Search NASA⌕ Search

SEARCH · Search NASA

Results for “orbits”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Status of Precise Orbit Determination for Jason-2 Using GPS

The JASON-2 satellite, launched in June 2008, is the latest follow-on to the successful TOPEX/Poseidon (T/P) and JASON-I altimetry missions. JASON-2 is equipped with a TRSR Blackjack GPS dual-frequency receiver, a laser retroreflector array, and a DORIS receiver for precise orbit determination (POD). The most recent time series of orbits computed at NASA GSFC, based on SLR/DORIS data have been completed using both ITRF2005 and ITRF2008. These orbits have been shown to agree radially at 1 cm RMS for dynamic vs SLRlDORIS reduced-dynamic orbits and in comparison with orbits produced by other analysis centers (Lemoine et al., 2010; Zelensky et al., 2010; Cerri et al., 2010). We have recently upgraded the GEODYN software to implement model improvements for GPS processing. We describe the implementation of IGS standards to the Jason2 GEODYN GPS processing, and other dynamical and measurement model improvements. Our GPS-only JASON-2 orbit accuracy is assessed using a number of tests including analysis of independent SLR and altimeter crossover residuals, orbit overlap differences, and direct comparison to orbits generated at GSFC using SLR and DORIS tracking, and to orbits generated externally at other centers. Tests based on SLR and the altimeter crossover residuals provide the best performance indicator for independent validation of the NASAlGSFC GPS-only reduced dynamic orbits. For the ITRF2005 and ITRF2008 implementation of our GPS-only obits we are using the IGS05 and IGS08 standards. Reduced dynamic versus dynamic orbit differences are used to characterize the remaining force model error and TRF instability. We evaluate the GPS vs SLR & DORIS orbits produced using the GEODYN software and assess in particular their consistency radially and the stability of the altimeter satellite reference frame in the Z direction for both ITRF2005 and ITRF2008 as a proxy to assess the consistency of the reference frame for altimeter satellite POD.

Melachroinos, S.↗
NASA NTRS2011

Limiting Future Collision Risk to Spacecraft: An Assessment of NASA's Meteoroid and Orbital Debris Programs

Over the past 50 years, various NASA communities have contributed significantly to maturing NASA s meteoroid and orbital debris (MMOD)1 programs to their current state. As a result of these community efforts, and to NASA s credit, NASA s MMOD programs and models are now widely used and respected by the providers and users of both government and commercial satellites, nationally as well as internationally. Satellites have been redesigned to protect critical components from MMOD damage by moving critical components from exterior surfaces to deep inside a satellite s structure. Orbits are monitored and altered to minimize the risk of collision with tracked orbital debris. MMOD shielding added to the International Space Station (ISS) protects critical components and astronauts from potentially catastrophic damage that might result from smaller, untracked debris and meteoroid impacts. The space shuttle, as it orbited Earth, and whether docked to the ISS or not, was optimally oriented to protect its fragile thermal protection and thermal radiation systems from MMOD damage. In addition, astronauts inspected its thermal protection system for MMOD damage before the shuttle reentered Earth s atmosphere; Orion, NASA s capsule to carry astronauts to low Earth orbit, includes designs to mitigate the threat of MMOD damage and provide increased safety to the crew. When a handful of reasonable assumptions are used in NASA s MMOD models, scenarios are uncovered that conclude that the current orbital debris environment has already reached a "tipping point." That is, the amount of debris - in terms of the population of large debris objects, as well as overall mass of debris in orbit - currently in orbit has reached a threshold where it will continually collide with itself, further increasing the population of orbital debris. This increase will lead to corresponding increases in spacecraft failures, which will only create more feedback into the system, increasing the debris population growth rate. The increase thus far has been most rapid in low Earth orbit (LEO), with geosynchronous Earth orbits (GEOs) potentially suffering the same fate, but over a much longer time period. The exact timing and pace of this exponential growth are uncertain, but the serious implications of such a scenario require careful attention because of the strategic importance of U.S. space operations. The Office of Science and Technology Policy and the Office of Management and Budget contracted with the National Research Council for a study to perform three tasks: review NASA s MMOD programs and efforts, recommend in which of those NASA should increase or decrease its effort or change focus, and determine whether NASA should pursue work in any new MMOD areas. The official letter requesting the study and the full statement of task for the Committee for the Assessment of NASA s Orbital Debris Programs are in Appendixes A and B, respectively.

