P is a gas-rich dwarf galaxy with active star formation, an underlying older population, and an extremely low oxygen abundance. We have obtained optical imaging with the Hubble Space Telescope to two magnitudes below the red clump in order to study the evolution of Leo P. We refine the distance measurement to Leo P to be 1.62 ± 0.15 Mpc, based on the luminosity of the horizontal branch stars and 10more » newly identified RR Lyrae candidates. This places the galaxy at the edge of the Local Group, ∼0.4 Mpc from Sextans B, the nearest galaxy in the NGC 3109 association of dwarf galaxies of which Leo P is clearly a member. The star responsible for ionizing the H ii region is most likely an O7V or O8V spectral type, with a stellar mass ≳25 M{sub ⊙}. The presence of this star provides observational evidence that massive stars at the upper end of the initial mass function are capable of being formed at star formation rates as low as ∼10{sup −5} M{sub ⊙} yr{sup −1}. The best-fitting star formation history (SFH) derived from the resolved stellar populations of Leo P using the latest PARSEC models shows a relatively constant star formation rate over the lifetime of the galaxy. The modeled luminosity characteristics of Leo P at early times are consistent with low-luminosity dSph Milky Way satellites, suggesting that Leo P is what a low-mass dSph would look like if it evolved in isolation and retained its gas. Despite the very low mass of Leo P, the imprint of reionization on its SFH is subtle at best, and consistent with being totally negligible. The isolation of Leo P, and the total quenching of star formation of Milky Way satellites of similar mass, implies that the local environment dominates the quenching of the Milky Way satellites.« less

Applicability of meteor radiant determination methods depending on orbit type. II. Low-eccentric orbits

NASA Astrophysics Data System (ADS)

Svoren, J.; Neslusan, L.; Porubcan, V.

1994-08-01

All known parent bodies of meteor showers belong to bodies moving in high-eccentricity orbits (e => 0.5). Recently, asteroids in low-eccentricity orbits (e < 0.5) approaching the Earth's orbit, were suggested as another population of possible parent bodies of meteor streams. This paper deals with the problem of calculation of meteor radiants connected with the bodies in low-eccentricity orbits from the point of view of optimal results depending on the method applied. The paper is a continuation of our previous analysis of high-eccentricity orbits (Svoren, J., Neslusan, L., Porubcan, V.: 1993, Contrib. Astron. Obs. Skalnate Pleso 23, 23). Some additional methods resulting from mathematical modelling are presented and discussed together with Porter's, Steel-Baggaley's and Hasegawa's methods. In order to be able to compare how suitable the application of the individual radiant determination methods is, it is necessary to determine the accuracy with which they approximate real meteor orbits. To verify the accuracy with which the orbit of a meteoroid with at least one node at 1 AU fits the original orbit of the parent body, the Southworth-Hawkins D-criterion (Southworth, R.B., Hawkins, G.S.: 1963, Smithson. Contr. Astrophys. 7, 261) was applied. D <= 0.1 indicates a very good fit of orbits, 0.1 < D <= 0.2 is considered for a good fit and D > 0.2 means that the fit is rather poor and the change of orbit unrealistic. The optimal method, i.e. the one which results in the smallest D values for the population of low-eccentricity orbits, is that of adjusting the orbit by varying both the eccentricity and perihelion distance. A comparison of theoretical radiants obtained by various methods was made for typical representatives from each group of the NEA (near-Earth asteroids) objects.

Real-time on-board orbit determination with DORIS

NASA Technical Reports Server (NTRS)

Berthias, J.-P.; Jayles, C.; Pradines, D.

1993-01-01

A spaceborne orbit determination system is being developed by the French Space Agency (CNES) for the SPOT 4 satellite. It processes DORIS measurements to produce an orbit with an accuracy of about 50O meters rms. In order to evaluate the reliability of the software, it was combined with the MERCATOR man/machine interface and used to process the TOPEX/Poseidon DORIS data in near real time during the validation phase of the instrument, at JPL and at CNES. This paper gives an overview of the orbit determination system and presents the results of the TOPEX/Poseidon experiment.

Improved solution accuracy for TDRSS-based TOPEX/Poseidon orbit determination

NASA Technical Reports Server (NTRS)

Doll, C. E.; Mistretta, G. D.; Hart, R. C.; Oza, D. H.; Bolvin, D. T.; Cox, C. M.; Nemesure, M.; Niklewski, D. J.; Samii, M. V.

1994-01-01

Orbit determination results are obtained by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) using a batch-least-squares estimator available in the Goddard Trajectory Determination System (GTDS) and an extended Kalman filter estimation system to process Tracking and Data Relay Satellite (TDRS) System (TDRSS) measurements. GTDS is the operational orbit determination system used by the FDD in support of the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft navigation and health and safety operations. The extended Kalman filter was implemented in an orbit determination analysis prototype system, closely related to the Real-Time Orbit Determination System/Enhanced (RTOD/E) system. In addition, the Precision Orbit Determination (POD) team within the GSFC Space Geodesy Branch generated an independent set of high-accuracy trajectories to support the TOPEX/Poseidon scientific data. These latter solutions use the geodynamics (GEODYN) orbit determination system with laser ranging and Doppler Orbitography and Radiopositioning integrated by satellite (DORIS) tracking measurements. The TOPEX/Poseidon trajectories were estimated for November 7 through November 11, 1992, the timeframe under study. Independent assessments were made of the consistencies of solutions produced by the batch and sequential methods. The batch-least-squares solutions were assessed based on the solution residuals, while the sequential solutions were assessed based on primarily the estimated covariances. The batch-least-squares and sequential orbit solutions were compared with the definitive POD orbit solutions. The solution differences were generally less than 2 meters for the batch-least-squares and less than 13 meters for the sequential estimation solutions. After the sequential estimation solutions were processed with a smoother algorithm, position differences with POD orbit solutions of less than 7 meters were obtained. The differences among the POD, GTDS, and filter

Astrodynamics. Volume 1 - Orbit determination, space navigation, celestial mechanics.

NASA Technical Reports Server (NTRS)

Herrick, S.

1971-01-01

Essential navigational, physical, and mathematical problems of space exploration are covered. The introductory chapters dealing with conic sections, orientation, and the integration of the two-body problem are followed by an introduction to orbit determination and design. Systems of units and constants, as well as ephemerides, representations, reference systems, and data are then dealt with. A detailed attention is given to rendezvous problems and to differential processes in observational orbit correction, and in rendezvous or guidance correction. Finally, the Laplacian methods for determining preliminary orbits, and the orbit methods of Lagrange, Gauss, and Gibbs are reviewed.

Fuel-Optimal Altitude Maintenance of Low-Earth-Orbit Spacecrafts by Combined Direct/Indirect Optimization

NASA Astrophysics Data System (ADS)

Kim, Kyung-Ha; Park, Chandeok; Park, Sang-Young

2015-12-01

This work presents fuel-optimal altitude maintenance of Low-Earth-Orbit (LEO) spacecrafts experiencing non-negligible air drag and J2 perturbation. A pseudospectral (direct) method is first applied to roughly estimate an optimal fuel consumption strategy, which is employed as an initial guess to precisely determine itself. Based on the physical specifications of KOrea Multi-Purpose SATellite-2 (KOMPSAT-2), a Korean artificial satellite, numerical simulations show that a satellite ascends with full thrust at the early stage of the maneuver period and then descends with null thrust. While the thrust profile is presumably bang-off, it is difficult to precisely determine the switching time by using a pseudospectral method only. This is expected, since the optimal switching epoch does not coincide with one of the collocation points prescribed by the pseudospectral method, in general. As an attempt to precisely determine the switching time and the associated optimal thrust history, a shooting (indirect) method is then employed with the initial guess being obtained through the pseudospectral method. This hybrid process allows the determination of the optimal fuel consumption for LEO spacecrafts and their thrust profiles efficiently and precisely.

Advantages of High vs. Low Earth Orbit for SIRTF

NASA Technical Reports Server (NTRS)

Eisenhardt, Peter; Werner, Michael W.

1989-01-01

While the subject of this workshop, which we will refer to as ET (for Enlightenment Telescope), is a dazzling successor to the Hubble Space Telescope, its location is unlikely to be the Low Earth Orbit (LEO) used by HST. Locations suggested for ET include High Earth Orbit (HEO) and the moon. The first space telescope to occupy HEO will be the liquid helium cooled Space Infrared Telescope Facility (SIRTF). The selection of HEO for SIRTF was the outcome of a recent study led by the Ames Research Center which showed significant advantages for SIRTF in HEO vs. LEO. This article summarizes the main results of that study. We begin with a review of SIRTF's rationale and requirements, in part because the IR capabilities and low temperature proposed for ET make it something of a successor to SIRTF as well as to HST. We conclude with some comments about another possible location for both SIRTF and ET, the Earth-Sun L2 Lagrangian point.

Hugoniot Models for Na and LiF from LEOS

DOE Office of Scientific and Technical Information (OSTI.GOV)

Whitley, Heather D.; Wu, Christine J.

2016-10-12

In this document, we provide the Hugoniot for sodium from two models: LEOS table L110 and Lynx table 110. We also provide the Hugoniot for lithium fluoride from LEOS (L2240) and Lynx (2240). The Hugoniot pressures are supplied for temperatures between 338.0 and 1.16×10 9 Kelvin and densities between 0.968 and 11.5 g/cc. These LEOS models were developed by the quotidian EOS methodology, which is a widely used and robust method for producing tabular EOS data. Tables list the model data for LEOS 110, Lynx 110, LEOS 2240, and Lynx 2240. The Lynx models follow the same methodology as themore » LEOS models; however, the Purgatorio average-atom DFT code was used to compute the electron thermal part of the EOS. The models for Lynx are only listed at high compression due to known issues with the Lynx library at lower pressures.« less

Evaluation of Low-Earth-Orbit Environmental Effects on International Space Station Thermal Control Materials

NASA Technical Reports Server (NTRS)

Dever, Joyce A.

1998-01-01

Many spacecraft thermal control coatings in low Earth orbit (LEO) can be affected by solar ultraviolet radiation and atomic oxygen. Ultraviolet radiation can darken some polymers and oxides commonly used in thermal control materials. Atomic oxygen can erode polymer materials, but it may reverse the ultraviolet-darkening effect on oxides. Maintaining the desired solar absorptance for thermal control coatings is important to assure the proper operating temperature of the spacecraft. Thermal control coatings to be used on the International Space Station (ISS) were evaluated for their performance after exposure in the NASA Lewis Research Center's Atomic Oxygen-Vacuum Ultraviolet Exposure (AO-VUV) facility. This facility simulated the LEO environments of solar vacuum ultraviolet (VUV) radiation (wavelength range, 115 to 200 nanometers (nm)) and VUV combined with atomic oxygen. Solar absorptance was measured in vacuo to eliminate the "bleaching" effects of ambient oxygen on VUV-induced degradation. The objective of these experiments was to determine solar absorptance increases of various thermal control materials due to exposure to simulated LEO conditions similar to those expected for ISS. Work was done in support of ISS efforts at the requests of Boeing Space and Defense Systems and Lockheed Martin Vought Systems.

Space environmental effects on spacecraft: LEO materials selection guide, part 2

NASA Astrophysics Data System (ADS)

Silverman, Edward M.

1995-08-01

This document provides performance properties on major spacecraft materials and subsystems that have been exposed to the low-Earth orbit (LEO) space environment. Spacecraft materials include metals, polymers, composites, white and black paints, thermal-control blankets, adhesives, and lubricants. Spacecraft subsystems include optical components, solar cells, and electronics. Information has been compiled from LEO short-term spaceflight experiments (e.g., space shuttle) and from retrieved satellites of longer mission durations (e.g., Long Duration Exposure Facility). Major space environment effects include atomic oxygen (AO), ultraviolet radiation, micrometeoroids and debris, contamination, and particle radiation. The main objective of this document is to provide a decision tool to designers for designing spacecraft and structures. This document identifies the space environments that will affect the performance of materials and components, e.g., thermal-optical property changes of paints due to UV exposures, AO-induced surface erosion of composites, dimensional changes due to thermal cycling, vacuum-induced moisture outgassing, and surface optical changes due to AO/UV exposures. Where appropriate, relationships between the space environment and the attendant material/system effects are identified. Part 2 covers thermal control systems, power systems, optical components, electronic systems, and applications.

Space environmental effects on spacecraft: LEO materials selection guide, part 2

NASA Technical Reports Server (NTRS)

Silverman, Edward M.

1995-01-01

This document provides performance properties on major spacecraft materials and subsystems that have been exposed to the low-Earth orbit (LEO) space environment. Spacecraft materials include metals, polymers, composites, white and black paints, thermal-control blankets, adhesives, and lubricants. Spacecraft subsystems include optical components, solar cells, and electronics. Information has been compiled from LEO short-term spaceflight experiments (e.g., space shuttle) and from retrieved satellites of longer mission durations (e.g., Long Duration Exposure Facility). Major space environment effects include atomic oxygen (AO), ultraviolet radiation, micrometeoroids and debris, contamination, and particle radiation. The main objective of this document is to provide a decision tool to designers for designing spacecraft and structures. This document identifies the space environments that will affect the performance of materials and components, e.g., thermal-optical property changes of paints due to UV exposures, AO-induced surface erosion of composites, dimensional changes due to thermal cycling, vacuum-induced moisture outgassing, and surface optical changes due to AO/UV exposures. Where appropriate, relationships between the space environment and the attendant material/system effects are identified. Part 2 covers thermal control systems, power systems, optical components, electronic systems, and applications.

Applicability of meteor radiant determination methods depending on orbit type. I. High-eccentric orbits

NASA Astrophysics Data System (ADS)

Svoren, J.; Neslusan, L.; Porubcan, V.

1993-07-01

It is evident that there is no uniform method of calculating meteor radiants which would yield reliable results for all types of cometary orbits. In the present paper an analysis of this problem is presented, together with recommended methods for various types of orbits. Some additional methods resulting from mathematical modelling are presented and discussed together with Porter's, Steel-Baggaley's and Hasegawa's methods. In order to be able to compare how suitable the application of the individual radiant determination methods is, it is necessary to determine the accuracy with which they approximate real meteor orbits. To verify the accuracy with which the orbit of a meteoroid with at least one node at 1 AU fits the original orbit of the parent body, we applied the Southworth-Hawkins D-criterion (Southworth, R.B., Hawkins, G.S.: 1963, Smithson. Contr. Astrophys 7, 261). D<=0.1 indicates a very good fit of orbits, 0.1orbits, and D>0.2 the fit is rather poor and the change of orbit unrealistic. The optimal methods with the smallest values of D for given types of orbits are shown in two series of six plots. The new method of rotation around the line of apsides we propose is very appropriate in the region of small inclinations. There is no doubt that Hasegawa's omega-adjustment method (Hasegawa, I.: 1990, Publ. Astron. Soc. Japan 42, 175) has the widest application. A comparison of the theoretical radiants with the observed radiants of seven known meteor showers is also presented.

Algorithms for Autonomous GPS Orbit Determination and Formation Flying: Investigation of Initialization Approaches and Orbit Determination for HEO

NASA Technical Reports Server (NTRS)

Axelrad, Penina; Speed, Eden; Leitner, Jesse A. (Technical Monitor)

2002-01-01

This report summarizes the efforts to date in processing GPS measurements in High Earth Orbit (HEO) applications by the Colorado Center for Astrodynamics Research (CCAR). Two specific projects were conducted; initialization of the orbit propagation software, GEODE, using nominal orbital elements for the IMEX orbit, and processing of actual and simulated GPS data from the AMSAT satellite using a Doppler-only batch filter. CCAR has investigated a number of approaches for initialization of the GEODE orbit estimator with little a priori information. This document describes a batch solution approach that uses pseudorange or Doppler measurements collected over an orbital arc to compute an epoch state estimate. The algorithm is based on limited orbital element knowledge from which a coarse estimate of satellite position and velocity can be determined and used to initialize GEODE. This algorithm assumes knowledge of nominal orbital elements, (a, e, i, omega, omega) and uses a search on time of perigee passage (tau(sub p)) to estimate the host satellite position within the orbit and the approximate receiver clock bias. Results of the method are shown for a simulation including large orbital uncertainties and measurement errors. In addition, CCAR has attempted to process GPS data from the AMSAT satellite to obtain an initial estimation of the orbit. Limited GPS data have been received to date, with few satellites tracked and no computed point solutions. Unknown variables in the received data have made computations of a precise orbit using the recovered pseudorange difficult. This document describes the Doppler-only batch approach used to compute the AMSAT orbit. Both actual flight data from AMSAT, and simulated data generated using the Satellite Tool Kit and Goddard Space Flight Center's Flight Simulator, were processed. Results for each case and conclusion are presented.

Design of a Representative Low Earth Orbit Satellite to Improve Existing Debris Models

NASA Technical Reports Server (NTRS)

Clark, S.; Dietrich, A.; Werremeyer, M.; Fitz-Coy, N.; Liou, J.-C.

2012-01-01

This paper summarizes the process and methodologies used in the design of a small-satellite, DebriSat, that represents materials and construction methods used in modern day Low Earth Orbit (LEO) satellites. This satellite will be used in a future hypervelocity impact test with the overall purpose to investigate the physical characteristics of modern LEO satellites after an on-orbit collision. The major ground-based satellite impact experiment used by DoD and NASA in their development of satellite breakup models was conducted in 1992. The target used for that experiment was a Navy Transit satellite (40 cm, 35 kg) fabricated in the 1960 s. Modern satellites are very different in materials and construction techniques from a satellite built 40 years ago. Therefore, there is a need to conduct a similar experiment using a modern target satellite to improve the fidelity of the satellite breakup models. The design of DebriSat will focus on designing and building a next-generation satellite to more accurately portray modern satellites. The design of DebriSat included a comprehensive study of historical LEO satellite designs and missions within the past 15 years for satellites ranging from 10 kg to 5000 kg. This study identified modern trends in hardware, material, and construction practices utilized in recent LEO missions, and helped direct the design of DebriSat.

Intial orbit determination results for Jason-1: towards a 1-cm orbit

NASA Technical Reports Server (NTRS)

Haines, B. J.; Haines, B.; Bertiger, W.; Desai, S.; Kuang, D.; Munson, T.; Reichert, A.; Young, L.; Willis, P.

2002-01-01

The U.S/France Jason-1 oceanographic mission is carrying state-of-the-art radiometric tracking systems (GPS and Doris) to support precise orbit determination (POD) requirements. The performance of the systems is strongly reflected in the early POD results. Results of both internal and external (e.g., satellite laser ranging) comparisons support that the 2.5 cm radial Rh4S requirement is being readily met, and provide reasons for optimism that 1 cm can be achieved. We discuss the POD strategy underlying these orbits, as well as the challenging issues that bear on the understanding and characterization of an orbit solution at the l-cm level. We also describe a system for producing science quality orbits in near real time in order to support emerging applications in operational oceanography.

AH Leo and the Blazhko Effect

NASA Astrophysics Data System (ADS)

Phillips, J.; Gay, P. L.

2004-12-01

We obtained 563 V-Band observations of AH Leo between January 27 and May 12, 2004. All observations were obtained with a 12-inch Schmidt-Cassegrain located on the island of Saipan, in the Commonwealth of the Northern Mariana Islands. We show that AH Leo is a type RRab RR Lyrae star with a minimum magnitude of V=14.658 magnitudes, a maximum amplitude of 0.989 magnitudes and a minimum amplitude of perhaps just 0.4 magnitudes. Its primary period is 0.4662609 days. Our observations also confirm the presence of the Blazhko effect, which had previously been detected by Smith and Gay (private communication) in 1993 and 1994. We estimate the Blazhko period to be roughly 20-days, however poor phase coverage at maximum light makes exact determination impossible. We also note that the bump during minimum, which is common in many RR Lyraes, varied throughout the Blazhko cycle, demonstrating amplitudes between 0 and 0.15 magnitudes. We would like to thank Sarah Maddison and Swinburne Astronomy Online for supporting this project

NASA's New Orbital Debris Engineering Model, ORDEM2010

NASA Technical Reports Server (NTRS)

Krisko, Paula H.

2010-01-01

This paper describes the functionality and use of ORDEM2010, which replaces ORDEM2000, as the NASA Orbital Debris Program Office (ODPO) debris engineering model. Like its predecessor, ORDEM2010 serves the ODPO mission of providing spacecraft designers/operators and debris observers with a publicly available model to calculate orbital debris flux by current-state-of-knowledge methods. The key advance in ORDEM2010 is the input file structure of the yearly debris populations from 1995-2035 of sizes 10 micron - 1 m. These files include debris from low-Earth orbits (LEO) through geosynchronous orbits (GEO). Stable orbital elements (i.e., those that do not randomize on a sub-year timescale) are included in the files as are debris size, debris number, material density, random error and population error. Material density is implemented from ground-test data into the NASA breakup model and assigned to debris fragments accordingly. The random and population errors are due to machine error and uncertainties in debris sizes. These high-fidelity population files call for a much higher-level model analysis than what was possible with the populations of ORDEM2000. Population analysis in the ORDEM2010 model consists of mapping matrices that convert the debris population elements to debris fluxes. One output mode results in a spacecraft encompassing 3-D igloo of debris flux, compartmentalized by debris size, velocity, pitch, and yaw with respect to spacecraft ram direction. The second output mode provides debris flux through an Earth-based telescope/radar beam from LEO through GEO. This paper compares the new ORDEM2010 with ORDEM2000 in terms of processes and results with examples of specific orbits.

