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NASA NTRS2016

Application of X-Ray Pulsar Navigation: A Characterization of the Earth Orbit Trade Space

The potential for pulsars as a navigation source has been studied since their discovery in 1967. X-ray pulsar navigation (XNAV) is a celestial navigation system that uses the consistent timing nature of x-ray photons from milli-second pulsars (MSP) to perform space navigation. By comparing the detected arrival of x-ray photons to a reference database of expected pulsar lightcurve timing models, one can infer a range and range rate measurement based on light time delay. Much of the challenge of XNAV comes from the faint signal, availability, and distant nature of pulsars. This is a study of potential pulsar XNAV measurements to measure extended Kalman filter (EKF) tracking performance with a wide trade space of bounded Earth orbits, using a simulation of existing x-ray detector space hardware. An example of an x-ray detector for XNAV is the NASA Station Explorer for X-ray Timing and Navigation (SEXTANT) mission, a technology demonstration of XNAV set to perform on the International Space Station (ISS) in late 2016early 2017. XNAV hardware implementation is driven by trajectory and environmental influences which add noise to the x-ray pulse signal. In a closed Earth orbit, the radiation environment can exponentially increase the signal noise from x-ray pulsar sources, decreasing the quality and frequency of measurements. The SEXTANT mission in particular improves on the signal to noise ratio by focusing an array of 56 x-ray silicon drift detectors at one pulsar target at a time. This reduces timing glitches and other timing noise contributions from ambient x-ray sources to within a 100 nanosecond resolution. This study also considers the SEXTANT scheduling challenges inherent in a single target observation. Finally, as the navigation sources are now relatively inertial targets, XNAV measurements are also subject to periods of occultation from various celestial bodies. This study focuses on the characterization of these drivers in closed Earth orbits and is not a tuning analysis of the EKF. The study shows that the closed Earth orbit for XNAV performance is reliant on the orbit semi-major axis and eccentricity as well as orbit inclination. These parameters are the primary drivers of pulsar measurement availability and significantly influence the natural spacecraft orbit dynamics. Sensitivity to initial orbit determination error growth due to the scarcity of XNAV measurements within an orbital period require appropriate timing of initial XNAV measurements. The orbit angles of argument of perigee and right ascension of the ascending node, alongside the other orbit parameters, complete the initial cadence of XNAV measurements. The performance of initial XNAV measurements then propagates throughout the experimental period. The study provides a basis to missions who wish to consider XNAV as a potential navigation source in a closed Earth orbit design. It provides a family of orbit trajectories as well as other modeling considerations needed to effectively evaluate if XNAV is an effective navigation source for a potential mission. As an EKF is sensitive to a linearized estimated state, this study has a direct benefit of providing effective XNAV measurements to maintain spacecraft tracking, independent of other navigation sources. In the particular use case of the SEXTANT mission, it also provides a novel scheduling algorithm which addresses the need to prioritize and manage pulsar observations for effective navigation.

PULSAR↗
NASA NTRS2016

Magnetospheric Multiscale Mission Attitude Dynamics: Observations from Flight Data

Extensive flight data is being collected throughout the MMS mission that includes quantities that are of interest for attitude dynamics studies such as spin rate, spin axis orientation nutation rate, etc. One example of such data is the long-term evolution of the spin rates of the four spacecraft. Spikes in these rates are observed that are separated by the MMS orbital period (just under 24 hr) and occur around perigee due to gravity-gradient torque. Periodic discontinuities in spin rate are caused by the controller resetting the spin rate approximately to the nominal 3.1 RPM value at the time of each maneuver. In between, a slow decay in spin rate can be seen to occur. The paper will discuss various disturbance torque mechanisms that could potentially be responsible for this behavior: these include magnetic hysteresis, eddy currents, solar radiation pressure, and a possible interaction between gravity-gradient and wire boom flexibility effects. One additional disturbance mechanism is produced by the Active Spacecraft Potential Control (ASPOC) devices: these emit positive indium ions to keep the MMS spacecraft electrically neutral, so as not to corrupt the electric field observations that are made by some of the on-board instruments. The spin rate decays that could be produced by these various mechanisms will be quantified in the paper, and their signatures described. Comparing these with the observations from flight data then allow the most likely candidate to be determined.