Source record↗
NASA NTRS2013

Dynamics of Orbits near 3:1 Resonance in the Earth-Moon System

The Interstellar Boundary Explorer (IBEX) spacecraft is currently in a highly elliptical orbit around Earth with a period near 3:1 resonance with the Moon. Its orbit is oriented so that apogee does not approach the Moon. Simulations show this orbit to be remarkably stable over the next twenty years. This article examines the dynamics of such orbits in the Circular Restricted 3-Body Problem (CR3BP). We look at three types of periodic orbits, each exhibiting a type of symmetry of the CR3BP. For each of the orbit types, we assess the local stability using Floquet analysis. Although not all of the periodic solutions are stable in the mathematical sense, any divergence is so slow as to produce practical stability over several decades. We use Poincare maps with twenty-year propagations to assess the nonlinear stability of the orbits, where the perturbation magnitudes are related to the orbit uncertainty for the IBEX mission. Finally we show that these orbits belong to a family of orbits connected in a bifurcation diagram that exhibits exchange of stability. The analysis of these families of period orbits provides a valuable starting point for a mission orbit trade study.

Dichmann, Donald J.↗
NASA NTRS2018

Dynamics of Planets Orbiting in the Alpha Centauri AB Stellar System

The Alpha Centauri system contains the Solar System's closest stellar neighbors. If an earth-like planet is present in the system, it could in principle be detected using a small space-based telescope. As Alpha Centauri is billions of years old, planets are only expected to be found in regions where their orbits are long-lived. We evaluate the extent of the regions within the Alpha Centauri AB system where small planets are able to orbit for billion-year timescales, as well as how closely-spaced planetary orbits can be within those regions in which individual planets can survive. Individual planets on nearly circular, coplanar orbits can survive throughout the habitable zones of both stars. However, perturbations from the companion star imply that the spacing of such planets in multi-planet systems must be significantly larger than the spacing of similar systems orbiting single stars in order to be long-lived. Because the binary companion induces a forced eccentricity upon the orbits of planets in orbit around either star, appropriately-phased circumstellar orbits with small initial eccentricities are stable to slightly larger initial semi-major axes than are initially circular orbits. Small initial eccentricities have a much larger effect on how closely planetary orbits can be spaced. Our results are of special interest as they can guide observers designing instrumentation and search strategies to attempt to discover planets orbiting the nearest sun-like stars. Results remain significantly higher than for planets orbiting single stars.

Lissauer, Jack J.↗
NASA NTRS2014

Orbit Determination Accuracy Analysis of the Magnetospheric Multiscale Mission During Perigee Raise

The Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) will provide orbit determination and prediction support for the Magnetospheric Multiscale (MMS) mission during the missions commissioning period. The spacecraft will launch into a highly elliptical Earth orbit in 2015. Starting approximately four days after launch, a series of five large perigee-raising maneuvers will be executed near apogee on a nearly every-other-orbit cadence. This perigee-raise operations concept requires a high-accuracy estimate of the orbital state within one orbit following the maneuver for performance evaluation and a high-accuracy orbit prediction to correctly plan and execute the next maneuver in the sequence. During early mission design, a linear covariance analysis method was used to study orbit determination and prediction accuracy for this perigee-raising campaign. This paper provides a higher fidelity Monte Carlo analysis using the operational COTS extended Kalman filter implementation that was performed to validate the linear covariance analysis estimates and to better characterize orbit determination performance for actively maneuvering spacecraft in a highly elliptical orbit. The study finds that the COTS extended Kalman filter tool converges on accurate definitive orbit solutions quickly, but prediction accuracy through orbits with very low altitude perigees is degraded by the unpredictability of atmospheric density variation.