Using Static Percentiles of AE9/AP9 to Approximate Dynamic Monte Carlo Runs for Radiation Analysis of Spiral Transfer Orbits

NASA Astrophysics Data System (ADS)

Kwan, Betty P.; O'Brien, T. Paul

2015-06-01

The Aerospace Corporation performed a study to determine whether static percentiles of AE9/AP9 can be used to approximate dynamic Monte Carlo runs for radiation analysis of spiral transfer orbits. Solar panel degradation is a major concern for solar-electric propulsion because solar-electric propulsion depends on the power output of the solar panel. Different spiral trajectories have different radiation environments that could lead to solar panel degradation. Because the spiral transfer orbits only last weeks to months, an average environment does not adequately address the possible transient enhancements of the radiation environment that must be accounted for in optimizing the transfer orbit trajectory. Therefore, to optimize the trajectory, an ensemble of Monte Carlo simulations of AE9/AP9 would normally be run for every spiral trajectory to determine the 95th percentile radiation environment. To avoid performing lengthy Monte Carlo dynamic simulations for every candidate spiral trajectory in the optimization, we found a static percentile that would be an accurate representation of the full Monte Carlo simulation for a representative set of spiral trajectories. For 3 LEO to GEO and 1 LEO to MEO trajectories, a static 90th percentile AP9 is a good approximation of the 95th percentile fluence with dynamics for 4-10 MeV protons, and a static 80th percentile AE9 is a good approximation of the 95th percentile fluence with dynamics for 0.5-2 MeV electrons. While the specific percentiles chosen cannot necessarily be used in general for other orbit trade studies, the concept of determining a static percentile as a quick approximation to a full Monte Carlo ensemble of simulations can likely be applied to other orbit trade studies. We expect the static percentile to depend on the region of space traversed, the mission duration, and the radiation effect considered.

A low Earth orbit molecular beam space simulation facility

NASA Technical Reports Server (NTRS)

Cross, J. B.

1984-01-01

A brief synopsis of the low Earth orbit (LEO) satellite environment is presented including neutral and ionic species. Two ground based atomic and molecular beam instruments are described which are capable of simulating the interaction of spacecraft surfaces with the LEO environment and detecting the results of these interactions. The first detects mass spectrometrically low level fluxes of reactively and nonreactively surface scattered species as a function of scattering angle and velocity while the second ultrahigh velocity (UHV) molecular beam, laser induced fluorescence apparatus is capable of measuring chemiluminescence produced by either gas phase or gas-surface interactions. A number of proposed experiments are described.

Long-Duration Human Habitation Beyond Low-Earth Orbit: Why is the Near Future Critical?

NASA Technical Reports Server (NTRS)

Thronson, Harley A.; Lester, Daniel; Hatfield, C. A.

2011-01-01

For more than a decade, habitation systems capable of comfortable human occupation and effective operations beyond low-Earth orbit (LEO) for more than a few weeks have been a priority recommendation to NASA. This capability is a lynch pin for human exploration beyond the Earth-Moon system. Here we describe briefly some relevant concepts and discuss justifications in the current political and financial environment for why near-term human habitation systems beyond LEO is an imperative.

Dawn Orbit Determination Team : Trajectory Modeling and Reconstruction Processes at Vesta

NASA Technical Reports Server (NTRS)

Abrahamson, Matt; Ardito, Alessandro; Han, Don; Haw, Robert; Kennedy, Brian; Mastrodemos, Nicholas; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew

2013-01-01

The NASA Dawn spacecraft was launched on September 27, 2007 on a mission to study the asteroid belt's two largest objects, Vesta and Ceres. It is the first deep space orbiting mission to demonstrate solar-electric ion propulsion, providing the necessary delta-V to enable capture and escape from two extraterrestrial bodies. At this time, Dawn has completed its science campaign at Vesta and is currently on its journey to Ceres, where it will arrive in mid-2015. The spacecraft spent over a year in orbit around Vesta from July 2011 through August 2012, capturing science data during four dedicated orbit phases. In order to maintain the reference orbits necessary for science and enable the transfers between those orbits, precise and timely orbit determination was required. The constraints associated with low-thrust ion propulsion coupled with the relatively unknown a priori gravity and rotation models for Vesta presented unique challenges for the Dawn orbit determination team. While [1] discusses the prediction performance of the orbit determination products, this paper discusses the dynamics models, filter configuration, and data processing implemented to deliver a rapid orbit determination capability to the Dawn project.

20 CFR 10.737 - How is a LEO claim filed, and who can file a LEO claim?

Code of Federal Regulations, 2012 CFR

2012-04-01

... Form CA-721; a death claim should be filed on Form CA-722. All claims should be submitted to the... 20 Employees' Benefits 1 2012-04-01 2012-04-01 false How is a LEO claim filed, and who can file a...' COMPENSATION ACT, AS AMENDED Special Provisions Non-Federal Law Enforcement Officers § 10.737 How is a LEO...

20 CFR 10.737 - How is a LEO claim filed, and who can file a LEO claim?

Code of Federal Regulations, 2010 CFR

2010-04-01

... Form CA-721; a death claim should be filed on Form CA-722. All claims should be submitted to the... 20 Employees' Benefits 1 2010-04-01 2010-04-01 false How is a LEO claim filed, and who can file a...' COMPENSATION ACT, AS AMENDED Special Provisions Non-Federal Law Enforcement Officers § 10.737 How is a LEO...

20 CFR 10.737 - How is a LEO claim filed, and who can file a LEO claim?

Code of Federal Regulations, 2011 CFR

2011-04-01

... Form CA-721; a death claim should be filed on Form CA-722. All claims should be submitted to the... 20 Employees' Benefits 1 2011-04-01 2011-04-01 false How is a LEO claim filed, and who can file a...' COMPENSATION ACT, AS AMENDED Special Provisions Non-Federal Law Enforcement Officers § 10.737 How is a LEO...

20 CFR 10.737 - How is a LEO claim filed, and who can file a LEO claim?

Code of Federal Regulations, 2014 CFR

2014-04-01

... Form CA-721; a death claim should be filed on Form CA-722. All claims should be submitted to the... 20 Employees' Benefits 1 2014-04-01 2012-04-01 true How is a LEO claim filed, and who can file a...' COMPENSATION ACT, AS AMENDED Special Provisions Non-Federal Law Enforcement Officers § 10.737 How is a LEO...

20 CFR 10.737 - How is a LEO claim filed, and who can file a LEO claim?

Code of Federal Regulations, 2013 CFR

2013-04-01

... Form CA-721; a death claim should be filed on Form CA-722. All claims should be submitted to the... 20 Employees' Benefits 1 2013-04-01 2012-04-01 true How is a LEO claim filed, and who can file a...' COMPENSATION ACT, AS AMENDED Special Provisions Non-Federal Law Enforcement Officers § 10.737 How is a LEO...

Real-Time and Post-Processed Orbit Determination and Positioning

NASA Technical Reports Server (NTRS)

Harvey, Nathaniel E. (Inventor); Lu, Wenwen (Inventor); Miller, Mark A. (Inventor); Bar-Sever, Yoaz E. (Inventor); Miller, Kevin J. (Inventor); Romans, Larry J. (Inventor); Dorsey, Angela R. (Inventor); Sibthorpe, Anthony J. (Inventor); Weiss, Jan P. (Inventor); Bertiger, William I. (Inventor);

Real-Time and Post-Processed Orbit Determination and Positioning

NASA Technical Reports Server (NTRS)

Bar-Sever, Yoaz E. (Inventor); Romans, Larry J. (Inventor); Weiss, Jan P. (Inventor); Gross, Jason (Inventor); Harvey, Nathaniel E. (Inventor); Lu, Wenwen (Inventor); Dorsey, Angela R. (Inventor); Miller, Mark A. (Inventor); Sibthorpe, Anthony J. (Inventor); Bertiger, William I. (Inventor);

Material interactions with the Low Earth Orbital (LEO) environment: Accurate reaction rate measurements

NASA Technical Reports Server (NTRS)

Visentine, James T.; Leger, Lubert J.

1987-01-01

To resolve uncertainties in estimated LEO atomic oxygen fluence and provide reaction product composition data for comparison to data obtained in ground-based simulation laboratories, a flight experiment has been proposed for the space shuttle which utilizes an ion-neutral mass spectrometer to obtain in-situ ambient density measurements and identify reaction products from modeled polymers exposed to the atomic oxygen environment. An overview of this experiment is presented and the methodology of calibrating the flight mass spectrometer in a neutral beam facility prior to its use on the space shuttle is established. The experiment, designated EOIM-3 (Evaluation of Oxygen Interactions with Materials, third series), will provide a reliable materials interaction data base for future spacecraft design and will furnish insight into the basic chemical mechanisms leading to atomic oxygen interactions with surfaces.

Method of resolving radio phase ambiguity in satellite orbit determination

NASA Technical Reports Server (NTRS)

Councelman, Charles C., III; Abbot, Richard I.

1989-01-01

For satellite orbit determination, the most accurate observable available today is microwave radio phase, which can be differenced between observing stations and between satellites to cancel both transmitter- and receiver-related errors. For maximum accuracy, the integer cycle ambiguities of the doubly differenced observations must be resolved. To perform this ambiguity resolution, a bootstrapping strategy is proposed. This strategy requires the tracking stations to have a wide ranging progression of spacings. By conventional 'integrated Doppler' processing of the observations from the most widely spaced stations, the orbits are determined well enough to permit resolution of the ambiguities for the most closely spaced stations. The resolution of these ambiguities reduces the uncertainty of the orbit determination enough to enable ambiguity resolution for more widely spaced stations, which further reduces the orbital uncertainty. In a test of this strategy with six tracking stations, both the formal and the true errors of determining Global Positioning System satellite orbits were reduced by a factor of 2.

Atmospheric density models comparison and impact on orbit solutions of GRACE-1, Sentinel-1A, TerraSAR-X

NASA Astrophysics Data System (ADS)

Colace, Marco; Hackel, Stefan; Kirschner, Michael; Kahle, Ralph; Circi, Christian

2017-04-01

Satellites in Low Earth Orbit (LEO) are notably affected by the presence of the atmosphere, a predominant source of perturbations of the Keplerian motion at the altitudes of interest. For spacecraft of this class the main source of error in propagated trajectories is due to the mismodeling of the neutral density in the thermosphere and the associated drag force, which steadily decelerates orbital motion with both secular and periodic effects. Thermospheric density varies significantly with space and time because of complex interactions between solar activity and the Earth's atmosphere and magnetic field. Properly reproducing this variability by means of empirical dynamic models has always represented a difficult task but is of vital importance for orbit determination and propagation. The present study shows the influence of different atmospheric density models, predicted space weather proxies, and their related uncertainties on the orbit solutions of representative satellite missions. The study has been carried out by using a routine-like orbit propagation scenario applied to GRACE-1, Sentinel-1A, and TerraSAR-X, three LEO orbiting spacecraft with operational altitudes well spaced within the 400-700 km range. Archived space weather data predictions and some of the most recent and promising empirical atmospheric models (Naval Research Laboratory's NRLMSISE-00 and Jacchia-Bowman 2008) were used side-by-side with the well-known Jacchia 1971 model in order to assess potential gains in prediction accuracy. To evaluate the influence of solar variability on the atmospheric density models and associated orbit quality, two 2-month test time frames, in high and low solar activity periods, have been selected. The scope of the presentation is a detailed comparison of atmospheric density models and their influence on the estimated orbits of GRACE-1, Sentinel-1A and TerraSAR-X.

[From "Løveapotekets kemiske Fabrik" to LEO Pharma A/S].

PubMed

Schrøder, Preben

2005-01-01

A chronological exposition of the history of LEO Pharma from the limited beginnings in 1908 in "København Løveapotek" to the international pharmaceutical company that exists today. The paper chronicles the various owners, from the pharmacist, August Kongsted, through his son-in-law, Knud Abildgaard to the present ownership by the LEO Foundation, which has during the entire period secured a company independent of foreign capital. Research, the life blood of the company, has resulted in a number of outstanding drugs such as hormones, heparin, diuretics, penicillin, Fucidin, alfacalcidol and calcipotriol. The company's headquarters are based in Ballerup close to Copenhagen. Further to a large research centre and production facilities in Denmark, LEO Pharma has affiliates in almost 100 countries and a number of European manufacturing sites. With 3500 employees across the globe and a yearly turnover of DKK 4.5 billion, LEO is now firmly established. Throughout the years LEO Pharma has been active within a number of other fields like agricultural feedstuff, veterinary medicine and medical utensils, but today LEO Pharma works only in human medicine. Over the years LEO Pharma has contributed to the foundation of companies like Nordisk Insulin, Ferrosan and LEO AB in Sweden.

Degradation of Silicon Carbide Reflective Surfaces in the LEO Environment

NASA Astrophysics Data System (ADS)

Mileti, Sandro; Coluzzi, Plinio; Marchetti, Mario

2009-01-01

Space mirrors in Low Earth Orbit (LEO) encounter a degradation problem caused by the impact of atomic oxygen (ATOX) in the space environment. This paper presents an experiment of the atomic oxygen impact degradation and UV synergic effects on ground simulation. The experiment was carried out in a dedicated ATOX simulation vacuum chamber. As target materials, a polished CVD Beta-silicon carbide (SiC) coating was investigated. The selection of silicon carbide is due to its high potential candidate as a mirror layer substrate material for its good reflectance at UV wavelengths and excellent thermal diffusivity. It has highly desirable mechanical and thermal properties and can achieve an excellent surface finish. The deposition of the coatings were on carbon-based material substrate; i.e., silicon impregnated carbon fiber composite (C/SiC). Mechanical and thermal properties of the coatings such as hardness and Coefficient of Thermal Expansion (CTE) were achieved. Several atomic oxygen impact angles were studied tilting the target samples respect to the flux direction. The various impact angles permitted to analyze the different erosion rates and typologies which the mirrors would encounter in LEO environment. The degradation was analyzed in various aspects. Macroscopic mass loss per unit area, surface roughness and morphology change were basically analyzed. The exposed surfaces of the materials were observed through a Scanning Electron Microscope (SEM). Secondly, optical diagnostic of the surfaces were performed in order to investigate their variation in optical properties as the evaluation of reflectance degradation. The presence of micro-cracks caused by shrinkage, grinding, polishing or thermal cycling and the porosity in the coatings, could have led to the undercutting phenomenon. Observation of uprising of undercutting was also conducted. Remarks are given regarding capabilities in short-term mission exposures to the LEO environment of this coating.

International Low-Earth-Orbit Spacecraft Materials Test Program Initiated for Better Prediction of Durability and Performance

NASA Technical Reports Server (NTRS)

Rutledge, Sharon K.

1999-01-01

Spacecraft in low Earth orbit (LEO) are subjected to many components of the environment, which can cause them to degrade much more rapidly than intended and greatly shorten their functional life. The atomic oxygen, ultraviolet radiation, and cross contamination present in LEO can affect sensitive surfaces such as thermal control paints, multilayer insulation, solar array surfaces, and optical surfaces. The LEO Spacecraft Materials Test (LEO-SMT) program is being conducted to assess the effects of simulated LEO exposure on current spacecraft materials to increase understanding of LEO degradation processes as well as to enable the prediction of in-space performance and durability. Using ground-based simulation facilities to test the durability of materials currently flying in LEO will allow researchers to compare the degradation evidenced in the ground-based facilities with that evidenced on orbit. This will allow refinement of ground laboratory test systems and the development of algorithms to predict the durability and performance of new materials in LEO from ground test results. Accurate predictions based on ground tests could reduce development costs and increase reliability. The wide variety of national and international materials being tested represent materials being functionally used on spacecraft in LEO. The more varied the types of materials tested, the greater the probability that researchers will develop and validate predictive models for spacecraft long-term performance and durability. Organizations that are currently participating in the program are ITT Research Institute (USA), Lockheed Martin (USA), MAP (France), SOREQ Nuclear Research Center (Israel), TNO Institute of Applied Physics (The Netherlands), and UBE Industries, Ltd. (Japan). These represent some of the major suppliers of thermal control and sensor materials currently flying in LEO. The participants provide materials that are exposed to selected levels of atomic oxygen, vacuum ultraviolet

On initial orbit determination

NASA Technical Reports Server (NTRS)

Taff, L. G.

1984-01-01

The classical methods of initial orbit determination are brought together within a larger viewpoint. This new synthesis stresses that all such techniques follow one of three approaches. Either they seek to compute the orbital element set, or its equivalent, by attacking the differential equations of motion (Laplace), the first integrals of the equations of motion (Taff), or the solution itself (Gauss). The particular technique pursued within a given type of approach should depend upon the nature of the observational data, the amount of a priori information one is willing to presume, and the object of the exercise. This might be a binary star system, a moon, a minor planet, or an artificial satellite. The efficacy of some algorithms for each approach is discussed briefly. Unfortunately, none of them work very well. Extensions of these techniques to radars or laser radars are trivial and have provided no new insights into the overall problem.

A simplex method for the orbit determination of maneuvering satellites

NASA Astrophysics Data System (ADS)

Chen, JianRong; Li, JunFeng; Wang, XiJing; Zhu, Jun; Wang, DanNa

2018-02-01

A simplex method of orbit determination (SMOD) is presented to solve the problem of orbit determination for maneuvering satellites subject to small and continuous thrust. The objective function is established as the sum of the nth powers of the observation errors based on global positioning satellite (GPS) data. The convergence behavior of the proposed method is analyzed using a range of initial orbital parameter errors and n values to ensure the rapid and accurate convergence of the SMOD. For an uncontrolled satellite, the orbit obtained by the SMOD provides a position error compared with GPS data that is commensurate with that obtained by the least squares technique. For low Earth orbit satellite control, the precision of the acceleration produced by a small pulse thrust is less than 0.1% compared with the calibrated value. The orbit obtained by the SMOD is also compared with weak GPS data for a geostationary Earth orbit satellite over several days. The results show that the position accuracy is within 12.0 m. The working efficiency of the electric propulsion is about 67% compared with the designed value. The analyses provide the guidance for subsequent satellite control. The method is suitable for orbit determination of maneuvering satellites subject to small and continuous thrust.

The Orbital Debris Problem and the Challenges for Environment Remediation

NASA Technical Reports Server (NTRS)

Liou, J.-C.

2013-01-01

Orbital debris scientists from major international space agencies, including JAXA and NASA, have worked together to predict the trend of the future environment. A summary presentation was given to the United Nations in February 2013. The orbital debris population in LEO will continue to increase. Catastrophic collisions will continue to occur every 5 to 9 years center dot To limit the growth of the future debris population and to better protect future spacecraft, active debris removal, should be considered.

Evaluation of semiempirical atmospheric density models for orbit determination applications

NASA Technical Reports Server (NTRS)

Cox, C. M.; Feiertag, R. J.; Oza, D. H.; Doll, C. E.

1994-01-01

This paper presents the results of an investigation of the orbit determination performance of the Jacchia-Roberts (JR), mass spectrometer incoherent scatter 1986 (MSIS-86), and drag temperature model (DTM) atmospheric density models. Evaluation of the models was performed to assess the modeling of the total atmospheric density. This study was made generic by using six spacecraft and selecting time periods of study representative of all portions of the 11-year cycle. Performance of the models was measured for multiple spacecraft, representing a selection of orbit geometries from near-equatorial to polar inclinations and altitudes from 400 kilometers to 900 kilometers. The orbit geometries represent typical low earth-orbiting spacecraft supported by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD). The best available modeling and orbit determination techniques using the Goddard Trajectory Determination System (GTDS) were employed to minimize the effects of modeling errors. The latest geopotential model available during the analysis, the Goddard earth model-T3 (GEM-T3), was employed to minimize geopotential model error effects on the drag estimation. Improved-accuracy techniques identified for TOPEX/Poseidon orbit determination analysis were used to improve the Tracking and Data Relay Satellite System (TDRSS)-based orbit determination used for most of the spacecraft chosen for this analysis. This paper shows that during periods of relatively quiet solar flux and geomagnetic activity near the solar minimum, the choice of atmospheric density model used for orbit determination is relatively inconsequential. During typical solar flux conditions near the solar maximum, the differences between the JR, DTM, and MSIS-86 models begin to become apparent. Time periods of extreme solar activity, those in which the daily and 81-day mean solar flux are high and change rapidly, result in significant differences between the models. During periods of high

Multi-GNSS orbit determination using satellite laser ranging

NASA Astrophysics Data System (ADS)

Bury, Grzegorz; Sośnica, Krzysztof; Zajdel, Radosław

2018-04-01

Galileo, BeiDou, QZSS, and NavIC are emerging global navigation satellite systems (GNSSs) and regional navigation satellite systems all of which are equipped with laser retroreflector arrays for range measurements. This paper summarizes the GNSS-intensive tracking campaigns conducted by the International Laser Ranging Service and provides results from multi-GNSS orbit determination using solely SLR observations. We consider the whole constellation of GLONASS, all active Galileo, four BeiDou satellites: 1 MEO, 3 IGSO, and one QZSS. We analyze the influence of the number of SLR observations on the quality of the 3-day multi-GNSS orbit solution. About 60 SLR observations are needed for obtaining MEO orbits of sufficient quality with the root mean square (RMS) of 3 cm for the radial component when compared to microwave-based orbits. From the analysis of a minimum number of tracking stations, when considering the 3-day arcs, 5 SLR stations do not provide a sufficient geometry of observations. The solution obtained using ten stations is characterized with RMS of 4, 9, and 18 cm in the radial, along-track, and cross-track direction, respectively, for MEO satellites. We also investigate the impact of the length of orbital arc on the quality of SLR-derived orbits. Hence, 5- and 7-day arcs constitute the best solution, whereas 3-day arcs are of inferior quality due to an insufficient number of SLR observations and 9-day arcs deteriorate the along-track component. The median RMS from the comparison between 7-day orbital arcs determined using SLR data with microwave-based orbits assumes values in the range of 3-4, 11-16, and 15-27 cm in radial, along-track, and cross-track, respectively, for MEO satellites. BeiDou IGSO and QZSS are characterized by RMS values higher by a factor of 8 and 24, respectively, than MEO orbits.

Space radiation dosimetry in low-Earth orbit and beyond.