DYNAMICS↗
NASA NTRS2016

Application of X-Ray Pulsar Navigation: A Characterization of the Earth Orbit Trade Space

The potential for pulsars as a navigation source has been studied since their discovery in 1967. X-ray pulsar navigation (XNAV) is a celestial navigation system that uses the consistent timing nature of x-ray photons from millisecond pulsars (MSP) to perform space navigation. By comparing the detected arrival of x-ray photons to a reference database of expected pulsar light-curve timing models, one can infer a range and range rate measurement based on light time delay. Much of the challenge of XNAV comes from the faint signal, availability, and distant nature of pulsars. This is a study of potential pulsar XNAV measurements to measure extended Kalman filter (EKF) tracking performance with a wide trade space of bounded Earth orbits, using a simulation of existing x-ray detector space hardware. An example of an x-ray detector for XNAV is the NASA Station Explorer for X-ray Timing and Navigation (SEXTANT) mission, a technology demonstration of XNAV set to perform on the International Space Station (ISS) in late 2016early 2017. XNAV hardware implementation is driven by trajectory and environmental influences which add noise to the x-ray pulse signal. In a closed Earth orbit, the radiation environment can exponentially increase the signal noise from x-ray pulsar sources, decreasing the quality and frequency of measurements. The SEXTANT mission in particular improves on the signal to noise ratio by focusing an array of 56 x-ray silicon drift detectors at one pulsar target at a time. This reduces timing glitches and other timing noise contributions from ambient x-ray sources to within a 100 nanosecond resolution. This study also considers the SEXTANT scheduling challenges inherent in a single target observation. Finally, as the navigation sources are now relatively inertial targets, XNAV measurements are also subject to periods of occultation from various celestial bodies. This study focuses on the characterization of these drivers in closed Earth orbits and is not a tuning analysis of the EKF. The study shows that the closed Earth orbit for XNAV performance is reliant on the orbit semi-major axis and eccentricity as well as orbit inclination. These parameters are the primary drivers of pulsar measurement availability and significantly influence the natural spacecraft orbit dynamics. Sensitivity to initial orbit determination error growth due to the scarcity of XNAV measurements within an orbital period require appropriate timing of initial XNAV measurements. The orbit angles of argument of perigee and right ascension of the ascending node, alongside the other orbit parameters, complete the initial cadence of XNAV measurements. The performance of initial XNAV measurements then propagates throughout the experimental period. The study provides a basis to missions who wish to consider XNAV as a potential navigation source in a closed Earth orbit design. It provides a family of orbit trajectories as well as other modeling considerations needed to effectively evaluate if XNAV is an effective navigation source for a potential mission. As an EKF is sensitive to a linearized estimated state, this study has a direct benefit of providing effective XNAV measurements to maintain spacecraft tracking, independent of other navigation sources. In the particular use case of the SEXTANT mission, it also provides a novel scheduling algorithm which addresses the need to prioritize and manage pulsar observations for effective navigation.