Determination↗
NASA NTRS2014

Orbit Determination Accuracy Analysis of the Magnetospheric Multiscale Mission During Perigee Raise

The Goddard Space Flight Center (GSFC) Flight Dynamics Facility (FDF) will provide orbit determination and prediction support for the Magnetospheric Multiscale (MMS) mission during the mission's commissioning period. The spacecraft will launch into a highly elliptical Earth orbit in 2015. Starting approximately four days after launch, a series of five large perigee-raising maneuvers will be executed near apogee on a nearly every-other-orbit cadence. This perigee-raise operations concept requires a high-accuracy estimate of the orbital state within one orbit following the maneuver for performance evaluation and a high-accuracy orbit prediction to correctly plan and execute the next maneuver in the sequence. During early mission design, a linear covariance analysis method was used to study orbit determination and prediction accuracy for this perigee-raising campaign. This paper provides a higher fidelity Monte Carlo analysis using the operational COTS extended Kalman filter implementation that was performed to validate the linear covariance analysis estimates and to better characterize orbit determination performance for actively maneuvering spacecraft in a highly elliptical orbit. The study finds that the COTS extended Kalman filter tool converges on accurate definitive orbit solutions quickly, but prediction accuracy through orbits with very low altitude perigees is degraded by the unpredictability of atmospheric density variation.

Multiscale↗
NASA NTRS2015

Towards the 1 mm/y Stability of the Radial Orbit Error at Regional Scales

An estimated orbit error budget for the Jason-1 and Jason-2 GDR-D solutions is constructed, using several measures of orbit error. The focus is on the long-term stability of the orbit time series for mean sea level applications on a regional scale. We discuss various issues related to the assessment of radial orbit error trends; in particular this study reviews orbit errors dependent on the tracking technique, with an aim to monitoring the long-term stability of all available tracking systems operating on Jason-1 and Jason-2 (GPS, DORIS, SLR). The reference frame accuracy and its effect on Jason orbit is assessed. We also examine the impact of analysis method on the inference of Geographically Correlated Errors as well as the significance of estimated radial orbit error trends versus the time span of the analysis. Thus a long-term error budget of the 10-year Jason-1 and Envisat GDR-D orbit time series is provided for two time scales: interannual and decadal. As the temporal variations of the geopotential remain one of the primary limitations in the Precision Orbit Determination modeling, the overall accuracy of the Jason-1 and Jason-2 GDR-D solutions is evaluated through comparison with external orbits based on different time-variable gravity models. This contribution is limited to an East-West "order-1" pattern at the 2 mm/y level (secular) and 4 mm level (seasonal), over the Jason-2 lifetime. The possibility of achieving sub-mm/y radial orbit stability over interannual and decadal periods at regional scales and the challenge of evaluating such an improvement using in situ independent data is discussed.

Precision Orbit Determination (POD)↗
NASA NTRS1994

Dynamic orbit determination using GPS measurements from TOPEX/POSEIDON

The GPAS data acquired by the TOPEX/POSEIDON (T/P) Demonstration Receiver (DR) have been used in a dynamic orbit determination, which was based on the description of the gravitational and nongravitational forces in the equations of motion. The GPS carrier phase data were processed in a double difference mode to remove clock errors, including the effects of Selective Availability. Simultaneous estimation of the T/P orbit and GPS orbits was performed using five 10-day cycles in the interval between December (1992) and April (1993). The resulting T/P orbits have been compared with the orbits determined from Satellite Laser Ranging, the French one-way Doppler tracking system, DORIS, and with the JPL reduced dynamic orbit determination strategies and force models with the GPS/DR to those used with SLR/DORIS, the radial component of the T/P orbit (based on JGM-2) was found to agree better than 30 mm (rms) and 35 mm with the JPL reduced dynamic orbit. An experiment gravity tuning was accomplished using four cycles of GPS/DR data. The resulting GPS./DR-orbits, determined by the dynamic technique with the experimental gravity field, are in better agreement with the JPL reduced dynamic orbits in both the radial component (21-25 mm) and altimeter crossover residuals than the JGM-2 orbits. (21-25 mm) and altimeter crossover residuals than the JGM-2 orbits.

Schultz, B. E.↗
NASA NTRS1998

Lunar Prospector Orbit Determination Results

The orbit support for Lunar Prospector (LP) consists of three main areas: (1) cislunar orbit determination, (2) rapid maneuver assessment using Doppler residuals, and (3) routine mapping orbit determination. The cislunar phase consisted of two trajectory correction maneuvers during the translunar cruise followed by three lunar orbit insertion burns. This paper will detail the cislunar orbit determination accuracy and the real-time assessment of the cislunar trajectory correction and lunar orbit insertion maneuvers. The non-spherical gravity model of the Moon is the primary influence on the mapping orbit determination accuracy. During the first two months of the mission, the GLGM-2 lunar potential model was used. After one month in the mapping orbit, a new potential model was developed that incorporated LP Doppler data. This paper will compare and contrast the mapping orbit determination accuracy using these two models. LP orbit support also includes a new enhancement - a web page to disseminate all definitive and predictive trajectory and mission planning information. The web site provides definitive mapping orbit ephemerides including moon latitude and longitude, and four week predictive products including: ephemeris, moon latitude/longitude, earth shadow, moon shadow, and ground station view periods. This paper will discuss the specifics of this web site.