PubMed

Benton, E R; Benton, E V

2001-09-01

Space radiation dosimetry presents one of the greatest challenges in the discipline of radiation protection. This is a result of both the highly complex nature of the radiation fields encountered in low-Earth orbit (LEO) and interplanetary space and of the constraints imposed by spaceflight on instrument design. This paper reviews the sources and composition of the space radiation environment in LEO as well as beyond the Earth's magnetosphere. A review of much of the dosimetric data that have been gathered over the last four decades of human space flight is presented. The different factors affecting the radiation exposures of astronauts and cosmonauts aboard the International Space Station (ISS) are emphasized. Measurements made aboard the Mir Orbital Station have highlighted the importance of both secondary particle production within the structure of spacecraft and the effect of shielding on both crew dose and dose equivalent. Roughly half the dose on ISS is expected to come from trapped protons and half from galactic cosmic rays (GCRs). The dearth of neutron measurements aboard LEO spacecraft and the difficulty inherent in making such measurements have led to large uncertainties in estimates of the neutron contribution to total dose equivalent. Except for a limited number of measurements made aboard the Apollo lunar missions, no crew dosimetry has been conducted beyond the Earth's magnetosphere. At the present time we are forced to rely on model-based estimates of crew dose and dose equivalent when planning for interplanetary missions, such as a mission to Mars. While space crews in LEO are unlikely to exceed the exposure limits recommended by such groups as the NCRP, dose equivalents of the same order as the recommended limits are likely over the course of a human mission to Mars. c2001 Elsevier Science B.V. All rights reserved.

View of ANDE release from orbiter Discovery payload bay

NASA Image and Video Library

2006-12-21

S116-E-07828 (21 Dec. 2006) --- As seen through windows on the aft flight deck of Space Shuttle Discovery, a Department of Defense pico-satellite known as Atmospheric Neutral Density Experiment (ANDE) is released from the shuttle's payload bay by STS-116 crewmembers. ANDE consists of two micro-satellites which will measure the density and composition of the low Earth orbit (LEO) atmosphere while being tracked from the ground. The data will be used to better predict the movement of objects in orbit.

Ground-based simulation of LEO environment: Investigations of a select LDEF material: FEP Teflon (trademark)

NASA Technical Reports Server (NTRS)

Cross, Jon B.; Koontz, Steven L.

1993-01-01

The Long Duration Exposure Facility (LDEF) has produced a wealth of data on materials degradation in the low earth orbit (LEO) space environment and has conclusively shown that surface chemistry (as opposed to surface physics-sputtering) is the key to understanding and predicting the degradation of materials in the LEO environment. It is also clear that materials degradation and spacecraft contamination are closely linked and that the fundamental mechanisms responsible for this linking are in general not well understood especially in the area of synergistic effects. The study of the fundamental mechanisms underlying materials degradation in LEO is hampered by the fact that the degradation process itself is not observed during the actual exposure to the environment. Rather the aftermath of the degradation process is studied, i.e., the material that remains after exposure is observed and mechanisms are proposed to explain the observed results. The EOIM-3 flight experiment is an attempt to bring sophisticated diagnostic equipment into the space environment and monitor the degradation process in real time through the use of mass spectrometry. More experiments of this nature which would include surface sensitive diagnostics (Auger and photoelectron spectroscopes) are needed to truly unravel the basic chemical mechanisms involved in the materials degradation process. Since these in-space capabilities will most likely not be available in the near future, ground-based LEO simulation facilities employing sophisticated diagnostics are needed to further advance the basic understanding of the materials degradation mechanisms. The LEO simulation facility developed at Los Alamos National Laboratory has been used to investigate the atomic oxygen/vacuum ultraviolet (AO/VUV) enhanced degradation of FEP Teflon. The results show that photo-ejection of polymer fragments occur at elevated temperature (200 C), that VUV synergistic rare gas sputtering of polymer fragments occur even at

Failure analysis of satellite subsystems to define suitable de-orbit devices

NASA Astrophysics Data System (ADS)

Palla, Chiara; Peroni, Moreno; Kingston, Jennifer

2016-11-01

Space missions in Low Earth Orbit (LEO) are severely affected by the build-up of orbital debris. A key practice, to be compliant with IADC (Inter-Agency Space Debris Coordination Committee) mitigation guidelines, is the removal of space systems that interfere with the LEO region not later than 25 years after the End of Mission. It is important to note that the current guidelines are not generally legally binding, even if different Space Agencies are now looking at the compliance for their missions. If the guidelines will change in law, it will be mandatory to have a postmission disposal strategy for all satellites, including micro and smaller classes. A potential increased number of these satellites is confirmed by different projections, in particular in the commercial sector. Micro and smaller spacecraft are, in general, not provided with propulsion capabilities to achieve a controlled re-entry, so they need different de-orbit disposal methods. When considering the utility of different debris mitigation methods, it is useful to understand which spacecraft subsystems are most likely to fail and how this may affect the operation of a de-orbit system. This also helps the consideration of which components are the most relevant or should be redundant depending on the satellite mass class. This work is based on a sample of LEO and MEO satellites launched between January 2000 and December 2014 with mass lower than 1000 kg. Failure analysis of satellite subsystems is performed by means of the Kaplan-Meier survival analysis; the parametric fits are conducted with Weibull distributions. The study is carried out by using the satellite database SpaceTrak™ which provides anomalies, failures, and trends information for spacecraft subsystems and launch vehicles. The database identifies five states for each satellite subsystem: three degraded states, one fully operational state, and one failed state (complete failure). The results obtained can guide the identification of the

Orbit Determination Toolbox

NASA Technical Reports Server (NTRS)

Carpenter, James R.; Berry, Kevin; Gregpru. Late; Speckman, Keith; Hur-Diaz, Sun; Surka, Derek; Gaylor, Dave

2010-01-01

The Orbit Determination Toolbox is an orbit determination (OD) analysis tool based on MATLAB and Java that provides a flexible way to do early mission analysis. The toolbox is primarily intended for advanced mission analysis such as might be performed in concept exploration, proposal, early design phase, or rapid design center environments. The emphasis is on flexibility, but it has enough fidelity to produce credible results. Insight into all flight dynamics source code is provided. MATLAB is the primary user interface and is used for piecing together measurement and dynamic models. The Java Astrodynamics Toolbox is used as an engine for things that might be slow or inefficient in MATLAB, such as high-fidelity trajectory propagation, lunar and planetary ephemeris look-ups, precession, nutation, polar motion calculations, ephemeris file parsing, and the like. The primary analysis functions are sequential filter/smoother and batch least-squares commands that incorporate Monte-Carlo data simulation, linear covariance analysis, measurement processing, and plotting capabilities at the generic level. These functions have a user interface that is based on that of the MATLAB ODE suite. To perform a specific analysis, users write MATLAB functions that implement truth and design system models. The user provides his or her models as inputs to the filter commands. The software provides a capability to publish and subscribe to a software bus that is compliant with the NASA Goddard Mission Services Evolution Center (GMSEC) standards, to exchange data with other flight dynamics tools to simplify the flight dynamics design cycle. Using the publish and subscribe approach allows for analysts in a rapid design center environment to seamlessly incorporate changes in spacecraft and mission design into navigation analysis and vice versa.

Technology development, demonstration, and orbital support requirements for manned lunar and Mars missions

NASA Technical Reports Server (NTRS)

Llewellyn, Charles P.; Brender, Karen D.

1990-01-01

An overview of the critical technology needs and the Space Station Freedom (SSF) focused support requirements for the Office of Exploration's (OEXP) manned lunar and Mars missions is presented. Major emphasis is directed at the technology needs associated with the low earth orbit (LEO) transportation node assembly and vehicle processing functions required by the lunar and Mars mission flight elements. The key technology areas identified as crucial to support the LEO node function include in-space assembly and construction, in-space vehicle processing and refurbishment, space storable cryogenics, and autonomous rendezvous and docking.

Technology needs development and orbital support requirements for manned lunar and Mars missions

NASA Technical Reports Server (NTRS)

Brender, Karen D.; Llewellyn, Charles P.

1990-01-01

This paper presents an overview of the critical technology needs and the Space Station Freedom focused support requirements for the Office of Exploration's manned lunar and Mars missions. The emphasis is on e directed at the technology needs associated with the low earth orbit (LEO) transportation node assembly and vehicle processing functions required by the lunar Mars mission flight elements. The key technology areas identified as crucial to support the LEO node function include in-space assembly and construction, in-space vehicle processing and refurbishment, space storable cryogenics, and autonomous rendezvous and docking.

A Brief History of Meteoroid and Orbital Debris Shielding Technology for US Manned Spacecraft

NASA Technical Reports Server (NTRS)

Bjorkman, Michael D.; Hyde, James L.

2008-01-01

Meteoroid and orbital debris shielding has played an important role from the beginning of manned spaceflight. During the early 60 s, meteoroid protection drove requirements for new meteor and micrometeoroid impact science. Meteoroid protection also stimulated advances in the technology of hypervelocity impact launchers and impact damage assessment methodologies. The first phase of meteoroid shielding assessments closed in the early 70 s with the end of the Apollo program. The second phase of meteoroid protection technology began in the early 80 s when it was determined that there is a manmade Earth orbital debris belt that poses a significant risk to LEO manned spacecraft. The severity of the Earth orbital debris environment has dictated changes in Space Shuttle and ISS operations as well as driven advances in shielding technology and assessment methodologies. A timeline of shielding technology and assessment methodology advances is presented along with a summary of risk assessment results.

Use of the VLBI delay observable for orbit determination of Earth-orbiting VLBI satellites

NASA Technical Reports Server (NTRS)

Ulvestad, J. S.

1992-01-01

Very long-baseline interferometry (VLBI) observations using a radio telescope in Earth orbit were performed first in the 1980s. Two spacecraft dedicated to VLBI are scheduled for launch in 1995; the primary scientific goals of these missions will be astrophysical in nature. This article addresses the use of space VLBI delay data for the additional purpose of improving the orbit determination of the Earth-orbiting spacecraft. In an idealized case of quasi-simultaneous observations of three radio sources in orthogonal directions, analytical expressions are found for the instantaneous spacecraft position and its error. The typical position error is at least as large as the distance corresponding to the delay measurement accuracy but can be much greater for some geometries. A number of practical considerations, such as system noise and imperfect calibrations, set bounds on the orbit-determination accuracy realistically achievable using space VLBI delay data. These effects limit the spacecraft position accuracy to at least 35 cm (and probably 3 m or more) for the first generation of dedicated space VLBI experiments. Even a 35-cm orbital accuracy would fail to provide global VLBI astrometry as accurate as ground-only VLBI. Recommended charges in future space VLBI missions are unlikely to make space VLBI competitive with ground-only VLBI in global astrometric measurements.

Space environmental effects on spacecraft: LEO materials selection guide, part 1

NASA Astrophysics Data System (ADS)

Silverman, Edward M.

1995-08-01

This document provides performance properties on major spacecraft materials and subsystems that have been exposed to the low-Earth orbit (LEO) space environment. Spacecraft materials include metals, polymers, composites, white and black paints, thermal-control blankets, adhesives, and lubricants. Spacecraft subsystems include optical components, solar cells, and electronics. Information has been compiled from LEO short-term spaceflight experiments (e.g., space shuttle) and from retrieved satellites of longer mission durations (e.g., Long Duration Exposure Facility). Major space environment effects include atomic oxygen (AO), ultraviolet radiation, micrometeoroids and debris, contamination, and particle radiation. The main objective of this document is to provide a decision tool to designers for designing spacecraft and structures. This document identifies the space environments that will affect the performance of materials and components, e.g., thermal-optical property changes of paints due to UV exposures, AO-induced surface erosion of composites, dimensional changes due to thermal cycling, vacuum-induced moisture outgassing, and surface optical changes due to AO/UV exposures. Where appropriate, relationships between the space environment and the attendant material/system effects are identified. Part 1 covers spacecraft design considerations for the space environment; advanced composites; polymers; adhesives; metals; ceramics; protective coatings; and lubricants, greases, and seals.

Space environmental effects on spacecraft: LEO materials selection guide, part 1

NASA Technical Reports Server (NTRS)

Silverman, Edward M.

1995-01-01

This document provides performance properties on major spacecraft materials and subsystems that have been exposed to the low-Earth orbit (LEO) space environment. Spacecraft materials include metals, polymers, composites, white and black paints, thermal-control blankets, adhesives, and lubricants. Spacecraft subsystems include optical components, solar cells, and electronics. Information has been compiled from LEO short-term spaceflight experiments (e.g., space shuttle) and from retrieved satellites of longer mission durations (e.g., Long Duration Exposure Facility). Major space environment effects include atomic oxygen (AO), ultraviolet radiation, micrometeoroids and debris, contamination, and particle radiation. The main objective of this document is to provide a decision tool to designers for designing spacecraft and structures. This document identifies the space environments that will affect the performance of materials and components, e.g., thermal-optical property changes of paints due to UV exposures, AO-induced surface erosion of composites, dimensional changes due to thermal cycling, vacuum-induced moisture outgassing, and surface optical changes due to AO/UV exposures. Where appropriate, relationships between the space environment and the attendant material/system effects are identified. Part 1 covers spacecraft design considerations for the space environment; advanced composites; polymers; adhesives; metals; ceramics; protective coatings; and lubricants, greases, and seals.

EcAMSat and BioSentinel: Autonomous Bio Nanosatellites Addressing Strategic Knowledge Gaps for Manned Spaceflight Beyond LEO

NASA Technical Reports Server (NTRS)

Padgen, Michael R.

2017-01-01

Manned missions beyond low Earth orbit (LEO) require that several strategic knowledge gaps about the effects of space travel on the human body be addressed. NASA Ames Research Center has been the leader in developing autonomous bio nanosatellites, including past successful missions for GeneSat, PharmaSat, and O/OREOS, that tackled some of these issues. These nanosatellites provide in situ measurements, which deliver insight into the dynamic changes in cell behavior in microgravity. In this talk, two upcoming bio nanosatellites developed at Ames, the E. coli Antimicrobial Satellite (EcAMSat) and BioSentinel, will be discussed. Both satellites contain microfluidic systems that precisely deliver nutrients to the microorganisms stored within wells of fluidic cards. Each well, in turn, has its own 3-color LED and detector system which is used to monitor changes in metabolic activity with alamarBlue, a redox indicator, and the optical density of the cells. EcAMSat investigates the effects of microgravity on bacterial resistance to antimicrobial drugs, vital knowledge for understanding how to maintain the health of astronauts in long-term and beyond LEO spaceflight. The behavior of wild type and mutant uropathic E. coli will be compared in microgravity and with ground data to help understand the molecular mechanisms behind antibiotic resistance and how these phenotypes might change in space. BioSentinel seeks to directly measure the effects of space radiation on budding yeast S. cerevisiae, particularly double strand breaks (DSB). While hitching a ride on the SLS EM-1 mission (Orion's first unmanned mission to the moon) in 2018, BioSentinel will be kicked off and enter into a heliocentric orbit, becoming the first study of the effects of radiation on living organisms outside LEO since the Apollo program. The yeast are stored in eighteen independent 16-well microfluidic cards, which will be individually activated over the 12 month mission duration. In addition to the wild

EcAMSat and BioSentinel: Autonomous Bio Nanosatellites Addressing Strategic Knowledge Gaps for Manned Spaceflight Beyond LEO

NASA Technical Reports Server (NTRS)

Padgen, Mike

2017-01-01

Manned missions beyond low Earth orbit (LEO) require that several strategic knowledge gaps about the effects of space travel on the human body be addressed. NASA Ames Research Center has been the leader in developing autonomous bio nanosatellites, including past successful missions for GeneSat, PharmaSat, and OOREOS, that tackled some of these issues. These nanosatellites provide in situ measurements, which deliver insight into the dynamic changes in cell behavior in microgravity. In this talk, two upcoming bio nanosatellites developed at Ames, the E. coli Antimicrobial Satellite (EcAMSat) and BioSentinel, will be discussed. Both satellites contain microfluidic systems that precisely deliver nutrients to the microorganisms stored within wells of fluidic cards. Each well, in turn, has its own 3-color LED and detector system which is used to monitor changes in metabolic activity with alamarBlue, a redox indicator, and the optical density of the cells. EcAMSat investigates the effects of microgravity on bacterial resistance to antimicrobial drugs, vital knowledge for understanding how to maintain the health of astronauts in long-term and beyond LEO spaceflight. The behavior of wild type and mutant uropathic E. coli will be compared in microgravity and with ground data to help understand the molecular mechanisms behind antibiotic resistance and how these phenotypes might change in space. BioSentinel seeks to directly measure the effects of space radiation on budding yeast S. cerevisiae, particularly double strand breaks (DSB). While hitching a ride on the SLS EM-1 mission (Orions first unmanned mission to the moon) in 2018, BioSentinel will be kicked off and enter into a heliocentric orbit, becoming the first study of the effects of radiation on living organisms outside LEO since the Apollo program. The yeast are stored in eighteen independent 16-well microfluidic cards, which will be individually activated over the 12 month mission duration. In addition to the wild

Characterization of the Catalog Fengyun-1C Fragments and Their Long-term Effect on the LEO Environment

NASA Technical Reports Server (NTRS)

Liou, J.-C.; Johnson, N. L.

2008-01-01

The intentional breakup of Fengyun-1C on 11 January 2007 created the most severe orbital debris cloud in history. More than 2500 large fragments were identified and tracked by the U.S. Space Surveillance Network by the end of the year. The altitude where the event occurred was probably the worst location for a major breakup in the low Earth orbit (LEO) region, since it was already highly populated with operational satellites and debris generated from previous breakups. The addition of so many fragments not only poses a realistic threat to operational satellites in the region, but also increases the instability (i.e., collision cascade effect) of the debris population there. Preliminary analysis of the large Fengyun-1C fragments indicates that their size and area-to-mass ratio (A/M) distributions are very different from those of other known events. About half of the fragments appear to be composed of light-weight materials and more than 100 of them have A/M values exceeding 1 square meter per kilogram, consistent with thermal blanket pieces. In addition, the orbital elements of the fragments suggest nontrivial velocity gain by the fragment cloud during the impact. These important characteristics were incorporated into a numerical simulation to assess the long-term impact of the Fengyun-1C fragments to the LEO debris environment. The main objectives of the simulation were to evaluate (1) the collision probabilities between the Fengyun-1C fragments and the rest of the catalog population and (2) the collision activities and population growth in the region in the next 100 years.

Characterization of the catalog Fengyun-1C fragments and their long-term effect on the LEO environment

NASA Astrophysics Data System (ADS)

Liou, J.-C.

The intentional breakup of Fengyun-1C on 11 January 2007 created the most severe orbital debris cloud in history. More than 2500 large fragments were identified and tracked by the U.S. Space Surveillance Network by the end of the year. The altitude where the event occurred was probably the worst location for a major breakup in the low Earth orbit (LEO) region, since it was already highly populated with operational satellites and debris generated from previous breakups. The addition of so many fragments not only poses a realistic threat to operational satellites in the region, but also increases the instability (i.e., collision cascade effect) of the debris population there. Preliminary analysis of the large Fengyun-1C fragments indicates that their size and area-tomass ratio (A/M) distributions are very different from those of other known events. About half of the fragments appear to be composed of light-weight materials and more than 100 of them have A/M values exceeding 1 m2 /kg, consistent with thermal blanket pieces. In addition, the orbital elements of the fragments suggest non-trivial velocity gain by the fragment cloud during the impact. These important characteristics were incorporated into a numerical simulation to assess the long-term impact of the Fengyun-1C fragments to the LEO debris environment. The main objectives of the simulation were to evaluate (1) the collision probabilities between the Fengyun-1C fragments and the rest of the catalog population and (2) the collision activities and population growth in the region in the next 100 years.

Estimated performance and future potential of solar dynamic and photovoltaic power systems for selected LEO and HEO missions

NASA Technical Reports Server (NTRS)

Bents, David J.; Lu, Cheng Y.

1989-01-01

Solar photovoltaic and thermal dynamic power systems for application to selected low-earth-orbit (LEO) and high-earth-orbit (HEO) missions are characterized in the regime 7 to 35 kWe. Input parameters to the characterization are varied to correspond to anticipated introduction of improved or new technologies. A comparative assessment is made of the two power system types for emerging technologies in cells and arrays, energy storage, optical surfaces, heat engines, thermal energy storage and thermal management. The assessment is made to common ground rules and assumptions. The four missions (Space Station, sun-synchronous, Van Allen belt, and GEO) are representative of the anticipated range of multikilowatt earth-orbit missions. The results give the expected performance, mass and drag of multikilowatt earth-orbiting solar power systems and show how the overall system figure of merit will improve as new component technologies are incorporated.

A laser-optical system to re-enter or lower low Earth orbit space debris

NASA Astrophysics Data System (ADS)

Phipps, Claude R.

2014-01-01

Collisions among existing Low Earth Orbit (LEO) debris are now a main source of new debris, threatening future use of LEO space. Due to their greater number, small (1-10 cm) debris are the main threat, while large (>10 cm) objects are the main source of new debris. Flying up and interacting with each large object is inefficient due to the energy cost of orbit plane changes, and quite expensive per object removed. Strategically, it is imperative to remove both small and large debris. Laser-Orbital-Debris-Removal (LODR), is the only solution that can address both large and small debris. In this paper, we briefly review ground-based LODR, and discuss how a polar location can dramatically increase its effectiveness for the important class of sun-synchronous orbit (SSO) objects. With 20% clear weather, a laser-optical system at either pole could lower the 8-ton ENVISAT by 40 km in about 8 weeks, reducing the hazard it represents by a factor of four. We also discuss the advantages and disadvantages of a space-based LODR system. We estimate cost per object removed for these systems. International cooperation is essential for designing, building and operating any such system.