x-ray↗
NASA NTRS2016

Optimizing Mars Sphere of Influence Maneuvers for NASA's Evolvable Mars Campaign

NASA's Human Spaceflight Architecture Team is refining human exploration architectures that will extend human presence to the Martian surface. For both Mars orbital and surface missions, NASA's Evolvable Mars Campaign assumes that cargo and crew can be delivered repeatedly to the same destination. Up to this point, interplanetary trajectories have been optimized to minimize the total propulsive requirements of the in-space transportation systems, while the pre-deployed assets and surface systems are optimized to minimize their respective propulsive requirements separate from the in-space transportation system. There is a need to investigate the coupled problem of optimizing the interplanetary trajectory and optimizing the maneuvers within Mars's sphere of influence. This paper provides a description of the ongoing method development, analysis and initial results of the effort to resolve the discontinuity between the interplanetary trajectory and the Mars sphere of influence trajectories. Assessment of Phobos and Deimos orbital missions shows the in-space transportation and crew taxi allocations are adequate for missions in the 2030s. Because the surface site has yet to be selected, the transportation elements must be sized to provide enough capability to provide surface access to all landing sites under consideration. Analysis shows access to sites from elliptical parking orbits with a lander that is designed for sub-periapsis landing location is either infeasible or requires expensive orbital maneuvers for many latitude ranges. In this case the locus of potential arrival perigee vectors identifies the potential maximum north or south latitudes accessible. Higher arrival velocities can decrease reorientation costs and increase landing site availability. Utilizing hyperbolic arrival and departure vectors in the optimization scheme will increase transportation site accessibility and provide more optimal solutions.

Merrill, Raymond G.↗
NASA NTRS2017

Minimum dV for Targeted Spacecraft Disposal

The density scale height of the Earth's atmosphere undergoes significant reduction in the final phases of a natural decay. It can be shown that for most realistic ballistic numbers, it is possible to exploit this effect to amplify available spacecraft dV by using it at the penultimate perigee to penetrate into higher drag regions at final apogee. The drag at this lower pass can more effectively propel a spacecraft towards the final target region than applying the same dV direct Hohmann transfer at that final apogee. This study analyzes the potential use of this effect-- in combination with small phasing burns--to calculate the absolute minimum delta-V that would be required to reliably guide a spacecraft to any specified safe unoccupied ocean region as a function of ballistic number, orbit inclination, and initial eccentricity. This calculation is made for controllable spacecraft in several orbit inclinations and eccentricities with arbitrary initial LAN and ArgP one week before final entry, under three-sigma atmospheric perturbations. The study analyzes the dV required under varying levels of final controllable altitude at which dV may be imparted, and various definitions of the length and location of a "safe" disposal area. The goal of such research is to improve public safety by creating assured safe disposal strategies for low-dV and/or low-thrust spacecraft that under more traditional strategies would need to be abandoned to a fully random decay.

Bacon, John↗
NASA NTRS2016

Trending of SNPP Ephemeris and Its Implications on VIIRS Geometric Performance

This paper describes trends in the Suomi National Polar-orbiting Partnership (SNPP) spacecraft ephemeris data over the four and half years of on-orbit operations. It then discusses the implications of these trends on the geometric performance of the Visible Infrared Imaging Radiometer Suite (VIIRS), one of the instruments onboard SNPP. The SNPP ephemeris data includes time stamped spacecraft positions and velocities that are used to calculate the spacecraft altitude and sub-satellite locations. Through drag make-up maneuvers (DMUs) the orbital mean altitude (spacecraft altitude averaged over an orbit) has been maintained at 838.8 kilometers to within plus or minus 0.2 kilometers and the orbital period at 101.5 minutes to within plus or minus 0.2 seconds. The corresponding orbital mean velocity in the terrestrial frame of reference has been maintained at 7524 meters per second to within plus or minus 0.5 meters per second. Within an orbit, the altitude varies from 828 kilometers near 15 degrees North latitude to 856 kilometers near the South Pole. Inclination adjust maneuvers (IAMs) have maintained the orbit inclination angle at 98.67 degrees to within plus or minus 0.07 degrees and the sun-synchronous local time at ascending node (LTAN) at 13:28 to within plus or minus 5 minutes. Besides these trends, it is interesting to observe that the orbit's elliptic shape has its major axis linking the perigee and apogee shorter than the line linking the ascending node and the descending node. This effect is caused by the Earth's oblate spheroid shape and deviates from a Keplerian orbit theory in which the two orbiting bodies are point masses. VIIRS has 5 imagery resolution bands, 16 moderate resolution bands and a day-night band, with 32, 16 and 16 detectors, respectively, aligned in the spacecraft flight (also known as track) direction. For each band's sample within a scan, the detectors sample the Earth's surface simultaneously in the track direction in the Earth Centered Inertial frame of reference. The distance between the center of the area sensed by the trailing detectors of one scan and the leading detectors of the next includes a component caused by earth rotation. This earth rotation component is relatively small (approximately 70 meters per second) for an orbit like SNPP, but must be taken into account in the design of low-Earth orbit scanning sensors similar to VIIRS to ensure contiguous coverage at nadir.