Beckman, Mark↗
NASA NTRS1996

Orbits Close to Asteroid 4769 Castalia

We use a radar-derived physical model of 4769 Castalia (1989 PB) to investigate close orbit dynamics around that kilometer- sized, uniformly rotating asteroid. Our methods of analysis provide a basis for systematic studies of particle dynamics close to any uniformly rotating asteroid. We establish that a Jacobi integral exists for particles orbiting this asteroid, examine the attendant zero-velocity surfaces, find families of periodic orbits, and determine their stability. All synchronous orbits and direct orbits within approx. 3 mean radii of Castalia are unstable and are subject to impact or escape from Castalia. Retrograde orbits are mostly stable and allow particles to orbit close to the asteroid surface. We derive a model which allows us to predict the escape conditions of a particle in orbit about Castalia and the (temporary) capture conditions for a hyperbolic interloper. Orbits within 1.5 km of Castalia are subject to immediate ejection from the system. Hyperbolic orbits with a V(sub infinity) less than 0.4 m/sec can potentially be captured by Castalia if their periapsis radius Is within approx. 2 km. For Castalia this capture region is small, but the results also apply to larger asteroids whose capture regions would also be larger. We determine bounds on ejecta speeds which either ensure ejecta escape or re-impact as functions of location on Castalia's surface. The speeds that ensure escape range from 0.28 to 0.84 m/sec and the speeds that ensure re-impact range from 0 to 0.18 m/sec. Speeds between these two bounds lead either to escape, re-impact, or potentially finite-time stable orbits. We develop a simple criterion which can establish whether a particle could have been ejected from the asteroid in the past and if it will impact the surface in the future.

Scheeres, D. J.↗
NASA NTRS1999

Orbit Determination for Mars Global Surveyor During Mapping

The Mars Global Surveyor (MGS) spacecraft reached a low-altitude circular orbit on February 4, 1999, after the termination of the second phase of aerobraking. The MGS spacecraft carries the Mars Orbiter Laser Altimeter (MOLA) whose primary goal is to derive a global, geodetically referenced 0.2 deg x 0.2 deg topographic grid of Mars with a vertical accuracy of better than 30 meters. During the interim science orbits in the' Hiatus mission phase (October - November 1997), and the Science Phasing Orbits (March - April, 1998, and June - July 1998) 208 passes of altimeter data were collected by the MOLA instrument. On March 1, 1999 the first ten orbits of MOLA altimeter data from the near-circular orbit were successfully returned from MGS by the Deep Space Network (DSN). Data will be collected from MOLA throughout the Mapping phase of the MCS mission, or for at least one Mars year (687 days). Whereas the interim orbits of Hiatus and SPO were highly eccentric, and altimeter data were only collected near periapsis when the spacecraft was below 785 km, the Mapping orbit of MGS is near circular, and altimeter data will be collected continuously at a rate of 10 Hz. The proper analysis of the altimeter data requires that the orbit of the MGS spacecraft be known to an accuracy comparable to that of the quality of the altimeter data. The altimeter has an ultimate precision of 30 cm on mostly flat surfaces, so ideally the orbits of the MGS spacecraft should be known to this level. This is a stringent requirement, and more realistic goals of orbit error for MGS are ten to thirty meters. In this paper we will discuss the force and measurement modelling required to achieve this objective. Issues in force modelling include the proper modelling of the gravity field of Mars, and the modelling of non-conservatives forces, including the development of a 'macro-model', in a similar fashion to TOPEX/POSEIDON and TDRSS. During Cruise and Aerobraking, the high gain antenna (HGA) was stowed on the +X face of the spacecraft. On March 29, 1999 the HGA will be deployed on a meter long boom which will remain Earth-pointing while the instrument panel (including the MOLA instrument) remains pointed at nadir. The tracking data must be corrected for the regular motion of the high gain antenna with respect to the center of mass, and the success of the MGS determination during Mapping will depend on correctly accounting for this offset in the measurement model.