Dawn Orbit Determination Team: Trajectory Modeling and Reconstruction Processes at Vesta

NASA Technical Reports Server (NTRS)

Abrahamson, Matthew J.; Ardito, Alessandro; Han, Dongsuk; Haw, Robert; Kennedy, Brian; Mastrodemos, Nick; Nandi, Sumita; Park, Ryan; Rush, Brian; Vaughan, Andrew

2013-01-01

The Dawn spacecraft spent over a year in orbit around Vesta from July 2011 through August 2012. In order to maintain the designated science reference orbits and enable the transfers between those orbits, precise and timely orbit determination was required. Challenges included low-thrust ion propulsion modeling, estimation of relatively unknown Vesta gravity and rotation models, track-ing data limitations, incorporation of real-time telemetry into dynamics model updates, and rapid maneuver design cycles during transfers. This paper discusses the dynamics models, filter configuration, and data processing implemented to deliver a rapid orbit determination capability to the Dawn project.

Dose in critical body organs in low Earth orbit

NASA Technical Reports Server (NTRS)

Wilson, J. W.; Cucinotta, F.

1984-01-01

Human exposure to trapped radiations in low Earth orbit (LEO) are evaluated on the basis of a simple approximation of the human geometry for spherical shell shields of varying thickness. A data base is presented that may be used to make preliminary assessment of the impact of radiation exposure constraints on human performance. A sample impact assessment is discussed.

Investigating fundamental physics and space environment with a dedicated Earth-orbiting spacecraft

NASA Astrophysics Data System (ADS)

Peron, Roberto

The near-Earth environment is a place of first choice for performing fundamental physics experiments, given its proximity to Earth and at the same time being relatively quiet dynamically for particular orbital arrangements. This environment also sees a rich phenomenology for what concerns gravitation. In fact, the general theory of relativity is an incredibly accurate description of gravitational phenomenology. However, its overall validity is being questioned by the theories that aim at reconciling it with the microscopic domain. Challenges come also from the ‘mysteries’ of Dark Matter and Dark Energy, though mainly at scales from the galactic up to the cosmological. It is therefore important to precisely test the consequences of the theory -- as well as those of competing ones -- at all the accessible scales. At the same time, the development of high-precision experimental space techniques, which are needed for tests in fundamental physics, opens the way to complementary applications. The growth of the (man-made) orbital debris population is creating problems to the future development of space. The year 2009 witnessed the first accidental collision between two satellites in orbit (Iridium and Cosmos) that led to the creation of more debris. International and national agencies are intervening by issuing and/or adopting guidelines to mitigate the growth of orbital debris. A central tenet of these guidelines requires a presence in space shorter than 25 years to satellites in low Earth orbit (LEO) after the conclusion of their operational lives. However, the determination of the natural lifetime of a satellite in LEO is very uncertain due to a large extent to the short-term and long-term variability of the atmospheric density in LEO and the comparatively low-accuracy of atmospheric density models. Many satellites orbiting in the 500-1200 km region with circular or elliptical orbits will be hard pressed to establish before flight whether or not they meet the 25

Preliminary orbit determination for lunar satellites.

NASA Technical Reports Server (NTRS)

Lancaster, E. R.

1973-01-01

Methods for the determination of orbits of artificial lunar satellites from earth-based range rate measurements developed by Koskela (1964) and Bateman et al. (1966) are simplified and extended to include range measurements along with range rate measurements. For illustration, a numerical example is presented.

Electrochemical Energy Storage for an Orbiting Space Station

NASA Technical Reports Server (NTRS)

Martin, R. E.

1981-01-01

The system weight of a multi hundred kilowatt fuel cell electrolysis cell energy storage system based upon alkaline electrochemical cell technology for use in a future orbiting space station in low Earth orbit (LEO) was studied. Preliminary system conceptual design, fuel cell module performance characteristics, subsystem and system weights, and overall system efficiency are identified. The impact of fuel cell module operating temperature and efficiency upon energy storage system weight is investigated. The weight of an advanced technology system featuring high strength filament wound reactant tanks and a fuel cell module employing lightweight graphite electrolyte reservoir plates is defined.

Low earth orbit satellite/terrestrial mobile service compatibility

NASA Technical Reports Server (NTRS)

Sheriff, R. E.; Gardiner, J. G.

1993-01-01

Digital cellular mobile 'second generation' systems are now gradually being introduced into service; one such example is GSM, which will provide a digital voice and data service throughout Europe. Total coverage is not expected to be achieved until the mid '90's, which has resulted in several proposals for the integration of GSM with a geostationary satellite service. Unfortunately, because terrestrial and space systems have been designed to optimize their performance for their particular environment, integration between a satellite and terrestrial system is unlikely to develop further than the satellite providing a back-up service. This lack of system compatibility is now being addressed by system designers of third generation systems. The next generation of mobile systems, referred to as FPLMTS (future public land mobile telecommunication systems) by CCIR and UMTS (universal mobile telecommunication system) in European research programs, are intended to provide inexpensive, hand-held terminals that can operate in either satellite, cellular, or cordless environments. This poses several challenges for system designers, not least in terms of the choice of multiple access technique and power requirements. Satellite mobile services have been dominated by the geostationary orbital type. Recently, however, a number of low earth orbit configurations have been proposed, for example Iridium. These systems are likely to be fully operational by the turn of the century, in time for the implementation of FPLMTS. The developments in LEO mobile satellite service technology were recognized at WARC-92 with the allocation of specific frequency bands for 'big' LEO's, as well as a frequency allocation for FPLMTS which included a specific satellite allocation. When considering integrating a space service into the terrestrial network, LEO's certainly appear to have their attractions: they can provide global coverage, the round trip delay is of the order of tens of milliseconds, and

The chemical abundances of the stellar populations in the Leo I and II dSph galaxies

NASA Astrophysics Data System (ADS)

Bosler, Tammy L.; Smecker-Hane, Tammy A.; Stetson, Peter B.

2007-06-01

We have obtained calcium abundances and radial velocities for 102 red giant branch (RGB) stars in the Leo I dwarf spheroidal galaxy (dSph) and 74 RGB stars in the Leo II dSph using the low-resolution spectrograph (LRIS) on the Keck I 10-m telescope. We report on the calcium abundances [Ca/H] derived from the strengths of the CaII triplet absorption lines at 8498, 8542 and 8662 Å in the stellar spectra using a new empirical CaII triplet calibration to [Ca/H]. The two galaxies have different average [Ca/H] values of -1.34 +/- 0.02 for Leo I and -1.65 +/- 0.02 for Leo II with intrinsic abundance dispersions of 1.2 and 1.0 dex, respectively. The typical random and total errors in derived abundances are 0.10 and 0.17 dex per star. For comparison to the existing literature, we also converted our CaII measurements to [Fe/H] on the scale of Carretta and Gratton (1997) though we discuss why this may not be the best determinant of metallicity; Leo I has a mean [Fe/H] = -1.34 and Leo II has a mean [Fe/H] = -1.59. The metallicity distribution function of Leo I is approximately Gaussian in shape with an excess at the metal-rich end, while that of Leo II shows an abrupt cut-off at the metal-rich end. The lower mean metallicity of Leo II is consistent with the fact that it has a lower luminosity, hence lower the total mass than Leo I; thus, the evolution of Leo II may have been affected more by mass lost in galactic winds. Our direct and independent measurement of the metallicity distributions in these dSph will allow a more accurate star-formation histories to be derived from future analysis of their colour-magnitude diagrams(CMDs). Data presented herein were obtained at the W.M. Keck Observatory, which is operated as a scientific partnership among the California Institute of Technology, the University of California and the National Aeronautics and Space Administration. The Observatory was made possible by the generous financial support of the W. M. Keck Foundation. E

TDRSS-user orbit determination using batch least-squares and sequential methods

NASA Astrophysics Data System (ADS)

Oza, D. H.; Jones, T. L.; Hakimi, M.; Samii, Mina V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.

1993-02-01

The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), and operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were less than 40 meters after the filter had reached steady state.

TDRSS-user orbit determination using batch least-squares and sequential methods

NASA Technical Reports Server (NTRS)

Oza, D. H.; Jones, T. L.; Hakimi, M.; Samii, Mina V.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.

1993-01-01

The Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) commissioned Applied Technology Associates, Incorporated, to develop the Real-Time Orbit Determination/Enhanced (RTOD/E) system on a Disk Operating System (DOS)-based personal computer (PC) as a prototype system for sequential orbit determination of spacecraft. This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using RTOD/E, operating on a PC, with the accuracy of an established batch least-squares system, the Goddard Trajectory Determination System (GTDS), and operating on a mainframe computer. The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and covariances and the first measurement residuals for the sequential case) of solutions produced by the batch and sequential methods. The forward-filtered RTOD/E orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were less than 40 meters after the filter had reached steady state.

Efficient Trajectory Propagation for Orbit Determination Problems

NASA Technical Reports Server (NTRS)

Roa, Javier; Pelaez, Jesus

2015-01-01

Regularized formulations of orbital motion apply a series of techniques to improve the numerical integration of the orbit. Despite their advantages and potential applications little attention has been paid to the propagation of the partial derivatives of the corresponding set of elements or coordinates, required in many orbit-determination scenarios and optimization problems. This paper fills this gap by presenting the general procedure for integrating the state-transition matrix of the system together with the nominal trajectory using regularized formulations and different sets of elements. The main difficulty comes from introducing an independent variable different from time, because the solution needs to be synchronized. The correction of the time delay is treated from a generic perspective not focused on any particular formulation. The synchronization using time-elements is also discussed. Numerical examples include strongly-perturbed orbits in the Pluto system, motivated by the recent flyby of the New Horizons spacecraft, together with a geocentric flyby of the NEAR spacecraft.

NiH2 Battery Reconditioning for LEO Applications

NASA Technical Reports Server (NTRS)

Armantrout, J. D.; Hafen, D. P.

1997-01-01

This paper summarizes reasons for and benefits of reconditioning nickel-hydrogen (NiH2) batteries used for Low Earth Orbit (LEO) applications. NiH2 battery cells do not have the classic discharge voltage problems more commonly associated with nickel-cadmium (NiCd) cells. This is due, in part, to use of hydrogen electrodes in place of cadmium electrodes. The nickel electrode, however, does have a similar discharge voltage signature for both cell designs. This can have an impact on LEO applications where peak loads at higher relative depths of discharge can impact operations. Periodic reconditioning provides information which can be used for analyzing long term performance trends to predict usable capacity to a specified voltage level. The reconditioning process described herein involves discharging NiH2 batteries at C/20 rates or less, to an average cell voltage of 1.0 volts or less. Recharge is performed at nominal C/5 rates to specified voltage/temperature (V/T) charge levels selected to restore required capacity with minimal overcharge. Reconditioning is a process of restoring reserve capacity lost on cycling, which is commonly called the memory effect in NiCd cells. This effect is characterized by decreases in the discharge voltage curve with operational life and cycling. The end effect of reconditioning NiH2 cells may be hidden in the versatility, of that design over the NiCd cell design and its associated negative electrode fading problem. The process of deep discharge at lower rates by way of reconditioning tends to redistribute electrolyte and water in the NiH2 cell electrode stack, while improving utilization and charge efficiency. NiH2 battery reconditioning effects on life are considered beneficial and may, in fact. extend life based on NiCd experience. In any case, usable capacity data obtained from reconditioning is required for performance evaluation and trend analysis. Characterization and life tests have provided the historical data base used to

The Orbital Design of Alpha Centauri Exoplanet Satellite (ACESat)

NASA Technical Reports Server (NTRS)

Weston, Sasha; Belikov, Rus; Bendek, Eduardo

2015-01-01

Exoplanet candidates discovered by Kepler are too distant for biomarkers to be detected with foreseeable technology. Alpha Centauri has high separation from other stars and is of close proximity to Earth, which makes the binary star system 'low hanging fruit' for scientists. Alpha Centauri Exoplanet Satellite (ACESat) is a mission proposed to Small Explorer Program (SMEX) that will use a coronagraph to search for an orbiting planet around one of the stars of Alpha Centauri. The trajectory design for this mission is presented here where three different trajectories are considered: Low Earth Orbit (LEO), Geosynchronous Orbit (GEO) and a Heliocentric Orbit. Uninterrupted stare time to Alpha Centauri is desirable for meeting science requirements, or an orbit that provides 90% stare time to the science target. The instrument thermal stability also has stringent requirements for proper function, influencing trajectory design.

Lunar Prospector Orbit Determination Uncertainties Using the High Resolution Lunar Gravity Models

NASA Technical Reports Server (NTRS)

Carranza, Eric; Konopliv, Alex; Ryne, Mark

1999-01-01

The Lunar Prospector (LP) mission began on January 6, 1998, when the LP spacecraft was launched from Cape Canaveral, Florida. The objectives of the mission were to determine whether water ice exists at the lunar poles, generate a global compositional map of the lunar surface, detect lunar outgassing, and improve knowledge of the lunar magnetic and gravity fields. Orbit determination of LP performed at the Jet Propulsion Laboratory (JPL) is conducted as part of the principal science investigation of the lunar gravity field. This paper will describe the JPL effort in support of the LP Gravity Investigation. This support includes high precision orbit determination, gravity model validation, and data editing. A description of the mission and its trajectory will be provided first, followed by a discussion of the orbit determination estimation procedure and models. Accuracies will be examined in terms of orbit-to-orbit solution differences, as a function of oblateness model truncation, and inclination in the plane-of-sky. Long term predictions for several gravity fields will be compared to the reconstructed orbits to demonstrate the accuracy of the orbit determination and oblateness fields developed by the Principal Gravity Investigator.

Improved solution accuracy for Landsat-4 (TDRSS-user) orbit determination

NASA Technical Reports Server (NTRS)

Oza, D. H.; Niklewski, D. J.; Doll, C. E.; Mistretta, G. D.; Hart, R. C.

1994-01-01

This paper presents the results of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft, Landsat-4, obtained using a Prototype Filter Smoother (PFS), with the accuracy of an established batch-least-squares system, the Goddard Trajectory Determination System (GTDS). The results of Landsat-4 orbit determination will provide useful experience for the Earth Observing System (EOS) series of satellites. The Landsat-4 ephemerides were estimated for the January 17-23, 1991, timeframe, during which intensive TDRSS tracking data for Landsat-4 were available. Independent assessments were made of the consistencies (overlap comparisons for the batch case and convariances for the sequential case) of solutions produced by the batch and sequential methods. The filtered and smoothed PFS orbit solutions were compared with the definitive GTDS orbit solutions for Landsat-4; the solution differences were generally less than 15 meters.

L-Mapping Solar Energetic Particles from LEO to High Altitudes at High Latitudes

NASA Astrophysics Data System (ADS)

Young, S. L.; Wilson, G.

2017-12-01

The current solar energetic particle (SEP) hazard specification is focused on geosynchronous orbit with some capability at LEO, but there is no specification for the large region between these orbital regimes. The L-mapping technique, which attempts to fill this capability gap, assumes that there is a simple relationship between magnetic L-shells and SEP penetration boundaries that can be exploited. A previous study compared POES observations that had been mapped to the Van Allen Probes with local observations. It found that more than 90% of the mapped and local fluxes were within a factor of four of each other; this is thought to be sucient for operational purposes. One concern with the previous study was the limited number of SEP events that have occurred during the Van Allen Probes mission. The current study examines the L-mapping method's accuracy at higher latitudes. Observations from a satellite that was launched into a HEO orbit with a 63° inclination before the peak of solar cycle 24 are compared to L-mapped POES observations. The larger number of events provides better statistics and the 63° orbit inclination allows us to examine the difference between mapping from POES to the magnetic equator, as in the previous study, and mapping from POES to higher latitudes.

Tethered body problems and relative motion orbit determination

NASA Technical Reports Server (NTRS)

Eades, J. B., Jr.; Wolf, H.

1972-01-01

Selected problems dealing with orbiting tethered body systems have been studied. In addition, a relative motion orbit determination program was developed. Results from these tasks are described and discussed. The expected tethered body motions were examined, analytically, to ascertain what influence would be played by the physical parameters of the tether, the gravity gradient and orbit eccentricity. After separating the motion modes these influences were determined; and, subsequently, the effects of oscillations and/or rotations, on tether force, were described. A study was undertaken, by examining tether motions, to see what type of control actions would be needed to accurately place a mass particle at a prescribed position relative to a main vehicle. Other applications for tethers were studied. Principally these were concerned with the producing of low-level gee forces by means of stabilized tether configurations; and, the initiation of free transfer trajectories from tether supported vehicle relative positions.

Application of Semi-analytical Satellite Theory orbit propagator to orbit determination for space object catalog maintenance

NASA Astrophysics Data System (ADS)

Setty, Srinivas J.; Cefola, Paul J.; Montenbruck, Oliver; Fiedler, Hauke

2016-05-01

Catalog maintenance for Space Situational Awareness (SSA) demands an accurate and computationally lean orbit propagation and orbit determination technique to cope with the ever increasing number of observed space objects. As an alternative to established numerical and analytical methods, we investigate the accuracy and computational load of the Draper Semi-analytical Satellite Theory (DSST). The standalone version of the DSST was enhanced with additional perturbation models to improve its recovery of short periodic motion. The accuracy of DSST is, for the first time, compared to a numerical propagator with fidelity force models for a comprehensive grid of low, medium, and high altitude orbits with varying eccentricity and different inclinations. Furthermore, the run-time of both propagators is compared as a function of propagation arc, output step size and gravity field order to assess its performance for a full range of relevant use cases. For use in orbit determination, a robust performance of DSST is demonstrated even in the case of sparse observations, which is most sensitive to mismodeled short periodic perturbations. Overall, DSST is shown to exhibit adequate accuracy at favorable computational speed for the full set of orbits that need to be considered in space surveillance. Along with the inherent benefits of a semi-analytical orbit representation, DSST provides an attractive alternative to the more common numerical orbit propagation techniques.

Accuracy assessment of BDS precision orbit determination and the influence analysis of site distribution

NASA Astrophysics Data System (ADS)

Chen, Ming; Guo, Jiming; Li, Zhicai; Zhang, Peng; Wu, Junli; Song, Weiwei

2017-04-01

BDS precision orbit determination is a key content of the BDS application, but the inadequate ground stations and the poor distribution of the network are the main reasons for the low accuracy of BDS precise orbit determination. In this paper, the BDS precise orbit determination results are obtained by using the IGS MGEX stations and the Chinese national reference stations，the accuracy of orbit determination of GEO, IGSO and MEO is 10.3cm, 2.8cm and 3.2cm, and the radial accuracy is 1.6cm,1.9cm and 1.5cm.The influence of ground reference stations distribution on BDS precise orbit determination is studied. The results show that the Chinese national reference stations contribute significantly to the BDS orbit determination, the overlap precision of GEO/IGSO/MEO satellites were improved by 15.5%, 57.5% and 5.3% respectively after adding the Chinese stations.Finally, the results of ODOP(orbit distribution of precision) and SLR are verified. Key words: BDS precise orbit determination; accuracy assessment；Chinese national reference stations；reference stations distribution；orbit distribution of precision

Low-Cost Propellant Launch to LEO from a Tethered Balloon - 'Propulsion Depots' Not 'Propellant Depots'

NASA Technical Reports Server (NTRS)

Wilcox, Brian H.; Schneider, Evan G.; Vaughan, David A.; Hall, Jeffrey L.; Yu, Chi Yau

2011-01-01

As we have previously reported, it may be possible to launch payloads into low-Earth orbit (LEO) at a per-kilogram cost that is one to two orders of magnitude lower than current launch systems, using only a relatively small capital investment (comparable to a single large present-day launch). An attractive payload would be large quantities of high-performance chemical rocket propellant (e.g. Liquid Oxygen/Liquid Hydrogen (LO2/LH2)) that would greatly facilitate, if not enable, extensive exploration of the moon, Mars, and beyond.

Characterization of Orbital Debris Via Hyper-Velocity Ground-Based Tests

NASA Technical Reports Server (NTRS)

Cowardin, Heather

2015-01-01

To replicate a hyper-velocity fragmentation event using modern-day spacecraft materials and construction techniques to better improve the existing DoD and NASA breakup models. DebriSat is intended to be representative of modern LEO satellites.Major design decisions were reviewed and approved by Aerospace subject matter experts from different disciplines. DebriSat includes 7 major subsystems. Attitude determination and control system (ADCS), command and data handling (C&DH), electrical power system (EPS), payload, propulsion, telemetry tracking and command (TT&C), and thermal management. To reduce cost, most components are emulated based on existing design of flight hardware and fabricated with the same materials. A key laboratory-based test, Satellite Orbital debris Characterization Impact Test (SOCIT), supporting the development of the DoD and NASA satellite breakup models was conducted at AEDC in 1992 .Breakup models based on SOCIT have supported many applications and matched on-orbit events reasonably well over the years.