SNPP↗
NASA NTRS2017

Shallow Lunar Seismic Activity and the Current Stress State of the Moon

A vast, global network of more than 3200 lobate thrust fault scarps has been revealed in high resolution Lunar Reconnaissance Orbiter Camera (LROC) images. The fault scarps are very young, less than 50 Ma, based on their small scale and crisp appearance, crosscutting relations with small-diameter impact craters, and rates of infilling of associated small, shallow graben and may be actively forming today. The population of young thrust fault scarps provides a window into the recent stress state of the Moon and offers insight into the origin of global lunar stresses. The distribution of orientations of the fault scarps is non-random, inconsistent with isotropic stresses from late-stage global contraction as the sole source of stress. Modeling shows that tidal stresses contribute significantly to the current stress state of the lunar crust. Tidal stresses (orbital recession and diurnal tides) superimposed on stresses from global contraction result in non-isotropic compressional stress and may produce thrust faults consistent with lobate scarp orientations. At any particular point on the lunar surface, peak compressive stress will be reached at a certain time in the diurnal cycle. Coseismic slip events on currently active thrust faults are expected to be triggered when peak stresses are reached. Analysis of the timing of the 28 the shallow moonquakes recorded by the Apollo seismic network shows that 19 indeed occur when the Moon is closer to apogee, while only 9 shallow events occur when the Moon is closer to perigee. Here we report efforts to refine the model for the current stress state of the Moon by investigating the contribution of polar wander. Progress on relocating the epicentral locations of the shallow moonquakes using an algorithm designed for sparse networks is also reported.

structure↗
NASA NTRS2017

Spacecraft Charging Considerations and Design Efforts for the Orion Crew Module

The Orion Crew Module (CM) is nearing completion for the next flight, designated as Exploration Mission 1 (EM-1). For the uncrewed mission, the flight path will take the CM through a Perigee Raise Maneuver (PRM) out to an altitude of approximately 1800 km, followed by a Trans-Lunar Injection burn, a pass through the Van Allen belts then out to the moon for a lunar flyby, a Distant Retrograde Insertion (DRI) burn, a Distant Retrograde Orbit (DRO), a Distant Retrograde Departure (DRD) burn, a second lunar flyby, an Earth Insertion (EI) burn, and finally entry and landing. All of this, with the exception of the DRO associated maneuvers, is similar to the previous Apollo 8 mission in late 1968. In recent discussions, it is now possible that EM-1 will be a crewed mission, and if this happens, the orbit may be quite different from that just described. In this case, the flight path may take the CM on an out and back pass through the Van Allen belts twice, then out to the moon, again passing through the Van Allen belts twice, then finally back home. Even if the current EM-1 mission doesn't end up as a crewed mission, EM-2 and subsequent missions will undoubtedly follow orbital trajectories that offer comparable exposures to heightened vehicle charging effects. Because of this, and regardless of flight path, the CM vehicle will likely experience a wide range of exposures to energetic ions and electrons, essentially covering the gamut between low earth orbit to geosynchronous orbit and beyond. National Aeronautical and Space Administration (NASA) and Lockheed Martin (LM) engineers and scientists have been working to fully understand and characterize the vehicle's immunity level with regard to surface and deep dielectric charging, and the ramifications of that immunity level pertaining to materials and impacts to operational avionics, communications, and navigational systems. This presentation attempts to chronicle these efforts in a summary fashion, and attempts to capture the results of that work as they pertain to the electrical and avionic systems on-board the Orion CM.