Lemoine, F. G.↗
NASA NTRS1999

Mars Orbit Rendezvous Strategy for the Mars 2003/2005 Sample Return Mission

The primary objective of the Mars 2003/2005 Sample Return Project is to return Martian surface materials to Earth from two different sites by the year 2008. The baseline mission plan relies heavily on the use of a Mars orbit rendezvous strategy similar to the lunar orbit rendezvous scheme used for the Apollo missions. The 2003 mission consists of a single spacecraft comprised of a Lander, Rover, and Mars ascent vehicle (MAV). The 2003 mission will be launched on a Delta-III-class launch vehicle in May/June 2003 and arrive at Mars in December 2003/January 2004. The Lander deploys the Rover to collect surface samples from several sites and return them to the Lander where they are transferred to a sample canister onboard the MAV. The MAV is launched into a low Mars orbit (targeted for 600 km circular, 45 deg inclination) and releases the sample canister to await retrieval by an Orbiter launched in 2005. (The sample canister is a passive vehicle with no maneuvering capability.) The duration of Mars surface operations is at most about 90 days. The 2005 mission consists of two separate spacecraft: a Lander/Rover/MAV spacecraft identical to that used for the 2003 mission and an Orbiter carrying an Earth Entry Vehicle (EEV). Both spacecraft will be launched on a single Ariane-5 in August 2005 and arrive at Mars in July/August 2006. A second sample canister is delivered to Mars orbit using the same scenario as was used for the 2003 mission. The Orbiter uses aerocapture for insertion into Mars orbit (targeted for 250 x 1400 km, 45 deg inclination). During its approximately one-year stay at Mars, the Orbiter will search for and attempt to rendezvous first with the 2003 sample canister and then with the 2005 sample canister. After retrieval, each sample canister is transferred to the EEV. The Orbiter departs Mars in July 2007 and returns to Earth in October 2008 on a trajectory targeted for landing at the Utah Test and Training Range (UTTR). After deploying the EEV, the Orbiter performs a deflection maneuver to avoid reentry into Earth's atmosphere.

DAmario, Louis A.↗
NASA NTRS1999

Man-Made Debris In and From Lunar Orbit

During 1966-1976, as part of the first phase of lunar exploration, 29 manned and robotic missions placed more than 40 objects into lunar orbit. Whereas several vehicles later successfully landed on the Moon and/or returned to Earth, others were either abandoned in orbit or intentionally sent to their destruction on the lunar surface. The former now constitute a small population of lunar orbital debris; the latter, including four Lunar Orbiters and four Lunar Module ascent stages, have contributed to nearly 50 lunar sites of man's refuse. Other lunar satellites are known or suspected of having fallen from orbit. Unlike Earth satellite orbital decays and deorbits, lunar satellites impact the lunar surface unscathed by atmospheric burning or melting. Fragmentations of lunar satellites, which would produce clouds of numerous orbital debris, have not yet been detected. The return to lunar orbit in the 1990's by the Hagoromo, Hiten, Clementine, and Lunar Prospector spacecraft and plans for increased lunar exploration early in the 21st century, raise questions of how best to minimize and to dispose of lunar orbital debris. Some of the lessons learned from more than 40 years of Earth orbit exploitation can be applied to the lunar orbital environment. For the near-term, perhaps the most important of these is postmission passivation. Unique solutions, e.g., lunar equatorial dumps, may also prove attractive. However, as with Earth satellites, debris mitigation measures are most effectively adopted early in the concept and design phase, and prevention is less costly than remediation.

Johnson, Nicholas L.↗
NASA NTRS2002

Orbit Determination Support for the Microwave Anisotropy Probe (MAP)