Recent Progress on the Second Generation CMORPH: LEO-IR Based Precipitation Estimates and Cloud Motion Vector

NASA Astrophysics Data System (ADS)

Xie, Pingping; Joyce, Robert; Wu, Shaorong

2015-04-01

As reported at the EGU General Assembly of 2014, a prototype system was developed for the second generation CMORPH to produce global analyses of 30-min precipitation on a 0.05olat/lon grid over the entire globe from pole to pole through integration of information from satellite observations as well as numerical model simulations. The second generation CMORPH is built upon the Kalman Filter based CMORPH algorithm of Joyce and Xie (2011). Inputs to the system include rainfall and snowfall rate retrievals from passive microwave (PMW) measurements aboard all available low earth orbit (LEO) satellites, precipitation estimates derived from infrared (IR) observations of geostationary (GEO) as well as LEO platforms, and precipitation simulations from numerical global models. Key to the success of the 2nd generation CMORPH, among a couple of other elements, are the development of a LEO-IR based precipitation estimation to fill in the polar gaps and objectively analyzed cloud motion vectors to capture the cloud movements of various spatial scales over the entire globe. In this presentation, we report our recent work on the refinement for these two important algorithm components. The prototype algorithm for the LEO IR precipitation estimation is refined to achieve improved quantitative accuracy and consistency with PMW retrievals. AVHRR IR TBB data from all LEO satellites are first remapped to a 0.05olat/lon grid over the entire globe and in a 30-min interval. Temporally and spatially co-located data pairs of the LEO TBB and inter-calibrated combined satellite PMW retrievals (MWCOMB) are then collected to construct tables. Precipitation at a grid box is derived from the TBB through matching the PDF tables for the TBB and the MWCOMB. This procedure is implemented for different season, latitude band and underlying surface types to account for the variations in the cloud - precipitation relationship. At the meantime, a sub-system is developed to construct analyzed fields of

Orbit Determination and Navigation of the Solar Terrestrial Relations Observatory (STEREO)

NASA Technical Reports Server (NTRS)

Mesarch, Michael A.; Robertson, Mika; Ottenstein, Neil; Nicholson, Ann; Nicholson, Mark; Ward, Douglas T.; Cosgrove, Jennifer; German, Darla; Hendry, Stephen; Shaw, James

2007-01-01

This paper provides an overview of the required upgrades necessary for navigation of NASA's twin heliocentric science missions, Solar TErestrial RElations Observatory (STEREO) Ahead and Behind. The orbit determination of the STEREO spacecraft was provided by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of the mission operations activities performed by the Johns Hopkins University Applied Physics Laboratory (APL). The changes to FDF's orbit determination software included modeling upgrades as well as modifications required to process the Deep Space Network X-band tracking data used for STEREO. Orbit results as well as comparisons to independently computed solutions are also included. The successful orbit determination support aided in maneuvering the STEREO spacecraft, launched on October 26, 2006 (00:52 Z), to target the lunar gravity assists required to place the spacecraft into their final heliocentric drift-away orbits where they are providing stereo imaging of the Sun.

The voltage threshold for arcing for solar cells in Leo - Flight and ground test results

NASA Technical Reports Server (NTRS)

Ferguson, Dale C.

1986-01-01

Ground and flight results of solar cell arcing in low earth orbit (LEO) conditions are compared and interpreted. It is shown that an apparent voltage threshold for arcing may be produced by a storage power law dependence of arc rate on voltage, combined with a limited observation time. The change in this apparent threshold with plasma density is a reflection of the density dependence of the arc rate. A nearly linear dependence of arc rate on density is inferred from the data. A real voltage threshold for arcing for 2 by 2 cm solar cells may exist however, independent of plasma density, near -230 V relative to the plasma. Here, arc rates may change by more than an order of magnitude for a change of only 30 V in array potential. For 5.9 by 5.9 solar cells, the voltage dependence of the arc rate is steeper, and the data are insufficient to indicate the existence of an arcing increased by an atomic oxygen plasma, as is found in LEO, and by arcing from the backs of welded-through substrates.

The voltage threshold for arcing for solar cells in LEO: Flight and ground test results

NASA Technical Reports Server (NTRS)

Ferguson, D. C.

1986-01-01

Ground and flight results of solar cell arcing in low Earth orbit (LEO) conditions are compared and interpreted. It is shown that an apparent voltage threshold for arcing may be produced by a strong power law dependence of arc rate on voltage, combined with a limited observation time. The change in this apparent threshold with plasma density is a reflection of the density dependence of the arc rate. A nearly linear dependence of arc rate on density is inferred from the data. A real voltage threshold for arcing for 2 by 2 cm solar cells may exist however, independent of plasma density, near -230 V relative to the plasma. Here, arc rates may change by more than an order of magnitude for a change of only 30 V in array potential. For 5.9 by 5.9 solar cells, the voltage dependence of the arc rate is steeper, and the data are insufficient to indicate the existence of an arcing increased by an atomic oxygen plasma, as is found in LEO, and by arcing from the backs of welded-through substrates.

Orbit Determination of the Lunar Reconnaissance Orbiter: Status and Recent Development

NASA Astrophysics Data System (ADS)

Neumann, G. A.; Mazarico, E.; Goossens, S. J.; Nicholas, J. B.; Wagner, R.; Speyerer, E. J.; Smith, D. E.; Zuber, M. T.

2016-12-01

The LRO mission has been operated since June 2009, and the productivity of its seven instruments has led to a wealth of new data and scientific results. The high-resolution data acquired benefit from precise orbit determination (OD), alleviating human intervention in their geolocation and co-registration. The initial position knowledge requirement (50 meters) was met with radio tracking data from the primary NASA White Sands ground station supported by USN, after combination with LOLA altimetric crossovers. LRO-specific gravity field solutions were thus determined and allowed radio-only OD to perform adequately, although secular inclination changes required frequent updates. The high-accuracy gravity fields from GRAIL, with <10 km resolution, further improved the radio-only orbit reconstruction quality. However, it is in part limited by the 0.3-0.5 mm/s measurement noise level in the S-band. One-way tracking through Laser Ranging can supplement the tracking available for OD with 28 Hz ranges with 20 cm single-shot precision, but is available only on the nearside. The LOLA altimetric data afford accurate, independent information about LRO's orbit, with a very different geometry that includes coverage over the lunar farside. With LOLA's highest-quality topographic model of the Moon and the Kaguya Terrain Camera stereo-derived elevation model, and their combination named SLDEM2015, another altimetric measurement is now possible to use in OD. This `direct altimetry' tracking type was developed to calibrate the laser boresight pointing of the IceSAT/GLAS altimeter, as differences in geolocated height of profiles with respect to an ocean surface reference geoid were primarily attributed to pointing errors. We extended this technique to short-scale, high-resolution targets, and can now use the SLDEM2015 topographic model as a basemap to match individual LOLA tracks during OD, adjusting both spacecraft position and pointing to minimize the discrepancies. Comparisons with

Detection and laser ranging of orbital objects using optical methods

NASA Astrophysics Data System (ADS)

Wagner, P.; Hampf, D.; Sproll, F.; Hasenohr, T.; Humbert, L.; Rodmann, J.; Riede, W.

2016-09-01

Laser ranging to satellites (SLR) in earth orbit is an established technology used for geodesy, fundamental science and precise orbit determination. A combined active and passive optical measurement system using a single telescope mount is presented which performs precise ranging measurements of retro reflector equipped objects in low earth orbit (LEO). The German Aerospace Center (DLR) runs an observatory in Stuttgart where a system has been assembled completely from commercial off-the-shelf (COTS) components. The visible light directed to the tracking camera is used to perform angular measurements of objects under investigation. This is done astrometrically by comparing the apparent target position with cataloged star positions. First successful satellite laser ranging was demonstrated recently using an optical fiber directing laser pulses onto the astronomical mount. The transmitter operates at a wavelength of 1064 nm with a repetition rate of 3 kHz and pulse energy of 25 μJ. A motorized tip/tilt mount allows beam steering of the collimated beam with μrad accuracy. The returning photons reflected from the object in space are captured with the tracking telescope. A special low aberration beam splitter unit was designed to separate the infrared from visible light. This allows passive optical closed loop tracking and operation of a single photon detector for time of flight measurements at a single telescope simultaneously. The presented innovative design yields to a compact and cost effective but very precise ranging system which allows orbit determination.

Experimental Tests of UltraFlex Array Designs in Low Earth Orbital and Geosynchronous Charging Environments

NASA Technical Reports Server (NTRS)

Galofaro, Joel T.; Vayner, Boris V.; Hillard, Grover B.

2011-01-01

The present ground based investigations give the first definitive look describing the expected on-orbit charging behavior of Orion UltraFlex array coupons in the Low Earth Orbital and Geosynchronous Environments. Furthermore, it is important to note that the LEO charging environment also applies to the International Space Station as well as to the lunar mission charging environments. The GEO charging environment includes the bounding case for all lunar orbital and lunar surface mission environments. The UltraFlex thin film photovoltaic array technology has been targeted to become the sole power system for life support and on-orbit power for the manned Aires Crew Exploration Vehicle. It is therefore, crucial to gain an understanding of the complex charging behavior to answer some of the basic performance and survivability issues in an attempt to ascertain that a single UltraFlex array design will be able to cope with the projected worst case LEO and GEO charging environments. Testing was limited to four array coupons, two coupons each from two different array manufactures, Emcore and Spectrolab. The layout of each array design is identical and varies only in the actual cell technology used. The individual array cells from each manufacturer have an antireflection layered coating and come in two different varieties either uncoated (only AR coating) or coated with a thin conducting ITO layer. The LEO Plasma tests revealed that all four coupons passed the arc threshold -120 V bias tests. GEO electron gun charging tests revealed that only front side area of ITO coated coupons passed tests. Only the Emcore AR array passed backside Stage 2 GEO Tests.

An intelligent interface for satellite operations: Your Orbit Determination Assistant (YODA)

NASA Technical Reports Server (NTRS)

Schur, Anne

1988-01-01

An intelligent interface is often characterized by the ability to adapt evaluation criteria as the environment and user goals change. Some factors that impact these adaptations are redefinition of task goals and, hence, user requirements; time criticality; and system status. To implement adaptations affected by these factors, a new set of capabilities must be incorporated into the human-computer interface design. These capabilities include: (1) dynamic update and removal of control states based on user inputs, (2) generation and removal of logical dependencies as change occurs, (3) uniform and smooth interfacing to numerous processes, databases, and expert systems, and (4) unobtrusive on-line assistance to users of concepts were applied and incorporated into a human-computer interface using artificial intelligence techniques to create a prototype expert system, Your Orbit Determination Assistant (YODA). YODA is a smart interface that supports, in real teime, orbit analysts who must determine the location of a satellite during the station acquisition phase of a mission. Also described is the integration of four knowledge sources required to support the orbit determination assistant: orbital mechanics, spacecraft specifications, characteristics of the mission support software, and orbit analyst experience. This initial effort is continuing with expansion of YODA's capabilities, including evaluation of results of the orbit determination task.

Implementation of a low-cost, commercial orbit determination system

NASA Technical Reports Server (NTRS)

Corrigan, Jim

1994-01-01

This paper describes the implementation and potential applications of a workstation-based orbit determination system developed by Storm Integration, Inc. called the Precision Orbit Determination System (PODS). PODS is offered as a layered product to the commercially-available Satellite Tool Kit (STK) produced by Analytical Graphics, Inc. PODS also incorporates the Workstation/Precision Orbit Determination (WS/POD) product offered by Van Martin System, Inc. The STK graphical user interface is used to access and invoke the PODS capabilities and to display the results. WS/POD is used to compute a best-fit solution to user-supplied tracking data. PODS provides the capability to simultaneously estimate the orbits of up to 99 satellites based on a wide variety of observation types including angles, range, range rate, and Global Positioning System (GPS) data. PODS can also estimate ground facility locations, Earth geopotential model coefficients, solar pressure and atmospheric drag parameters, and observation data biases. All determined data is automatically incorporated into the STK data base, which allows storage, manipulation and export of the data to other applications. PODS is offered in three levels: Standard, Basic GPS and Extended GPS. Standard allows processing of non-GPS observation types for any number of vehicles and facilities. Basic GPS adds processing of GPS pseudo-ranging data to the Standard capabilities. Extended GPS adds the ability to process GPS carrier phase data.

Orbit determination and prediction of GEO satellite of BeiDou during repositioning maneuver

NASA Astrophysics Data System (ADS)

Cao, Fen; Yang, XuHai; Li, ZhiGang; Sun, BaoQi; Kong, Yao; Chen, Liang; Feng, Chugang

2014-11-01

In order to establish a continuous GEO satellite orbit during repositioning maneuvers, a suitable maneuver force model has been established associated with an optimal orbit determination method and strategy. A continuous increasing acceleration is established by constructing a constant force that is equivalent to the pulse force, with the mass of the satellite decreasing throughout maneuver. This acceleration can be added to other accelerations, such as solar radiation, to obtain the continuous acceleration of the satellite. The orbit determination method and strategy are illuminated, with subsequent assessment of the orbit being determined and predicted accordingly. The orbit of the GEO satellite during repositioning maneuver can be determined and predicted by using C-Band pseudo-range observations of the BeiDou GEO satellite with COSPAR ID 2010-001A in 2011 and 2012. The results indicate that observations before maneuver do affect orbit determination and prediction, and should therefore be selected appropriately. A more precise orbit and prediction can be obtained compared to common short arc methods when observations starting 1 day prior the maneuver and 2 h after the maneuver are adopted in POD (Precise Orbit Determination). The achieved URE (User Range Error) under non-consideration of satellite clock errors is better than 2 m within the first 2 h after maneuver, and less than 3 m for further 2 h of orbit prediction.

Filter Strategies for Mars Science Laboratory Orbit Determination

NASA Technical Reports Server (NTRS)

Thompson, Paul F.; Gustafson, Eric D.; Kruizinga, Gerhard L.; Martin-Mur, Tomas J.

2013-01-01

The Mars Science Laboratory (MSL) spacecraft had ambitious navigation delivery and knowledge accuracy requirements for landing inside Gale Crater. Confidence in the orbit determination (OD) solutions was increased by investigating numerous filter strategies for solving the orbit determination problem. We will discuss the strategy for the different types of variations: for example, data types, data weights, solar pressure model covariance, and estimating versus considering model parameters. This process generated a set of plausible OD solutions that were compared to the baseline OD strategy. Even implausible or unrealistic results were helpful in isolating sensitivities in the OD solutions to certain model parameterizations or data types.

Orbit Determination and Navigation of the Solar Terrestrial Relations Observatory (STEREO)

NASA Technical Reports Server (NTRS)

Mesarch, Michael; Robertson, Mika; Ottenstein, Neil; Nicholson, Ann; Nicholson, Mark; Ward, Douglas T.; Cosgrove, Jennifer; German, Darla; Hendry, Stephen; Shaw, James

2007-01-01

This paper provides an overview of the required upgrades necessary for navigation of NASA's twin heliocentric science missions, Solar TErestrial RElations Observatory (STEREO) Ahead and Behind. The orbit determination of the STEREO spacecraft was provided by the NASA Goddard Space Flight Center's (GSFC) Flight Dynamics Facility (FDF) in support of the mission operations activities performed by the Johns Hopkins University Applied Physics Laboratory (APL). The changes to FDF s orbit determination software included modeling upgrades as well as modifications required to process the Deep Space Network X-band tracking data used for STEREO. Orbit results as well as comparisons to independently computed solutions are also included. The successful orbit determination support aided in maneuvering the STEREO spacecraft, launched on October 26, 2006 (00:52 Z), to target the lunar gravity assists required to place the spacecraft into their final heliocentric drift-away orbits where they are providing stereo imaging of the Sun.

Achieving and Validating the 1-centimeter Orbit: JASON-1 Precision Orbit Determination Using GPS, SLR, DORIS and Altimeter data

NASA Technical Reports Server (NTRS)

Luthcke, Scott B.; Zelensky, Nikita P.; Rowlands, David D.; Lemoine, Frank G.; Williams, Teresa A.

2003-01-01

Jason-1, launched on December 7, 2001, is continuing the time series of centimeter level ocean topography observations as the follow-on to the highly successful TOPEX/POSEIDON (T/P) radar altimeter satellite. The precision orbit determination (POD) is a critical component to meeting the ocean topography goals of the mission. Jason-1 is no exception and has set a 1 cm radial orbit accuracy goal, which represents a factor of two improvement over what is currently being achieved for T/P. The challenge to precision orbit determination (POD) is both achieving the 1 cm radial orbit accuracy and evaluating and validating the performance of the 1 cm orbit. Fortunately, Jason-1 POD can rely on four independent tracking data types including near continuous tracking data from the dual frequency codeless BlackJack GPS receiver. In addition, to the enhanced GPS receiver, Jason-1 carries significantly improved SLR and DORIS tracking systems along with the altimeter itself. We demonstrate the 1 cm radial orbit accuracy goal has been achieved using GPS data alone in a reduced dynamic solution. It is also shown that adding SLR data to the GPS-based solutions improves the orbits even further. In order to assess the performance of these orbits it is necessary to process all of the available tracking data (GPS, SLR, DORIS and altimeter crossover differences) as either dependent or independent of the orbit solutions. It was also necessary to compute orbit solutions using various combinations of the four available tracking data in order to independently assess the orbit performance. Towards this end, we have greatly improved orbits determined solely from SLR+DORIS data by applying the reduced dynamic solution strategy. In addition, we have computed reduced dynamic orbits based on SLR, DORIS and crossover data that are a significant improvement over the SLR and DORIS based dynamic solutions. These solutions provide the best performing orbits for independent validation of the GPS

A Low Earth Orbit satellite marine communication system demonstration

NASA Technical Reports Server (NTRS)

Elms, T. Keith; Butt, Kenneth A.; Asmus, Ken W.

1995-01-01

An application of Low Earth Orbit (LEO) satellite communications technology was investigated during a joint Canadian/American scientific expedition to the north pole in the summer of 1994. The Canadian ice breaker involved, was equipped with a store-and-forward LEO satellite terminal which was linked to a ground station in St. John's, Newfoundland, via the near-polar-orbiting satellite, HealthSat-l. The objective was to evaluate the performance of such a system while providing an alternate means of communications in the far north. The system performed well, given its inherent limitations. All 151 attempts to send data files to the ship were successful. Only two (2) of the 35 attempts to send files from the ship were unsuccessful. The files ranged in size from 0.1 to 60 Kbytes. In the high arctic, above 80 deg north, this system often provided the only practical means of data communications. This experiment demonstrated the potential of such a system for not-real-time communications with remote and/or mobile stations, and highlighted the many issues involved. This paper describes the project objectives, system configuration and experimental procedure used, related technical issues, trial results, future work, and conclusions.

An advanced analysis method of initial orbit determination with too short arc data

NASA Astrophysics Data System (ADS)

Li, Binzhe; Fang, Li

2018-02-01

This paper studies the initial orbit determination (IOD) based on space-based angle measurement. Commonly, these space-based observations have short durations. As a result, classical initial orbit determination algorithms give poor results, such as Laplace methods and Gauss methods. In this paper, an advanced analysis method of initial orbit determination is developed for space-based observations. The admissible region and triangulation are introduced in the method. Genetic algorithm is also used for adding some constraints of parameters. Simulation results show that the algorithm can successfully complete the initial orbit determination.

Comparison of ERBS orbit determination accuracy using batch least-squares and sequential methods

NASA Technical Reports Server (NTRS)

Oza, D. H.; Jones, T. L.; Fabien, S. M.; Mistretta, G. D.; Hart, R. C.; Doll, C. E.

1991-01-01

The Flight Dynamics Div. (FDD) at NASA-Goddard commissioned a study to develop the Real Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination of spacecraft on a DOS based personal computer (PC). An overview is presented of RTOD/E capabilities and the results are presented of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOS/E on a PC with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. RTOD/E was used to perform sequential orbit determination for the Earth Radiation Budget Satellite (ERBS), and the Goddard Trajectory Determination System (GTDS) was used to perform the batch least squares orbit determination. The estimated ERBS ephemerides were obtained for the Aug. 16 to 22, 1989, timeframe, during which intensive TDRSS tracking data for ERBS were available. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for ERBS; the solution differences were less than 40 meters after the filter had reached steady state.

Comparison of ERBS orbit determination accuracy using batch least-squares and sequential methods

NASA Astrophysics Data System (ADS)

Oza, D. H.; Jones, T. L.; Fabien, S. M.; Mistretta, G. D.; Hart, R. C.; Doll, C. E.

1991-10-01

The Flight Dynamics Div. (FDD) at NASA-Goddard commissioned a study to develop the Real Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination of spacecraft on a DOS based personal computer (PC). An overview is presented of RTOD/E capabilities and the results are presented of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOS/E on a PC with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. RTOD/E was used to perform sequential orbit determination for the Earth Radiation Budget Satellite (ERBS), and the Goddard Trajectory Determination System (GTDS) was used to perform the batch least squares orbit determination. The estimated ERBS ephemerides were obtained for the Aug. 16 to 22, 1989, timeframe, during which intensive TDRSS tracking data for ERBS were available. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for ERBS; the solution differences were less than 40 meters after the filter had reached steady state.

Nuclear reactor power for an electrically powered orbital transfer vehicle

NASA Technical Reports Server (NTRS)

Jaffe, L.; Beatty, R.; Bhandari, P.; Chow, E.; Deininger, W.; Ewell, R.; Fujita, T.; Grossman, M.; Kia, T.; Nesmith, B.

1987-01-01

To help determine the systems requirements for a 300-kWe space nuclear reactor power system, a mission and spacecraft have been examined which utilize electric propulsion and this nuclear reactor power for multiple transfers of cargo between low earth orbit (LEO) and geosynchronous earth orbit (GEO). A propulsion system employing ion thrusters and xenon propellant was selected. Propellant and thrusters are replaced after each sortie to GEO. The mass of the Orbital Transfer Vehicle (OTV), empty and dry, is 11,000 kg; nominal propellant load is 5000 kg. The OTV operates between a circular orbit at 925 km altitude, 28.5 deg inclination, and GEO. Cargo is brought to the OTV by Shuttle and an Orbital Maneuvering Vehicle (OMV); the OTV then takes it to GEO. The OTV can also bring cargo back from GEO, for transfer by OMV to the Shuttle. OTV propellant is resupplied and the ion thrusters are replaced by the OMV before each trip to GEO. At the end of mission life, the OTV's electric propulsion is used to place it in a heliocentric orbit so that the reactor will not return to earth. The nominal cargo capability to GEO is 6000 kg with a transit time of 120 days; 1350 kg can be transferred in 90 days, and 14,300 kg in 240 days. These capabilities can be considerably increased by using separate Shuttle launches to bring up propellant and cargo, or by changing to mercury propellant.