Scully, Bob↗
NASA NTRS2017

Testing the Definition of the ESC Envelope

The previous effort, including a successful Change Control Request, addressed shrinking the size of the Earth Science Constellations' (ESC) Envelope by reducing the Margin. Fundamental to the purpose of the Envelope is the case where the argument of perigee of the secondary object circulates from 90 degrees to 270 degrees. This ("outside of the envelope, always outside the envelope") case was tested both numerically in a spreadsheet and analytically. Results showed how it is important to include the fact that a secondary with a different semi-major axis has a different frozen eccentricity value.

ESC envelope↗
NASA NTRS2017

Results of the Apogee-Raising Campaign of the Magnetospheric Multiscale Mission

This paper describes the apogee-raising campaign of the Magnetospheric Multiscale mission, where the spacecraft increased their apogee radii from 12 to 25 Earth radii in a total of 98 maneuvers. These maneuvers included an initial formation resize set to spread the spacecraft apart for safety, 32 apogee-raise delta-v maneuvers, their associated slews, four perigee-raise maneuvers and the associated slews, and finally a set of maneuvers to get back into formation. These activities were all accomplished successfully and on schedule with no anomalies, and at a fuel consumption somewhat less than predicted. As a result, MMS was set up ready to carry out in situ studies of magnetic reconnection in the magnetotail, with sufficient fuel remaining for a significant extended mission.

Mission Design↗
NASA NTRS2018

Arcus Mission Design: Stable Lunar Resonant HEO for X-ray Astronomy

The Arcus mission, proposed for NASA's 2016 Astrophysics Medium Explorer (MIDEX) announcement of opportunity, will use X-ray spectroscopy to detect previously unaccounted quantities of normal matter in the Universe. The Arcus mission design uses 4:1 lunar resonance to provide a stable orbit for visibility of widely-dispersed targets, in a low background radiation environment, above the Van Allen belts for the minimum two-year science mission. Additional ad-vantages of 4:1 resonance are long term stability without maintenance maneuvers, eclipses under 4.5 hours, perigee radius approximately 12 Re for data download, and streamlined operational cadence with approximately 1 week or-bit period.

Arcus↗
NASA NTRS2018

Arcus Mission Design: Stable Lunar-Resonant High Earth Orbit for X-Ray Astronomy

The Arcus mission, proposed for NASA's 2016 Astrophysics Medium Explorer (MIDEX) announcement of opportunity, will use X-ray spectroscopy to detect previously unaccounted quantities of normal matter in the Universe. The Arcus mission design uses 4:1 lunar resonance to provide a stable orbit for visibility of widely-dispersed targets, in a low background radiation environment, above the Van Allen belts for the minimum two-year science mission. Additional ad-vantages of 4:1 resonance are long term stability without maintenance maneu-vers, eclipses under 4.5 hours, perigee radius approximately 12 Re for data download, and streamlined operational cadence with approximately 1 week orbit period.

Arcus↗
NASA NTRS2016

In-Space Production of Storable Propellants

Our exploration of the feasibility of in-space production of storable propellants from resourcesavailable on asteroids, and also on Mars and the Moon, has considered the process of sampleacquisition; the energy requirements for heating and volatile release; the thermodynamic behaviorof gas release from minerals and organic polymers containing hydrogen, oxygen, carbon, sulfur,and nitrogen; purification of the released water and carbon dioxide; the storage and transportation of these materials in the condensed state; synthesis of fuels including methanol and dimethyl ether(DME); and the co-production, concentration, stability, and storage of the complementary oxidizer high-test hydrogen peroxide (HTP). We call attention to the ability of the storable propellant/oxidizer combination of DME and HTP to carry out deep-space missions and to perform retrieval and relocation of any and all space-derived resources, such as retrieving asteroidal metal to high Earth orbit. This study's analyses are based on returning resources to a storage/processing/dispensing facility in a Highly Eccentric Earth Orbit (HEEO) with a perigee above geosynchronous orbit and an apogee approach or beyond the Moon's orbit.The methane/LOX option, notable for good engine performance, has not been included in this study because our scope includes only fully storable propellants.