The Microwave Anisotropy Probe (MAP) is a Medium Class Explorers (MIDEX) mission produced in partnership between Goddard Space Flight Center (GSFC) and Princeton University. The main science objective of the MAP mission is to produce an accurate full-sky map of the cosmic microwave background temperature fluctuations anisotropy. MAP was launched from the Cape Canaveral Air Force Station Complex 17 aboard a Delta II 7425-10 expendable launch vehicle at exactly 19:46:46.183 UTC on June 30, 2001. The spacecraft received a nominal direct insertion by the Delta into a 185 km circular orbit. MAP was then maneuvered into a sequence of phasing loops designed to set up a lunar swingby (gravity-assisted acceleration) of the spacecraft onto a transfer trajectory to a Lissajous orbit about the Earth-Sun L2 point. The mission duration is approximately 27 months with 3 to 4 months of transfer time to the final mission orbit about L2. The MAP transfer orbit consisted of 3.5 phasing loops: the first loop has a 7-day period, the second and third loops have a 9-day period, and the last half loop has a 4-day period as illustrated in Figure 1, which also indicates the placement of maneuvers. A Pfinal correction maneuver was performed 18 hours after the last perigee to more closely achieve the targeted lissajous orbit. The lunar encounter or swingby took place approximately 30 days after launch. After the lunar encounter, the spacecraft will cruise for approximately 120 days before it arrives at L2. A Mid-Course Correction (MCC) maneuver was executed seven days after the swingby to further refine the trajectory. Once the MAP satellite is injected into the L2 Lissajous orbit, it will perform occasional stationkeeping maneuvers to maintain the Lissajous orbit for a minimum of two years (and a goal of four years). Because of its complex orbital characteristics, the mission provided a unique challenge to orbit determination (OD) support in many orbital regimes. Extensive trajectory error covariance analysis was performed to predict ephemeris accuracy for the OD process using a Bayesian least-squares technique. The orbit determination error analysis is essential for maneuver planning and maneuver recovery study. Several tracking scenarios were investigated for each phase of the mission. This paper provides a summary of the premission trajectory covariance error analysis, as well as actual real-time OD results. The use and impact of the various tracking stations, systems, and measurements will be discussed. Details of the operational OD support and the inferred OD accuracy will be presented, and the results will be compared to the premission covariance analysis, In addition, there will be a discussion of the challenges presented to OD support including delta-Vs at apogee as well as perigee, and effects due to spacecraft attitude mode, in light of their implications to the OD accuracy and covariance analysis.

Truong, Son H.↗
NASA NTRS2006

TRMM On-Orbit Performance Re-Accessed After Control Change

The Tropical Rainfall Measuring Mission (TRMM) spacecraft, a joint mission between the U.S. and Japan, launched onboard an HI1 rocket on November 27,1997 and transitioned in August, 2001 from an average operating altitude of 350 kilometers to 402.5 kilometers. Due to problems using the Earth Sensor Assembly (ESA) at the higher altitude, TRMM switched to a backup attitude control mode. Prior to the orbit boost TRMM controlled pitch and roll to the local vertical using ESA measurements while using gyro data to propagate yaw attitude between yaw updates from the Sun sensors. After the orbit boost, a Kalman filter used 3-axis gyro data with Sun sensor and magnetometers to estimate onboard attitude. While originally intended to meet a degraded attitude accuracy of 0.7 degrees, the new control mode met the original 0.2 degree attitude accuracy requirement after improving onboard ephemeris prediction and adjusting the magnetometer calibration onboard. Independent roll attitude checks using a science instrument, the Precipitation Radar (PR) which was built in Japan, provided a novel insight into the pointing performance. The PR data helped identify the pointing errors after the orbit boost, track the performance improvements, and show subtle effects from ephemeris errors and gyro bias errors. It also helped identify average bias trends throughout the mission. Roll errors tracked by the PR from sample orbits pre-boost and post-boost are shown in Figure 1. Prior to the orbit boost the largest attitude errors were due to occasional interference in the ESA. These errors were sometime larger than 0.2 degrees in pitch and roll, but usually less, as estimated from a comprehensive review of the attitude excursions using gyro data. Sudden jumps in the onboard roll show up as spikes in the reported attitude since the control responds within tens of seconds to null the pointing error. The PR estimated roll tracks well with an estimate of the roll history propagated using gyro data. After the orbit boost, the attitude errors shown by the PR roll have a smooth sine-wave type signal because of the way that attitude errors propagate with the use of gyro data. Yaw errors couple at orbit period to roll with '/4 orbit lag. By tracking the amplitude, phase, and bias of the sinusoidal PR roll error signal, it was shown that the average pitch rotation axis tends to be offset from orbit normal in a direction perpendicular to the Sun direction, as shown in Figure 2 for a 200 day period following the orbit boost. This is a result of the higher accuracy and stability of the Sun sensor measurements relative to the magnetometer measurements used in the Kalman filter. In November, 2001 a magnetometer calibration adjustment was uploaded which improved the pointing performance, keeping the roll and yaw amplitudes within about 0.1 degrees. After the boost, onboard ephemeris errors had a direct effect on the pitch pointing, being used to compute the Earth pointing reference frame. Improvements after the orbit boost have kept the the onboard ephemeris errors generally below 20 kilometers. Ephemeris errors have secondary effects on roll and yaw, especially during high beta angle when pitch effects can couple into roll and yaw. This is illustrated in figure 3. The onboard roll bias trends as measured by PR data show correlations with the Kalman filter's gyro bias error. This particularly shows up after yaw turns (every 2 to 4 weeks) as shown in Figure 3, when a slight roll bias is observed while the onboard computed gyro biases settle to new values. As for longer term trends, the PR data shows that the roll bias was influenced by Earth horizon radiance effects prior to the boost, changing values at yaw turns, and indicated a long term drift as shown in Figure 4. After the boost, the bias variations were smaller and showed some possible correlation with solar beta angle, probably due to sun sensor misalignment effects.