A Numerical Approach to Estimate the Ballistic Coefficient of Space Debris from TLE Orbital Data

NASA Technical Reports Server (NTRS)

Narkeliunas, Jonas

2016-01-01

Low Earth Orbit (LEO) is full of space debris, which consist of spent rocket stages, old satellites and fragments from explosions and collisions. As of 2009, more than 21,000 orbital debris larger than 10 cm are known to exist], and while it is hard to track anything smaller than that, the estimated population of particles between 1 and 10 cm in diameter is approximately 500,000, whereas small as 1 cm exceeds 100 million. These objects orbit Earth with huge kinetic energies speeds usually exceed 7 kms. The shape of their orbit varies from almost circular to highly elliptical and covers all LEO, a region in space between 160 and 2,000 km above sea level. Unfortunately, LEO is also the place where most of our active satellites are situated, as well as, International Space Station (ISS) and Hubble Space Telescope, whose orbits are around 400 and 550 km above sea level, respectively.This poses a real threat as debris can collide with satellites and deal substantial damage or even destroy them.Collisions between two or more debris create clouds of smaller debris, which are harder to track and increase overall object density and collision probability. At some point, the debris density couldthen reach a critical value, which would start a chain reaction and the number of space debris would grow exponentially. This phenomenon was first described by Kessler in 1978 and he concluded that it would lead to creation of debris belt, which would vastly complicate satellite operations in LEO. The debris density is already relatively high, as seen from several necessary debris avoidance maneuvers done by Shuttle, before it was discontinued, and ISS. But not all satellites have a propulsion system to avoid collision, hence different methods need to be applied. One of the proposed collision avoidance concepts is called LightForce and it suggests using photon pressure to induce small orbital corrections to deflect debris from colliding. This method is very efficient as seen from

A demonstration of high precision GPS orbit determination for geodetic applications

NASA Technical Reports Server (NTRS)

Lichten, S. M.; Border, J. S.

1987-01-01

High precision orbit determination of Global Positioning System (GPS) satellites is a key requirement for GPS-based precise geodetic measurements and precise low-earth orbiter tracking, currently under study at JPL. Different strategies for orbit determination have been explored at JPL with data from a 1985 GPS field experiment. The most successful strategy uses multi-day arcs for orbit determination and includes fine tuning of spacecraft solar pressure coefficients and station zenith tropospheric delays using the GPS data. Average rms orbit repeatability values for 5 of the GPS satellites are 1.0, 1.2, and 1.7 m in altitude, cross-track, and down-track componenets when two independent 5-day fits are compared. Orbit predictions up to 24 hours outside the multi-day arcs agree within 4 m of independent solutions obtained with well tracked satellites in the prediction interval. Baseline repeatability improves with multi-day as compared to single-day arc orbit solutions. When tropospheric delay fluctuations are modeled with process noise, significant additional improvement in baseline repeatability is achieved. For a 246-km baseline, with 6-day arc solutions for GPS orbits, baseline repeatability is 2 parts in 100 million (0.4-0.6 cm) for east, north, and length components and 8 parts in 100 million for the vertical component. For 1314 and 1509 km baselines with the same orbits, baseline repeatability is 2 parts in 100 million for the north components (2-3 cm) and 4 parts in 100 million or better for east, length, and vertical components.

Orbit determination and orbit control for the Earth Observing System (EOS) AM spacecraft

NASA Technical Reports Server (NTRS)

Herberg, Joseph R.; Folta, David C.

1993-01-01

Future NASA Earth Observing System (EOS) Spacecraft will make measurements of the earth's clouds, oceans, atmosphere, land and radiation balance. These EOS Spacecraft will be part of the NASA Mission to Planet Earth. This paper specifically addresses the EOS AM Spacecraft, referred to as 'AM' because it has a sun-synchronous orbit with a 10:30 AM descending node. This paper describes the EOS AM Spacecraft mission orbit requirements, orbit determination, orbit control, and navigation system impact on earth based pointing. The EOS AM Spacecraft will be the first spacecraft to use the TDRSS Onboard Navigation System (TONS) as the primary means of navigation. TONS flight software will process one-way forward Doppler measurements taken during scheduled TDRSS contacts. An extended Kalman filter will estimate spacecraft position, velocity, drag coefficient correction, and ultrastable master oscillator frequency bias and drift. The TONS baseline algorithms, software, and hardware implementation are described in this paper. TONS integration into the EOS AM Spacecraft Guidance, Navigation, and Control (GN&C) System; TONS assisted onboard time maintenance; and the TONS Ground Support System (TGSS) are also addressed.

Autonomous Path Planning for On-Orbit Servicing Vehicles

NASA Astrophysics Data System (ADS)

McInnes, C. R.

On-orbit servicing has long been considered as a means of reducing mission costs. While automated on-orbit servicing of satellites in LEO and GEO has yet to be realised, the International Space Station (ISS) will require servicing in a number of forms for re-supply, external visual inspection and maintenance. This paper will discuss a unified approach to path planning for such servicing vehicles using artificial potential field methods. In particular, path constrained rendezvous and docking of the ESA Automated Transfer Vehicle (ATV) at the ISS will be investigated as will mission and path planning tools for the Daimler-Chrysler Aerospace ISS Inspector free-flying camera. Future applications for free-flying microcameras and co-operative control between multiple free-flyers for on-orbit assembly will also be considered.

Low-Cost Propellant Launch to Earth Orbit from a Tethered Balloon

NASA Technical Reports Server (NTRS)

Wilcox, Brian H.

2006-01-01

Propellant will be more than 85% of the mass that needs to be lofted into Low Earth Orbit (LEO) in the planned program of Exploration of the Moon, Mars, and beyond. This paper describes a possible means for launching thousands of tons of propellant per year into LEO at a cost 15 to 30 times less than the current launch cost per kilogram. The basic idea is to mass-produce very simple, small and relatively low-performance rockets at a cost per kilogram comparable to automobiles, instead of the 25X greater cost that is customary for current launch vehicles that are produced in small quantities and which are manufactured with performance near the limits of what is possible. These small, simple rockets can reach orbit because they are launched above 95% of the atmosphere, where the drag losses even on a small rocket are acceptable, and because they can be launched nearly horizontally with very simple guidance based primarily on spin-stabilization. Launching above most of the atmosphere is accomplished by winching the rocket up a tether to a balloon. A fuel depot in equatorial orbit passes over the launch site on every orbit (approximately every 90 minutes). One or more rockets can be launched each time the fuel depot passes overhead, so the launch rate can be any multiple of 6000 small rockets per year, a number that is sufficient to reap the benefits of mass production.

Benefits of high aerodynamic efficiency to orbital transfer vehicles

NASA Technical Reports Server (NTRS)

Andrews, D. G.; Norris, R. B.; Paris, S. W.

1984-01-01

The benefits and costs of high aerodynamic efficiency on aeroassisted orbital transfer vehicles (AOTV) are analyzed. Results show that a high lift to drag (L/D) AOTV can achieve significant velocity savings relative to low L/D aerobraked OTV's when traveling round trip between low Earth orbits (LEO) and alternate orbits as high as geosynchronous Earth orbit (GEO). Trajectory analysis is used to show the impact of thermal protection system technology and the importance of lift loading coefficient on vehicle performance. The possible improvements in AOTV subsystem technologies are assessed and their impact on vehicle inert weight and performance noted. Finally, the performance of high L/D AOTV concepts is compared with the performances of low L/D aeroassisted and all propulsive OTV concepts to assess the benefits of aerodynamic efficiency on this class of vehicle.

Autonomous space processor for orbital debris

NASA Technical Reports Server (NTRS)

Ramohalli, Kumar; Marine, Micky; Colvin, James; Crockett, Richard; Sword, Lee; Putz, Jennifer; Woelfle, Sheri

1991-01-01

The development of an Autonomous Space Processor for Orbital Debris (ASPOD) was the goal. The nature of this craft, which will process, in situ, orbital debris using resources available in low Earth orbit (LEO) is explained. The serious problem of orbital debris is briefly described and the nature of the large debris population is outlined. The focus was on the development of a versatile robotic manipulator to augment an existing robotic arm, the incorporation of remote operation of the robotic arms, and the formulation of optimal (time and energy) trajectory planning algorithms for coordinated robotic arms. The mechanical design of the new arm is described in detail. The work envelope is explained showing the flexibility of the new design. Several telemetry communication systems are described which will enable the remote operation of the robotic arms. The trajectory planning algorithms are fully developed for both the time optimal and energy optimal problems. The time optimal problem is solved using phase plane techniques while the energy optimal problem is solved using dynamic programming.

Operational Challenges In TDRS Post-Maneuver Orbit Determination

NASA Technical Reports Server (NTRS)

Laing, Jason; Myers, Jessica; Ward, Douglas; Lamb, Rivers

2015-01-01

The GSFC Flight Dynamics Facility (FDF) is responsible for daily and post maneuver orbit determination for the Tracking and Data Relay Satellite System (TDRSS). The most stringent requirement for this orbit determination is 75 meters total position accuracy (3-sigma) predicted over one day for Terra's onboard navigation system. To maintain an accurate solution onboard Terra, a solution is generated and provided by the FDF Four hours after a TDRS maneuver. A number of factors present challenges to this support, such as maneuver prediction uncertainty and potentially unreliable tracking from User satellities. Reliable support is provided by comparing an extended Kalman Filter (estimated using ODTK) against a Batch Least Squares system (estimated using GTDS).

An Independent Orbit Determination Simulation for the OSIRIS-REx Asteroid Sample Return Mission

NASA Technical Reports Server (NTRS)

Getzandanner, Kenneth; Rowlands, David; Mazarico, Erwan; Antreasian, Peter; Jackman, Coralie; Moreau, Michael

2016-01-01

After arriving at the near-Earth asteroid (101955) Bennu in late 2018, the OSIRIS-REx spacecraft will execute a series of observation campaigns and orbit phases to accurately characterize Bennu and ultimately collect a sample of pristine regolith from its surface. While in the vicinity of Bennu, the OSIRIS-REx navigation team will rely on a combination of ground-based radiometric tracking data and optical navigation (OpNav) images to generate and deliver precision orbit determination products. Long before arrival at Bennu, the navigation team is performing multiple orbit determination simulations and thread tests to verify navigation performance and ensure interfaces between multiple software suites function properly. In this paper, we will summarize the results of an independent orbit determination simulation of the Orbit B phase of the mission performed to test the interface between the OpNav image processing and orbit determination software packages.

GPS-Based Reduced Dynamic Orbit Determination Using Accelerometer Data

NASA Technical Reports Server (NTRS)

VanHelleputte, Tom; Visser, Pieter

2007-01-01

Currently two gravity field satellite missions, CHAMP and GRACE, are equipped with high sensitivity electrostatic accelerometers, measuring the non-conservative forces acting on the spacecraft in three orthogonal directions. During the gravity field recovery these measurements help to separate gravitational and non-gravitational contributions in the observed orbit perturbations. For precise orbit determination purposes all these missions have a dual-frequency GPS receiver on board. The reduced dynamic technique combines the dense and accurate GPS observations with physical models of the forces acting on the spacecraft, complemented by empirical accelerations, which are stochastic parameters adjusted in the orbit determination process. When the spacecraft carries an accelerometer, these measured accelerations can be used to replace the models of the non-conservative forces, such as air drag and solar radiation pressure. This approach is implemented in a batch least-squares estimator of the GPS High Precision Orbit Determination Software Tools (GHOST), developed at DLR/GSOC and DEOS. It is extensively tested with data of the CHAMP and GRACE satellites. As accelerometer observations typically can be affected by an unknown scale factor and bias in each measurement direction, they require calibration during processing. Therefore the estimated state vector is augmented with six parameters: a scale and bias factor for the three axes. In order to converge efficiently to a good solution, reasonable a priori values for the bias factor are necessary. These are calculated by combining the mean value of the accelerometer observations with the mean value of the non-conservative force models and empirical accelerations, estimated when using these models. When replacing the non-conservative force models with accelerometer observations and still estimating empirical accelerations, a good orbit precision is achieved. 100 days of GRACE B data processing results in a mean orbit fit of

Accurate orbit determination strategies for the tracking and data relay satellites

NASA Technical Reports Server (NTRS)

Oza, D. H.; Bolvin, D. T.; Lorah, J. M.; Lee, T.; Doll, C. E.

1995-01-01

The National Aeronautics and Space Administration (NASA) has developed the Tracking and Data Relay Satellite (TDRS) System (TDRSS) for tracking and communications support of low Earth-orbiting satellites. TDRSS has the operational capability of providing 85% coverage for TDRSS-user spacecraft. TDRSS currently consists of five geosynchronous spacecraft and the White Sands Complex (WSC) at White Sands, New Mexico. The Bilateration Ranging Transponder System (BRTS) provides range and Doppler measurements for each TDRS. The ground-based BRTS transponders are tracked as if they were TDRSS-user spacecraft. Since the positions of the BRTS transponders are known, their radiometric tracking measurements can be used to provide a well-determined ephemeris for the TDRS spacecraft. For high-accuracy orbit determination of a TDRSS user, such as the Ocean Topography Experiment (TOPEX)/Poseidon spacecraft, high-accuracy TDRS orbits are required. This paper reports on successive refinements in improved techniques and procedures leading to more accurate TDRS orbit determination strategies using the Goddard Trajectory Determination System (GTDS). These strategies range from the standard operational solution using only the BRTS tracking measurements to a sophisticated iterative process involving several successive simultaneous solutions for multiple TDRSs and a TDRSS-user spacecraft. Results are presented for GTDS-generated TDRS ephemerides produced in simultaneous solutions with the TOPEX/Poseidon spacecraft. Strategies with different user spacecraft, as well as schemes for recovering accurate TDRS orbits following a TDRS maneuver, are also presented. In addition, a comprehensive assessment and evaluation of alternative strategies for TDRS orbit determination, excluding BRTS tracking measurements, are presented.

Orbit determination and gravity field recovery from Doppler tracking data to the Lunar Reconnaissance Orbiter

NASA Astrophysics Data System (ADS)

Maier, Andrea; Baur, Oliver

2016-03-01

We present results for Precise Orbit Determination (POD) of the Lunar Reconnaissance Orbiter (LRO) based on two-way Doppler range-rates over a time span of ~13 months (January 3, 2011 to February 9, 2012). Different orbital arc lengths and various sets of empirical parameters were tested to seek optimal parametrization. An overlap analysis covering three months of Doppler data shows that the most precise orbits are obtained using an arc length of 2.5 days and estimating arc-wise constant empirical accelerations in along track direction. The overlap analysis over the entire investigated time span of 13 months indicates an orbital precision of 13.79 m, 14.17 m, and 1.28 m in along track, cross track, and radial direction, respectively, with 21.32 m in total position. We compare our orbits to the official science orbits released by the US National Aeronautics and Space Administration (NASA). The differences amount to 9.50 m, 6.98 m, and 1.50 m in along track, cross track, and radial direction, respectively, as well as 12.71 m in total position. Based on the reconstructed LRO orbits, we estimated lunar gravity field coefficients up to spherical harmonic degree and order 60. The results are compared to gravity field solutions derived from data collected by other lunar missions.

8-12 GHz Radio Observations of Flare Activity On M dwarf CN Leo

NASA Astrophysics Data System (ADS)

Wofford, Alia; Villadsen, Jackie; Quintana, Elisa; Barclay, Thomas; Thackeray, Beverly

2018-01-01

Red dwarfs are cool stars that make up 70% of all stars. Red dwarfs can be utilized to detect potentially habitable planets but they have particularly strong magnetic activity that can be detrimental to orbiting planets’ atmospheres and habitability. A coronal mass ejection (CME) is an eruption of magnetized plasma from the star that is ejected into the interplanetary medium which can erode a planet’s atmosphere daily. Based on the sun CMEs are expected to produce very bright radio bursts along with optical flares. We are using M dwarf CN Leo, a well studied flare star that was in the K2 campaign field in summer 2017, as a template to understand the relationship between radio and optical flares and the space weather conditions impacting M dwarf planets. Using radio frequencies ranging from 0.22 GHz-12 GHz we search for simultaneous radio bursts and optical flares to infer if CMEs, flares or aurorae are occurring on the star. I will present the 8-12 GHz radio data from eight 1.5-hour observations with simultaneous optical data. CN Leo produced a bright non-thermal radio flare that lasted approximately for a day during two consecutive observations, with a gyrosynchrotron emission mechanism.

Application of GPS tracking techniques to orbit determination for TDRS

NASA Technical Reports Server (NTRS)

Haines, B. J.; Lichten, S. M.; Malla, R. P.; Wu, S. C.

1993-01-01

In this paper, we evaluate two fundamentally different approaches to TDRS orbit determination utilizing Global Positioning System (GPS) technology and GPS-related techniques. In the first, a GPS flight receiver is deployed on the TDRSS spacecraft. The TDRS ephemerides are determined using direct ranging to the GPS spacecraft, and no ground network is required. In the second approach, the TDRSS spacecraft broadcast a suitable beacon signal, permitting the simultaneous tracking of GPS and TDRSS satellites from a small ground network. Both strategies can be designed to meet future operational requirements for TDRS-2 orbit determination.

Enhanced orbit determination filter sensitivity analysis: Error budget development

NASA Technical Reports Server (NTRS)

Estefan, J. A.; Burkhart, P. D.

1994-01-01

An error budget analysis is presented which quantifies the effects of different error sources in the orbit determination process when the enhanced orbit determination filter, recently developed, is used to reduce radio metric data. The enhanced filter strategy differs from more traditional filtering methods in that nearly all of the principal ground system calibration errors affecting the data are represented as filter parameters. Error budget computations were performed for a Mars Observer interplanetary cruise scenario for cases in which only X-band (8.4-GHz) Doppler data were used to determine the spacecraft's orbit, X-band ranging data were used exclusively, and a combined set in which the ranging data were used in addition to the Doppler data. In all three cases, the filter model was assumed to be a correct representation of the physical world. Random nongravitational accelerations were found to be the largest source of error contributing to the individual error budgets. Other significant contributors, depending on the data strategy used, were solar-radiation pressure coefficient uncertainty, random earth-orientation calibration errors, and Deep Space Network (DSN) station location uncertainty.

Space Launch System Mission Flexibility Assessment

NASA Technical Reports Server (NTRS)

Monk, Timothy; Holladay, Jon; Sanders, Terry; Hampton, Bryan

2012-01-01

The Space Launch System (SLS) is envisioned as a heavy lift vehicle that will provide the foundation for future beyond low Earth orbit (LEO) missions. While multiple assessments have been performed to determine the optimal configuration for the SLS, this effort was undertaken to evaluate the flexibility of various concepts for the range of missions that may be required of this system. These mission scenarios include single launch crew and/or cargo delivery to LEO, single launch cargo delivery missions to LEO in support of multi-launch mission campaigns, and single launch beyond LEO missions. Specifically, we assessed options for the single launch beyond LEO mission scenario using a variety of in-space stages and vehicle staging criteria. This was performed to determine the most flexible (and perhaps optimal) method of designing this particular type of mission. A specific mission opportunity to the Jovian system was further assessed to determine potential solutions that may meet currently envisioned mission objectives. This application sought to significantly reduce mission cost by allowing for a direct, faster transfer from Earth to Jupiter and to determine the order-of-magnitude mass margin that would be made available from utilization of the SLS. In general, smaller, existing stages provided comparable performance to larger, new stage developments when the mission scenario allowed for optimal LEO dropoff orbits (e.g. highly elliptical staging orbits). Initial results using this method with early SLS configurations and existing Upper Stages showed the potential of capturing Lunar flyby missions as well as providing significant mass delivery to a Jupiter transfer orbit.

Low Earth orbit atomic oxygen simulation for durability evaluation of solar reflector surfaces

NASA Technical Reports Server (NTRS)

Degroh, Kim K.; Banks, Bruce A.

1992-01-01

To evaluate the performance and durability of solar reflector surfaces in the atomic oxygen environment typical of low Earth orbit (LEO), one must expose the reflector surface either directly to LEO or to ground-laboratory atomic oxygen environments. Although actual LEO exposures are most desired, such opportunities are typically scarce, expensive, and of limited duration. As a result, ground-laboratory exposures must be relied upon as the most practical long-term durability evaluation technique. Plasma ashers are widely used as LEO simulation facilities by producing atomic oxygen environments for durability evaluation of potential spacecraft materials. Atomic oxygen arrival differs between ground and space exposure in that plasma asher exposure produces isotropic arrival and space solar tracking produces sweeping arrival. Differences in initial impact reaction probability occur, dependent upon the energy and species existing in these environments. Due to the variations in ground-laboratory and space atomic oxygen, quantification of in-space performance based on plasma asher testing is not straightforward. The various atomic oxygen interactions that can occur with reflector surfaces, such as undercutting in organic substrates at protective coating defect sites, ground-laboratory techniques recommended for evaluating the atomic oxygen durability of reflectors based on asher exposures, and computational techniques which make use of ground-laboratory atomic oxygen exposure to predict in-space LEO durability are addressed.

Direct Data Distribution From Low-Earth Orbit

NASA Technical Reports Server (NTRS)

Budinger, James M.; Fujikawa, Gene; Kunath, Richard R.; Nguyen, Nam T.; Romanofsky, Robert R.; Spence, Rodney L.

1997-01-01

NASA Lewis Research Center (LeRC) is developing the space and ground segment technologies necessary to demonstrate a direct data distribution (1)3) system for use in space-to-ground communication links from spacecraft in low-Earth orbit (LEO) to strategically located tracking ground terminals. The key space segment technologies include a K-band (19 GHz) MMIC-based transmit phased array antenna, and a multichannel bandwidth- and power-efficient digital encoder/modulate with an aggregate data rate of 622 Mb/s. Along with small (1.8 meter), low-cost tracking terminals on the ground, the D3 system enables affordable distribution of data to the end user or archive facility through interoperability with commercial terrestrial telecommunications networks. The D3 system is applicable to both government and commercial science and communications spacecraft in LEO. The features and benefits of the D3 system concept are described. Starting with typical orbital characteristics, a set of baseline requirements for representative applications is developed, including requirements for onboard storage and tracking terminals, and sample link budgets are presented. Characteristics of the transmit array antenna and digital encoder/modulator are described. The architecture and components of the tracking terminal are described, including technologies for the next generation terminal. Candidate flights of opportunity for risk mitigation and space demonstration of the D3 features are identified.