Bilanow, Steve↗
NASA NTRS2003

Mars orbits with daily repeating ground traces

This paper derives orbits at Mars with ground traces that repeat at the same times every solar day (sol). A relay orbiter in such an orbit would pass over insitu probes at the same times every sol, ensuring consistent coverage and simplifying mission design and operations. 42 orbits in five classes are characteried: 14 cicular equatorial prograde orbits; 14 circular equatorial retrograde orbits; 11 circular sun synchrounous orbits; 2 eccentroc equatorial orbits; 1 eccentric critcally inclined orbit. the paper reports on the performance of a relay orbiter in some of the orbits.

Mars↗
NASA NTRS1992

Method of delivering lunar generated fluid to earth orbit using an external tank

A method and apparatus are provided for delivering lunar generated fluid to Earth orbit from lunar orbit. Transport takes place in an external tank of a shuttle which has been suitably outfitted in Earth orbit for reusable travel between Earth orbit and a lunar orbit. The outfitting of the external tank includes the adding of an engine, an electrical system, a communication system, a guidance system, an aerobraking device, and a plurality of interconnected fluid storage tanks to the hydrogen and oxygen tanks of the external tank. The external tank is then propelled to lunar orbit the first time using Earth-based propellant. In lunar orbit, the storage tanks are filled with the lunar generated fluid with the remainder tank volumes filled with lunar generated liquid oxygen and hydrogen which serve as propellants for returning the tank to Earth orbit where the fluid is off-loaded. The remaining lunar generated oxygen and hydrogen is then sufficient to return the external tank to lunar orbit so that a subsequent cycle of fluid delivery is repeated. A space station in a higher Earth orbit is preferably used to outfit the external tank, and a lunar node in lunar orbit is used to store and transfer the fluid and liquid oxygen and hydrogen to the external tank. The lunar generated fluid is preferably .sup.3 He.

Butterfield, Ansel J.↗
NASA NTRS2009

Orbital Debris Quarterly News, Volume 13, Issue 4

Although NASA has conducted research on orbital debris since the 1960s, the NASA Orbital Debris Program Office is now considered to have been established in October 1979, following the recognition by senior NASA officials of orbital debris as a space environmental issue and the allocation by NASA Headquarters Advanced Programs Office to the Lyndon B. Johnson Space Center (JSC) of funds specifically dedicated for orbital debris investigations. In the 30 years since, the NASA Orbital Debris Program Office has pioneered the characterization of the orbital debris environment and its potential effects on current and future space systems, has developed comprehensive orbital debris mitigation measures, and has led efforts by the international aerospace community in addressing the challenges posed by orbital debris. In 1967 the Flight Analysis Branch at the Manned Spacecraft Center (renamed the Lyndon B. Johnson Space Center in 1973) evaluated the risks of collisions between an Apollo spacecraft and orbital debris. Three years later the same group calculated collision risks for the forthcoming Skylab space station, which was launched in 1973. By 1976, the nucleus of NASA s yet-to-be-formed orbital debris research efforts, including Andrew Potter, Burton Cour-Palais, and Donald Kessler, was found in JSC s Environmental Effects Office, examining the potential threat of orbital debris to large space platforms, in particular the proposed Solar Power Satellites (SPS).

Liou, Jer-Chyi↗