Optimal reentry prediction of space objects from LEO using RSM and GA

NASA Astrophysics Data System (ADS)

Mutyalarao, M.; Raj, M. Xavier James

2012-07-01

The accurate estimation of the orbital life time (OLT) of decaying near-Earth objects is of considerable importance for the prediction of risk object re-entry time and hazard assessment as well as for mitigation strategies. Recently, due to the reentries of large number of risk objects, which poses threat to the human life and property, a great concern is developed in the space scientific community all over the World. The evolution of objects in Low Earth Orbit (LEO) is determined by a complex interplay of the perturbing forces, mainly due to atmospheric drag and Earth gravity. These orbits are mostly in low eccentric (eccentricity < 0.2) and have variations in perigee and apogee altitudes due to perturbations during a revolution. The changes in the perigee and apogee altitudes of these orbits are mainly due to the gravitational perturbations of the Earth and the atmospheric density. It has become necessary to use extremely complex force models to match with the present operational requirements and observational techniques. Further the re-entry time of the objects in such orbits is sensitive to the initial conditions. In this paper the problem of predicting re-entry time is attempted as an optimal estimation problem. It is known that the errors are more in eccentricity for the observations based on two line elements (TLEs). Thus two parameters, initial eccentricity and ballistic coefficient, are chosen for optimal estimation. These two parameters are computed with response surface method (RSM) using a genetic algorithm (GA) for the selected time zones, based on rough linear variation of response parameter, the mean semi-major axis during orbit evolution. Error minimization between the observed and predicted mean Semi-major axis is achieved by the application of an optimization algorithm such as Genetic Algorithm (GA). The basic feature of the present approach is that the model and measurement errors are accountable in terms of adjusting the ballistic coefficient

Copernicus POD Service: Orbit Determination of the Sentinel Satellites

NASA Astrophysics Data System (ADS)

Peter, Heike; Fernández, Jaime; Ayuga, Francisco; Féménias, Pierre

2016-04-01

The Copernicus POD (Precise Orbit Determination) Service is part of the Copernicus Processing Data Ground Segment (PDGS) of the Sentinel-1, -2 and -3 missions. A GMV-led consortium is operating the Copernicus POD Service being in charge of generating precise orbital products and auxiliary data files for their use as part of the processing chains of the respective Sentinel PDGS. Sentinel-1A was launched in April 2014 while Sentinel-2A was on June 2015 and both are routinely operated since then. Sentinel-3A is expected to be launched in February 2016 and Sentinel-1B is planned for spring 2016. Thus the CPOD Service will be operating three to four satellites simultaneously in spring 2016. The satellites of the Sentinel-1, -2, and -3 missions are all equipped with dual frequency high precision GPS receivers delivering the main observables for POD. Sentinel-3 satellites will additionally be equipped with a laser retro reflector for Satellite Laser Ranging and a receiver for DORIS tracking. All three types of observables (GPS, SLR and DORIS) will be used routinely for POD. The POD core of the CPOD Service is NAPEOS (Navigation Package for Earth Orbiting Satellites) the leading ESA/ESOC software for precise orbit determination. The careful selection of models and inputs is important to achieve the different but very demanding requirements in terms of orbital accuracy and timeliness for the Sentinel -1, -2 & -3 missions. The three missions require orbital products with various latencies from 30 minutes up to 20-30 days. The accuracy requirements are also different and partly very challenging, targeting 5 cm in 3D for Sentinel-1 and 2-3 cm in radial direction for Sentinel-3. Although the characteristics and the requirements are different for the three missions the same core POD setup is used to the largest extent possible. This strategy facilitates maintenance of the complex system of the CPOD Service. Updates in the dynamical modelling of the satellite orbits, e

Orbit Determination with Very Short Arcs: Admissible Regions

NASA Astrophysics Data System (ADS)

Gronchi, G. F.; Milani, A.; de'Michieli Vitturi, M.; Knezevic, Z.

2004-05-01

Contemporary observational surveys provide a huge number of detections of small solar system bodies, in particular of asteroids. These have to be reduced in real time in order to optimize the observational strategy and to select the targets for the follow-up and for the subsequent determination of an orbit. Typically, reported astrometry consists of few positions over a short time span, and this information is often not enough to compute a preliminary orbit and perform an identification. Classical methods for preliminary orbit determination based on three observations fail in such cases, and a new approach is necessary to cope with the problem. We introduce the concept of attributable, which is a vector composed by two angles and two angular velocities at a given time. It is then shown that the missing values (geocentric range and range rate), necessary for the computation of an orbit, can be constrained to a compact set that we call admissible region (AR). The latter is defined on the basis of requirements that the body belongs to the solar system, that it is not a satellite of the Earth, and that it is not a "shooting star" (very close and very small). A mathematical description of the AR is given, together with the proof of its topological properties: it turns out that the AR cannot have more than two connected components. A sampling of the AR can be performed by means of a Delaunay triangulation. A finite number of six-parameter sets of initial conditions are thus defined, with each node of triangulation representing a Virtual Asteroid for which it is possible to propagate the corresponding orbit and to predict ephemerides.

Modification of an impulse-factoring orbital transfer technique to account for orbit determination and maneuver execution errors

NASA Technical Reports Server (NTRS)

Kibler, J. F.; Green, R. N.; Young, G. R.; Kelly, M. G.

1974-01-01

A method has previously been developed to satisfy terminal rendezvous and intermediate timing constraints for planetary missions involving orbital operations. The method uses impulse factoring in which a two-impulse transfer is divided into three or four impulses which add one or two intermediate orbits. The periods of the intermediate orbits and the number of revolutions in each orbit are varied to satisfy timing constraints. Techniques are developed to retarget the orbital transfer in the presence of orbit-determination and maneuver-execution errors. Sample results indicate that the nominal transfer can be retargeted with little change in either the magnitude (Delta V) or location of the individual impulses. Additonally, the total Delta V required for the retargeted transfer is little different from that required for the nominal transfer. A digital computer program developed to implement the techniques is described.

James Webb Space Telescope Orbit Determination Analysis

NASA Technical Reports Server (NTRS)

Yoon, Sungpil; Rosales, Jose; Richon, Karen

2014-01-01

The James Webb Space Telescope (JWST) is designed to study and answer fundamental astrophysical questions from an orbit about the Sun-Earth/Moon L2 libration point, 1.5 million km away from Earth. This paper describes the results of an orbit determination (OD) analysis of the JWST mission emphasizing the challenges specific to this mission in various mission phases. Three mid-course correction (MCC) maneuvers during launch and early orbit phase and transfer orbit phase are required for the spacecraft to reach L2. These three MCC maneuvers are MCC-1a at Launch+12 hours, MCC-1b at L+2.5 days and MCC-2 at L+30 days. Accurate OD solutions are needed to support MCC maneuver planning. A preliminary analysis shows that OD performance with the given assumptions is adequate to support MCC maneuver planning. During the nominal science operations phase, the mission requires better than 2 cm/sec velocity estimation performance to support stationkeeping maneuver planning. The major challenge to accurate JWST OD during the nominal science phase results from the unusually large solar radiation pressure force acting on the huge sunshield. Other challenges are stationkeeping maneuvers at 21-day intervals to keep JWST in orbit around L2, frequent attitude reorientations to align the JWST telescope with its targets and frequent maneuvers to unload momentum accumulated in the reaction wheels. Monte Carlo analysis shows that the proposed OD approach can produce solutions that meet the mission requirements.

Gamma guidance of trajectories for coplanar, aeroassisted orbital transfer

NASA Technical Reports Server (NTRS)

Miele, A.; Wang, T.

1990-01-01

The optimization and guidance of trajectories for coplaner, aeroassisted orbital transfer (AOT) from high Earth orbit (HEO) to low Earth orbit (LEO) are examined. In particular, HEO can be a geosynchronous Earth orbit (GEO). It is assumed that the initial and final orbits are circular, that the gravitational field is central and is governed by the inverse square law, and that at most three impulses are employed: one at HEO exit, one at atmospheric exit, and one at LEO entry. It is also assumed that, during the atmospheric pass, the trajectory is controlled via the lift coefficient. The presence of upper and lower bounds on the lift coefficient is considered. First, optimal trajectories are computed by minimizing the total velocity impulse (hence, the propellant consumption) required for AOT transfer. The sequential gradient-restoration algorithm (SGRA) is used for optimal control problems. The optimal trajectory is shown to include two branches: a relatively short descending flight branch (branch 1) and a long ascending flight branch (branch 2). Next, attention is focused on guidance trajectories capable of approximating the optimal trajectories in real time, while retaining the essential characteristics of simplicity, ease of implementation, and reliability. For the atmospheric pass, a feedback control scheme is employed and the lift coefficient is adjusted according to a two-stage gamma guidance law. Further improvements are possible via a modified gamma guidance which is more stable with respect to dispersion effects arising from navigation errors, variations of the atmospheric density, and uncertainties in the aerodynamic coefficients than gamma guidance trajectory. A byproduct of the studies on dispersion effects is the following design concept. For coplaner aeroassisted orbital transfer, the lift-range-to-weight ratio appears to play a more important role than the lift-to-drag ratio. This is because the lift-range-to-weight ratio controls mainly the minimum

Expected orbit determination performance for the TOPEX/Poseidon mission

NASA Technical Reports Server (NTRS)

Nerem, R. S.; Putney, Barbara H.; Marshall, J. A.; Lerch, Francis J.; Pavlis, Erricos C.; Klosko, Steven M.; Luthcke, Scott B.; Patel, Girish B.; Williamson, Ronald G.; Zelensky, Nikita P.

1993-01-01

Each of the components required for the computation of precise orbits for the TOPEX/Poseidon (T/P) spacecraft - gravity field modeling, nonconservative force modeling, and satellite tracking technologies - is examined. The research conducted in the Space Geodesy Branch at Goddard Space Flight Center in preparation for meeting the 13-cm radial orbit accuracy requirement for the T/P mission is outlined. New developments in modeling the earth's gravitational field and modeling the complex nonconservative forces acting on T/P are highlighted. The T/P error budget is reviewed, and a prelaunch assessment of the predicted orbit determination accuracies is summarized.

PRIMA Platform capability for satellite missions in LEO and MEO (SAR, Optical, GNSS, TLC, etc.)

NASA Astrophysics Data System (ADS)

Logue, T.; L'Abbate, M.

2016-12-01

PRIMA (Piattaforma Riconfigurabile Italiana Multi Applicativa) is a multi-mission 3-axis stabilized Platform developed by Thales Alenia Space Italia under ASI contract.PRIMA is designed to operate for a wide variety of applications from LEO, MEO up to GEO and for different classes of satellites Platform Family. It has an extensive heritage in flight heritage (LEO and MEO Satellites already fully operational) in which it has successfully demonstrated the flexibility of use, low management costs and the ability to adapt to changing operational conditions.The flexibility and modularity of PRIMA provides unique capability to satisfy different Payload design and mission requirements, thanks to the utilization of recurrent adaptable modules (Service Module-SVM, Propulsion Module-PPM, Payload Module-PLM) to obtain mission dependent configuration. PRIMA product line development is continuously progressing, and is based on state of art technology, modular architecture and an Integrated Avionics. The aim is to maintain and extent multi-mission capabilities to operate in different environments (LEO to GEO) with different payloads (SAR, Optical, GNSS, TLC, etc.). The design is compatible with a wide range of European and US equipment suppliers, thus maximising cooperation opportunity. Evolution activities are mainly focused on the following areas: Structure: to enable Spacecraft configurations for multiple launch; Thermal Control: to guarantee thermal limits for new missions, more demanding in terms of environment and payload; Electrical: to cope with higher power demand (e.g. electrical propulsion, wide range of payloads, etc.) considering orbital environment (e.g. lighting condition); Avionics : AOCS solutions optimized on mission (LEO observation driven by agility and pointing, agility not a driver for GEO). Use of sensors and actuators tailored for specific mission and related environments. Optimised Propulsion control. Data Handling, SW and FDIR mission customization

LEO-to-ground polarization measurements aiming for space QKD using Small Optical TrAnsponder (SOTA).

PubMed

Carrasco-Casado, Alberto; Kunimori, Hiroo; Takenaka, Hideki; Kubo-Oka, Toshihiro; Akioka, Maki; Fuse, Tetsuharu; Koyama, Yoshisada; Kolev, Dimitar; Munemasa, Yasushi; Toyoshima, Morio

2016-05-30

Quantum communication, and more specifically Quantum Key Distribution (QKD), enables the transmission of information in a theoretically secure way, guaranteed by the laws of quantum physics. Although fiber-based QKD has been readily available since several years ago, a global quantum communication network will require the development of space links, which remains to be demonstrated. NICT launched a LEO satellite in 2014 carrying a lasercom terminal (SOTA), designed for in-orbit technological demonstrations. In this paper, we present the results of the campaign to measure the polarization characteristics of the SOTA laser sources after propagating from LEO to ground. The most-widely used property for encoding information in free-space QKD is the polarization, and especially the linear polarization. Therefore, studying its behavior in a realistic link is a fundamental step for proving the feasibility of space quantum communications. The results of the polarization preservation of two highly-polarized lasers are presented here, including the first-time measurement of a linearly-polarized source at λ = 976 nm and a circularly-polarized source at λ = 1549 nm from space using a realistic QKD-like receiver, installed in the Optical Ground Station at the NICT Headquarters, in Tokyo, Japan.

Selected Gravity Models in Terms of the fit to the GOCE Kinematic Orbit in the Dynamic Orbit Determination Process

NASA Astrophysics Data System (ADS)

Bobojć, Andrzej; Drożyner, Andrzej; Rzepecka, Zofia

2017-04-01

The work includes the comparison of performance of selected geopotential models in the dynamic orbit estimation of the satellite of the Gravity Field and Steady-State Ocean Circulation Explorer (GOCE) mission. This was realized by fitting estimated orbital arcs to the official centimeter-accuracy GOCE kinematic orbit which is provided by the European Space Agency. The Cartesian coordinates of kinematic orbit were treated as observations in the orbit estimation. The initial satellite state vector components were corrected in an iterative process with respect to the J2000.0 inertial reference frame using the given geopotential model, the models describing the remaining gravitational perturbations and the solar radiation pressure. Taking the obtained solutions into account, the RMS values of orbital residuals were computed. These residuals result from the difference between the determined orbit and the reference one - the GOCE kinematic orbit. The performance of selected gravity models was also determined using various orbital arc lengths. Additionally, the RMS fit values were obtained for some gravity models truncated at given degree and order of spherical harmonic coefficients. The advantage of using the kinematic orbit is its independence from any a priori dynamical models. For the research such GOCE-independent gravity models as HUST-Grace2016s, ITU_GRACE16, ITSG-Grace2014s, ITSG-Grace2014k, GGM05S, Tongji-GRACE01, ULUX_CHAMP2013S, ITG-GRACE2010S, EIGEN-51C, EIGEN5S, EGM2008 and EGM96 were adopted.

A comprehensive study of Mercury and MESSENGER orbit determination

NASA Astrophysics Data System (ADS)

Genova, Antonio; Mazarico, Erwan; Goossens, Sander; Lemoine, Frank G.; Neumann, Gregory A.; Nicholas, Joseph B.; Rowlands, David D.; Smith, David E.; Zuber, Maria; Solomon, Sean C.

2016-10-01

The MErcury, Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft orbited the planet Mercury for more than 4 years. The probe started its science mission in orbit around Mercury on 18 March 2011. The Mercury Laser Altimeter (MLA) and radio science system were the instruments dedicated to geodetic observations of the topography, gravity field, orientation, and tides of Mercury. X-band radio-tracking range-rate data collected by the NASA Deep Space Network (DSN) allowed the determination of Mercury's gravity field to spherical harmonic degree and order 100, the planet's obliquity, and the Love number k2.The extensive range data acquired in orbit around Mercury during the science mission (from April 2011 to April 2015), and during the three flybys of the planet in 2008 and 2009, provide a powerful dataset for the investigation of Mercury's ephemeris. The proximity of Mercury's orbit to the Sun leads to a significant perihelion precession attributable to the gravitational flattening of the Sun (J2) and the Parameterized Post-Newtonian (PPN) coefficients γ and β, which describe the space curvature produced by a unit rest mass and the nonlinearity in superposition of gravity, respectively. Therefore, the estimation of Mercury's ephemeris can provide crucial information on the interior structure of the Sun and Einstein's general theory of relativity. However, the high correlation among J2, γ, and β complicates the combined recovery of these parameters, so additional assumptions are required, such as the Nordtvedt relationship η = 4β - γ - 3.We have modified our orbit determination software, GEODYN II, to enable the simultaneous integration of the spacecraft and central body trajectories. The combined estimation of the MESSENGER and Mercury orbits allowed us to determine a more accurate gravity field, orientation, and tides of Mercury, and the values of GM and J2 for the Sun, where G is the gravitational constant and M is the solar mass

A Study on the Effects of J2 Perturbations on a Drag-Free Control System for Spacecraft in Low Earth Orbit

NASA Technical Reports Server (NTRS)

Vess, Melissa Fleck; Starin, Scott R.

2003-01-01

Low Earth Orbit (LEO) missions provide a unique means of gathering information about many of Earth s aspects such as climate, atmosphere, and gravitational field. Among the greatest challenges of LEO missions are designing, predicting, and maintaining the spacecraft orbit. The predominant perturbative forces acting on a spacecraft in LEO are J2 and higher order gravitational components, the effects of which are fairly easy to predict, and atmospheric drag, which causes the greatest uncertainty in predicting spacecraft ephemeris. The continuously varying atmospheric drag requires increased spacecraft tracking in order to accurately predict spacecraft location. In addition, periodic propulsive maneuvers typically must be planned and performed to counteract the effects of drag on the spacecraft orbit. If the effects of drag could be continuously and autonomously counteracted, the uncertainty in ephemeris due to atmospheric drag would essentially be eliminated from the spacecraft dynamics. One method of autonomous drag compensation that has been implemented on some missions is drag-free control. Drag-free control of a spacecraft was initially proposed in the 1960's and is discussed extensively by Lange. His drag-free control architecture consists of a free-floating proof mass enclosed within a spacecraft, isolating it from external disturbance forces such as atmospheric drag and solar radiation pressure. Under ideal conditions, internal disturbance forces can be ignored or mitigated, and the orbit of the proof mass depends only on gravitational forces. A sensor associated with the proof mass senses the movement of the spacecraft relative to the proof mass. Using the sensor measurements, the spacecraft is forced to follow the orbit of the proof mass by using low thrust propulsion, thus counteracting any non-gravitational disturbance forces. If the non-gravitational disturbance forces are successfully removed, the spacecraft s orbit will be affected only by well

Precise Orbit Determination of BeiDou Navigation Satellite System

NASA Astrophysics Data System (ADS)

He, Lina; Ge, Maorong; Wang, Jiexian; Wickert, Jens; Schuh, Harald

2013-04-01

China has been developing its own independent satellite navigation system since decades. Now the COMPASS system, also known as BeiDou, is emerging and gaining more and more interest and attention in the worldwide GNSS communities. The current regional BeiDou system is ready for its operational service around the end of 2012 with a constellation including five Geostationary Earth Orbit satellites (GEO), five Inclined Geosynchronous Orbit satellites (IGSO) and four Medium Earth orbit (MEO) satellites in operation. Besides the open service with positioning accuracy of around 10m which is free to civilian users, both precise relative positioning, and precise point positioning are demonstrated as well. In order to enhance the BeiDou precise positioning service, Precise Orbit Determination (POD) which is essential of any satellite navigation system has been investigated and studied thoroughly. To further improving the orbits of different types of satellites, we study the impact of network coverage on POD data products by comparing results from tracking networks over the Chinese territory, Asian-Pacific, Asian and of global scale. Furthermore, we concentrate on the improvement of involving MEOs on the orbit quality of GEOs and IGSOs. POD with and without MEOs are undertaken and results are analyzed. Finally, integer ambiguity resolution which brings highly improvement on orbits and positions with GPS data is also carried out and its effect on POD data products is assessed and discussed in detail. Seven weeks of BeiDou data from a ground tracking network, deployed by Wuhan University is employed in this study. The test constellation includes four GEO, five IGSO and two MEO satellites in operation. The three-day solution approach is employed to enhance its strength due to the limited coverage of the tracking network and the small movement of most of the satellites. A number of tracking scenarios and processing schemas are identified and processed and overlapping orbit

Laser Ranging for Effective and Accurate Tracking of Space Debris in Low Earth Orbits

NASA Astrophysics Data System (ADS)

Blanchet, Guillaume; Haag, Herve; Hennegrave, Laurent; Assemat, Francois; Vial, Sophie; Samain, Etienne

2013-08-01

The paper presents the results of preliminary design options for an operational laser ranging system adapted to the measurement of the distance of space debris. Thorough analysis of the operational parameters is provided with identification of performance drivers and assessment of enabling design options. Results from performance simulation demonstrate how the range measurement enables improvement of the orbit determination when combined with astrometry. Besides, experimental results on rocket-stage class debris in LEO were obtained by Astrium beginning of 2012, in collaboration with the Observatoire de la Côte d'Azur (OCA), by operating an experimental laser ranging system supported by the MéO (Métrologie Optique) telescope.

SPECS: Orbital debris removal

NASA Technical Reports Server (NTRS)

1991-01-01

The debris problem has reached a stage at which the risk to satellites and spacecraft has become substantial in low Earth orbit (LEO). This research discovered that small particles posed little threat to spacecraft because shielding can effectively prevent these particles from damaging the spacecraft. The research also showed that, even though collision with a large piece of debris could destroy the spacecraft, the large pieces of debris pose little danger because they can be tracked and the spacecraft can be maneuvered away from these pieces. Additionally, there are many current designs to capture and remove large debris particles from the space environment. From this analysis, it was decided to concentrate on the removal of medium-sized orbital debris, that is, those pieces ranging from 1 cm to 50 cm in size. The current design incorporates a transfer vehicle and a netting vehicle to capture the medium-sized debris. The system is based near an operational space station located at 28.5 deg inclination and 400 km altitude. The system uses ground-based tracking to determine the location of a satellite breakup or debris cloud. These data are uploaded to the transfer vehicle, which proceeds to rendezvous with the debris at a lower altitude parking orbit. Next, the netting vehicle is deployed, tracks the targeted debris, and captures it. After expending the available nets, the netting vehicle returns to the transfer vehicle for a new netting module and continues to capture more debris in the target area. Once all the netting modules are expended, the transfer vehicle returns to the space station's orbit where it is resupplied with new netting modules from a space shuttle load. The new modules are launched by the shuttle from the ground and the expended modules are taken back to Earth for removal of the captured debris, refueling, and repacking of the nets. Once the netting modules are refurbished, they are taken back into orbit for reuse. In a typical mission, the

Determination of Eros Physical Parameters for Near Earth Asteroid Rendezvous Orbit Phase Navigation

NASA Technical Reports Server (NTRS)

Miller, J. K.; Antreasian, P. J.; Georgini, J.; Owen, W. M.; Williams, B. G.; Yeomans, D. K.

1995-01-01

Navigation of the orbit phase of the Near Earth steroid Rendezvous (NEAR) mission will re,quire determination of certain physical parameters describing the size, shape, gravity field, attitude and inertial properties of Eros. Prior to launch, little was known about Eros except for its orbit which could be determined with high precision from ground based telescope observations. Radar bounce and light curve data provided a rough estimate of Eros shape and a fairly good estimate of the pole, prime meridian and spin rate. However, the determination of the NEAR spacecraft orbit requires a high precision model of Eros's physical parameters and the ground based data provides only marginal a priori information. Eros is the principal source of perturbations of the spacecraft's trajectory and the principal source of data for determining the orbit. The initial orbit determination strategy is therefore concerned with developing a precise model of Eros. The original plan for Eros orbital operations was to execute a series of rendezvous burns beginning on December 20,1998 and insert into a close Eros orbit in January 1999. As a result of an unplanned termination of the rendezvous burn on December 20, 1998, the NEAR spacecraft continued on its high velocity approach trajectory and passed within 3900 km of Eros on December 23, 1998. The planned rendezvous burn was delayed until January 3, 1999 which resulted in the spacecraft being placed on a trajectory that slowly returns to Eros with a subsequent delay of close Eros orbital operations until February 2001. The flyby of Eros provided a brief glimpse and allowed for a crude estimate of the pole, prime meridian and mass of Eros. More importantly for navigation, orbit determination software was executed in the landmark tracking mode to determine the spacecraft orbit and a preliminary shape and landmark data base has been obtained. The flyby also provided an opportunity to test orbit determination operational procedures that will be

Evaluation of advanced geopotential models for operational orbit determination

NASA Technical Reports Server (NTRS)

Radomski, M. S.; Davis, B. E.; Samii, M. V.; Engel, C. J.; Doll, C. E.

1988-01-01

To meet future orbit determination accuracy requirements for different NASA projects, analyses are performed using Tracking and Data Relay Satellite System (TDRSS) tracking measurements and orbit determination improvements in areas such as the modeling of the Earth's gravitational field. Current operational requirements are satisfied using the Goddard Earth Model-9 (GEM-9) geopotential model with the harmonic expansion truncated at order and degree 21 (21-by-21). This study evaluates the performance of 36-by-36 geopotential models, such as the GEM-10B and Preliminary Goddard Solution-3117 (PGS-3117) models. The Earth Radiation Budget Satellite (ERBS) and LANDSAT-5 are the spacecraft considered in this study.

AGB stars in Leo P and their use as metallicity probes

NASA Astrophysics Data System (ADS)

Lee, Chien-Hsiu

2016-09-01

Leo P is the most metal-poor yet star-forming galaxy in the local volume, and has the potential to serve as a local counterpart to interpret the properties of distant galaxies in the early universe. We present a comprehensive search of asymptotic giant branch (AGB) stars in Leo P using deep infrared imaging. AGB stars are the major dust contributors; the metal poor nature of Leo P can help to shed light on the dust formation process in very low-metallicity environments, similar to the early Universe. We select and classify oxygen-rich and carbon-rich candidate AGB stars using J - K versus K colour-magnitude diagram. To filter out contaminations from background galaxies, we exploit the high-resolution Hubble Space Telescope imaging and identify 9 oxygen-rich AGBs and 13 carbon-rich AGB stars in Leo P. We then use the ratio of carbon-rich and oxygen-rich AGB stars (C/M ratio) as an indicator of on-site metallicity and derive the global metallicity [Fe/H] = -1.8 dex for Leo P, in good agreement with previous studies using isochrone fitting. Follow-up observations of these Leo P AGB stars in the mid-infrared [e.g. Spitzer, James Webb Space Telescope (JWST)] will be invaluable to measure the dust formation rates using Spectral energy distribution (SED) fitting.

Threats to U.S. National Security Interests in Space: Orbital Debris Mitigation and Removal

DTIC Science & Technology

2014-01-08

objects larger than the size of a softball and hundreds of thousands of smaller fragments. This population of space debris potentially threatens U.S...catalogues objects as small as about 10 cm ( softball size) in LEO and as small as 1 meter in Geosynchronous Orbit.12 Today, the Space Surveillance

Low-Earth orbit effects on organic composite materials flown on LDEF

NASA Technical Reports Server (NTRS)

George, Pete E.; Dursch, Harry W.

1993-01-01

Over 35 different types of organic matrix composites were flown as part of 11 different experiments onboard the NASA Long Duration Exposure Facility (LDEF) satellite. This materials and systems experiment satellite flew in low-earth orbit (LEO) for 69 months. For that period, the experiments were subjected to the LEO environment including atomic oxygen (AO), ultraviolet (UV) radiation, thermal cycling, microvacuum, meteoroid and space debris (M&D), and particle radiation. Since retrieval of the satellite in January of 1990, the principal experiment investigators have been deintegrating, examining, and testing the materials specimens flown. The most detrimental environmental effect on all organic matrix composites was material loss due to AO erosion. AO erosion of uncoated organic matrix composites (OMC) facing the satellite ram direction was responsible for significant mechanical property degradations. Also, thermal cycling-induced microcracking was observed in some nonunidirectional reinforced OMC's. Thermal cycling and outgassing caused significant but predictable dimensional changes as measured in situ on one experiment. Some metal and metal oxide-based coatings were found to be very effective at preventing AO erosion of OMC's. However, M&D impacts and coating fractures which compromised these coatings allowed AO erosion of the underlying OMC substrates. The findings for organic matrix composites flown on the LDEF are summarized and the LEO environmental factors, their effects, and the influence on space hardware design factors for LEO applications are identified.

Comparison of Hyperthermal Ground Laboratory Atomic Oxygen Erosion Yields With Those in Low Earth Orbit

NASA Technical Reports Server (NTRS)

Banks, Bruce A.; Dill, Grace C.; Loftus, Ryan J.; deGroh, Kim K.; Miller, Sharon K.

2013-01-01

The atomic oxygen erosion yields of 26 materials (all polymers except for pyrolytic graphite) were measured in two directed hyperthermal radio frequency (RF) plasma ashers operating at 30 or 35 kHz with air. The hyperthermal asher results were compared with thermal energy asher results and low Earth orbital (LEO) results from the Materials International Space Station Experiment 2 and 7 (MISSE 2 and 7) flight experiments. The hyperthermal testing was conducted to a significant portion of the atomic oxygen fluence similar polymers were exposed to during the MISSE 2 and 7 missions. Comparison of the hyperthermal asher prediction of LEO erosion yields with thermal energy asher erosion yields indicates that except for the fluorocarbon polymers of PTFE and FEP, the hyperthermal energy ashers are a much more reliable predictor of LEO erosion yield than thermal energy asher testing, by a factor of four.

Study of multi-kW solar arrays for Earth orbit application

NASA Technical Reports Server (NTRS)

1980-01-01

Low cost low Earth orbit (LOW) and geosynchronous Earth orbit (GEO) Solar Array concepts in the 300 to 1000 kW range which could be reduced to hardware in the mid 1980's, are identified. Size scaling factors and longer life demands are recognized as the prime drivers for the designs if low life cycle costs for energy are to be achieved. Technology is identified which requires further development in order to assure component readiness and availability. Use of the low concentration ratio (CR) concentrator, which uses gallium arsenide solar cells for both LEO and GEO applications, is recommended.

A Deep Space Orbit Determination Software: Overview and Event Prediction Capability

NASA Astrophysics Data System (ADS)

Kim, Youngkwang; Park, Sang-Young; Lee, Eunji; Kim, Minsik

2017-06-01

This paper presents an overview of deep space orbit determination software (DSODS), as well as validation and verification results on its event prediction capabilities. DSODS was developed in the MATLAB object-oriented programming environment to support the Korea Pathfinder Lunar Orbiter (KPLO) mission. DSODS has three major capabilities: celestial event prediction for spacecraft, orbit determination with deep space network (DSN) tracking data, and DSN tracking data simulation. To achieve its functionality requirements, DSODS consists of four modules: orbit propagation (OP), event prediction (EP), data simulation (DS), and orbit determination (OD) modules. This paper explains the highest-level data flows between modules in event prediction, orbit determination, and tracking data simulation processes. Furthermore, to address the event prediction capability of DSODS, this paper introduces OP and EP modules. The role of the OP module is to handle time and coordinate system conversions, to propagate spacecraft trajectories, and to handle the ephemerides of spacecraft and celestial bodies. Currently, the OP module utilizes the General Mission Analysis Tool (GMAT) as a third-party software component for highfidelity deep space propagation, as well as time and coordinate system conversions. The role of the EP module is to predict celestial events, including eclipses, and ground station visibilities, and this paper presents the functionality requirements of the EP module. The validation and verification results show that, for most cases, event prediction errors were less than 10 millisec when compared with flight proven mission analysis tools such as GMAT and Systems Tool Kit (STK). Thus, we conclude that DSODS is capable of predicting events for the KPLO in real mission applications.

Hubble Peers into the Mouth of Leo A

NASA Image and Video Library

2017-12-08

At first glance, this NASA/ESA Hubble Space Telescope image seems to show an array of different cosmic objects, but the speckling of stars shown here actually forms a single body — a nearby dwarf galaxy known as Leo A. Its few million stars are so sparsely distributed that some distant background galaxies are visible through it. Leo A itself is at a distance of about 2.5 million light-years from Earth and a member of the Local Group of galaxies; a group that includes the Milky Way and the well-known Andromeda galaxy. Astronomers study dwarf galaxies because they are very numerous and are simpler in structure than their giant cousins. However, their small size makes them difficult to study at great distances. As a result, the dwarf galaxies of the Local Group are of particular interest, as they are close enough to study in detail. As it turns out, Leo A is a rather unusual galaxy. It is one of the most isolated galaxies in the Local Group, has no obvious structural features beyond being a roughly spherical mass of stars, and shows no evidence for recent interactions with any of its few neighbors. However, the galaxy’s contents are overwhelmingly dominated by relatively young stars, something that would normally be the result of a recent interaction with another galaxy. Around 90% of the stars in Leo A are less than eight billion years old — young in cosmic terms! This raises a number of intriguing questions about why star formation in Leo A did not take place on the “usual” timescale, but instead waited until it was good and ready. Image credit: ESA/Hubble & NASA; Acknowledgment: Judy Schmidt

Genetic Algorithm for Initial Orbit Determination with Too Short Arc (Continued)

NASA Astrophysics Data System (ADS)

Li, Xin-ran; Wang, Xin

2017-04-01

When the genetic algorithm is used to solve the problem of too short-arc (TSA) orbit determination, due to the difference of computing process between the genetic algorithm and the classical method, the original method for outlier deletion is no longer applicable. In the genetic algorithm, the robust estimation is realized by introducing different loss functions for the fitness function, then the outlier problem of the TSA orbit determination is solved. Compared with the classical method, the genetic algorithm is greatly simplified by introducing in different loss functions. Through the comparison on the calculations of multiple loss functions, it is found that the least median square (LMS) estimation and least trimmed square (LTS) estimation can greatly improve the robustness of the TSA orbit determination, and have a high breakdown point.

Analysis of orbit determination from Earth-based tracking for relay satellites in a perturbed areostationary orbit

NASA Astrophysics Data System (ADS)

Romero, P.; Pablos, B.; Barderas, G.

2017-07-01

Areostationary satellites are considered a high interest group of satellites to satisfy the telecommunications needs of the foreseen missions to Mars. An areostationary satellite, in an areoequatorial circular orbit with a period of 1 Martian sidereal day, would orbit Mars remaining at a fixed location over the Martian surface, analogous to a geostationary satellite around the Earth. This work addresses an analysis of the perturbed orbital motion of an areostationary satellite as well as a preliminary analysis of the aerostationary orbit estimation accuracy based on Earth tracking observations. First, the models for the perturbations due to the Mars gravitational field, the gravitational attraction of the Sun and the Martian moons, Phobos and Deimos, and solar radiation pressure are described. Then, the observability from Earth including possible occultations by Mars of an areostationary satellite in a perturbed areosynchronous motion is analyzed. The results show that continuous Earth-based tracking is achievable using observations from the three NASA Deep Space Network Complexes in Madrid, Goldstone and Canberra in an occultation-free scenario. Finally, an analysis of the orbit determination accuracy is addressed considering several scenarios including discontinuous tracking schedules for different epochs and different areoestationary satellites. Simulations also allow to quantify the aerostationary orbit estimation accuracy for various tracking series durations and observed orbit arc-lengths.

What's New for Laser Orbital Debris Removal

NASA Astrophysics Data System (ADS)

Phipps, Claude; Lander, Mike

2011-11-01

Orbital debris in low Earth orbit (LEO) are now sufficiently dense that the use of space is threatened by runaway collision cascading. A problem predicted more than thirty years ago, the threat from debris larger than about 1cm is now a reality that we ignore at our peril. The least costly, and most comprehensive, solution is Laser Orbital Debris Removal (LODR). In this approach, a high power pulsed laser on the Earth creates a laser-ablation jet on the debris object's surface which provides the small impulse required to cause it to re-enter and burn up in the atmosphere. The LODR system should be located near the Equator, and includes the laser, a large, agile mirror, and systems for active detection, tracking and atmospheric path correction. In this paper, we discuss advances that have occurred since LODR was first proposed, which make this solution to the debris problem look quite realistic.

Determination of Orbiter and Carrier Aerodynamic Coefficients from Load Cell Measurements

NASA Technical Reports Server (NTRS)

Glenn, G. M.

1976-01-01

A method of determining orbiter and carrier total aerodynamic coefficients from load cell measurements is required to support the inert and the captive active flights of the ALT program. A set of equations expressing the orbiter and carrier total aerodynamic coefficients in terms of the load cell measurements, the sensed dynamics of the Boeing 747 (carrier) aircraft, and the relative geometry of the orbiter/carrier is derived.

Using Onboard Telemetry for MAVEN Orbit Determination

NASA Technical Reports Server (NTRS)

Lam, Try; Trawny, Nikolas; Lee, Clifford

2013-01-01

Determination of the spacecraft state has been traditional done using radiometric tracking data before and after the atmosphere drag pass. This paper describes our approach and results to include onboard telemetry measurements in addition to radiometric observables to refine the reconstructed trajectory estimate for the Mars Atmosphere and Volatile Evolution Mission (MAVEN). Uncertainties in the Mars atmosphere models, combined with non-continuous tracking degrade navigation accuracy, making MAVEN a key candidate for using onboard telemetry data to help complement its orbit determination process.

Forecasting and remote sensing outer belt relativistic electrons from low Earth orbit

DOE PAGES

Chen, Yue; Reeves, Geoffrey D.; Cunningham, Gregory S.; ...

2016-02-15

Our study demonstrates the feasibility and reliability of using observations from low Earth orbit (LEO) to forecast and nowcast relativistic electrons in the outer radiation belt. Furthermore, we first report a high cross-energy, cross-pitch-angle coherence discovered between the trapped MeV electrons and precipitating approximately hundreds (~100s) of keV electrons—observed by satellites with very different altitudes—with correlation coefficients as high as ≳ 0.85. We then tested the feasibility of applying linear prediction filters to LEO data to predict the arrival of new MeV electrons during geomagnetic storms, as well as their evolving distributions afterward, based on the coherence. Reliability of thesemore » predictive filters is quantified by the performance efficiency with values as high as 0.74 when driven merely by LEO observations (or up to 0.94 with the inclusion of in situ MeV electron measurements). Finally, a hypothesis based upon the wave-particle resonance theory is proposed to explain the coherence, and a first-principle electron tracing model yields supporting evidence.« less

Evaluation of TDRSS-user orbit determination accuracy using batch least-squares and sequential methods

NASA Technical Reports Server (NTRS)

Oza, D. H.; Jones, T. L.; Hodjatzadeh, M.; Samii, M. V.; Doll, C. E.; Hart, R. C.; Mistretta, G. D.

1991-01-01

The development of the Real-Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination on a Disk Operating System (DOS) based Personal Computer (PC) is addressed. The results of a study to compare the orbit determination accuracy of a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOD/E with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), is addressed. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for the Earth Radiation Budget Satellite (ERBS); the maximum solution differences were less than 25 m after the filter had reached steady state.

A Cryogenic Propellant Production Depot for Low Earth Orbit

NASA Technical Reports Server (NTRS)

Potter, Seth D.; Henley, Mark; Guitierrez, Sonia; Fikes, John; Carrington, Connie; Smitherman, David; Gerry, Mark; Sutherlin, Steve; Beason, Phil; Howell, Joe (Technical Monitor)

2001-01-01

The cost of access to space beyond low Earth orbit can be lowered if vehicles can refuel in orbit. The power requirements for a propellant depot that electrolyzes water and stores cryogenic oxygen and hydrogen can be met using technology developed for space solar power. A propellant depot is described that will be deployed in a 400 km circular equatorial orbit, receive tanks of water launched into a lower orbit from Earth by gun launch or reusable launch vehicle, convert the water to liquid hydrogen and oxygen, and store Lip to 500 metric tonnes of cryogenic propellants. The propellant stored in the depot can support transportation from low Earth orbit to geostationary Earth orbit, the Moon, LaGrange points, Mars, etc. The tanks are configured in an inline gravity-gradient configuration to minimize drag and settle the propellant. Temperatures can be maintained by body-mounted radiators; these will also provide some shielding against orbital debris. Power is supplied by a pair of solar arrays mounted perpendicular to the orbital plane, which rotate once per orbit to track the Sun. In the longer term, cryogenic propellant production technology can be applied to a larger LEO depot, as well as to the use of lunar water resources at a similar depot elsewhere.

Geomagnetic transmission disturbances and heavy-ion fluences observed in low Earth orbit during the solar energetic particle events of October 1989.

PubMed

Boberg, P R; Tylka, A J; Adams, J H; Beahm, L P; Fluckiger, E O; Kleis, T; Kobel, E

1996-01-01

The large solar energetic particle (SEP) events and simultaneous large geomagnetic disturbances observed during October 1989 posed a significant, rapidly evolving space radiation hazard. Using data from the GOES-7, NOAA-10, IMP-8 and LDEF satellites, we determined the geomagnetic transmission, heavy ion fluences, mean Fe ionic charge state, and effective radiation hazard observed in low Earth orbit (LEO) for these SEPs. We modeled the geomagnetic transmission by tracing particles through the combination of the internal International Geomagnetic Reference Field (IGRF) and the Tsyganenko (1989) magnetospheric field models, extending the modeling to large geomagnetic disturbances. We used our results to assess the radiation hazard such very large SEP events would pose in the anticipated 52 degrees inclination space station orbit.

Requirements for an Advanced Low Earth Orbit (LEO) Sounder (ALS) for Improved Regional Weather Prediction and Monitoring of Greenhouse Gases

NASA Technical Reports Server (NTRS)

Pagano, Thomas S.; Chahine, Moustafa T.; Susskind, Joel

2008-01-01

Hyperspectral infrared atmospheric sounders (e.g., the Atmospheric Infrared Sounder (AIRS) on Aqua and the Infrared Atmospheric Sounding Interferometer (IASI) on Met Op) provide highly accurate temperature and water vapor profiles in the lower to upper troposphere. These systems are vital operational components of our National Weather Prediction system and the AIRS has demonstrated over 6 hrs of forecast improvement on the 5 day operational forecast. Despite the success in the mid troposphere to lower stratosphere, a reduction in sensitivity and accuracy has been seen in these systems in the boundary layer over land. In this paper we demonstrate the potential improvement associated with higher spatial resolution (1 km vs currently 13.5 km) on the accuracy of boundary layer products with an added consequence of higher yield of cloud free scenes. This latter feature is related to the number of samples that can be assimilated and has also shown to have a significant impact on improving forecast accuracy. We also present a set of frequencies and resolutions that will improve vertical resolution of temperature and water vapor and trace gas species throughout the atmosphere. Development of an Advanced Low Earth Orbit (LEO) Sounder (ALS) with these improvements will improve weather forecast at the regional scale and of tropical storms and hurricanes. Improvements are also expected in the accuracy of the water vapor and cloud properties products, enhancing process studies and providing a better match to the resolution of future climate models. The improvements of technology required for the ALS are consistent with the current state of technology as demonstrated in NASA Instrument Incubator Program and NOAA's Hyperspectral Environmental Suite (HES) formulation phase development programs.

Earth-to-Geostationary Orbit Transportation for Space Solar Power System Development