Lewis, John S.↗
NASA NTRS1959

An Analysis of the Corridor and Guidance Requirements for Supercircular Entry into Planetary Atmospheres

The analysis of super-circular entry is developed around a new dimensionless parameter which combines certain conditions at the conic perigee altitude with certain characteristics of the vehicle; this parameter conveniently determines either deceleration-limited or heating limited corridor widths for elliptic, parabolic, or hyperbolic approach trajectories. Illustrative calculations of corridor widths and the associated guidance problems are presented for Venus, Earth, Mars, Jupiter, and Titan. Generalized curves are presented for application to various entry conditions.

Chapman, Dean R.↗
NASA NTRSDate not supplied

Space Launch System Launch Window and Day of Launch Processes

Lunar missions benefit from varying the launch azimuth as a function of launch time to allow longer launch windows with minimum performance impacts. This variable azimuth approach allows the vehicle to track the Moon’s apparent motion due to Earth’s rotation . The Space Launch System (SLS) Block 1 vehicle design requires the mission to launch into an elliptical parking orbit to provide sufficient energy to insert Orion into a trans-lunar trajectory. The primary benefit of varying the launch azimuth, and as a result the parking orbit inclination, allows the SLS Interim Cryogenic Propulsion Stage (ICPS) to perform its Trans-Lunar Injection (TLI) burn closer to perigee and take advantage of performing a burn in a location where the burn will primarily raise apogee.

A. S. Craig↗
NASA NTRSDate not supplied

Space Launch System Launch Windows and Day of Launch Processes

Lunar missions benefit from varying the launch azimuth as a function of launch time to allow longer launch windows with minimum performance impacts. This variable azimuth approach allows the vehicle to track the Moon’s apparent motion due to Earth’s rotation. The Space Launch System (SLS) Block 1 vehicle design requires the mission to launch into an elliptical parking orbit to provide sufficient energy to insert Orion into a Trans-Lunar Injection (TLI) orbit. The primary benefit of varying the launch azimuth, and as a result the achieved orbit inclination, allows the SLS Interim Cryogenic Propulsion Stage (ICPS) to perform its TLI burn closer to perigee and take advantage of performing a burn in a location where the burn will optimally raise apogee.

launch window↗
NASA NTRS2018

In-Flight Operation of the Dawn Ion Propulsion System Through Completion of the Final Orbit Around Dwarf Planet Ceres

The Dawn mission, part of NASA’s Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, transfer between Ceres science orbits, and orbit maintenance maneuvers. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional DV of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in August 2012 and completed in late December 2014. From December 2014 through June 2016 the IPS was used for transiting the spacecraft to all science orbits at Ceres including the final orbit for Dawn’s primary mission, called the low altitude mapping orbit (LAMO), a circular orbit at a mean altitude above Ceres of approximately 385 km. Dawn met or exceeded all pre-launch science requirements and Dawn's prime mission concluded on June 30, 2016. Dawn subsequently received NASA approval for two extended missions at Ceres, called XM1 and XM2. During XM1 IPS was used for orbit maintenance and to transit the spacecraft to several new, higher-altitude science orbits, ending in an elliptical orbit at a maximum altitude of approximately 38,000 km. In XM2, IPS operations included transiting the Dawn spacecraft to its final, highly elliptical orbit at Ceres ranging from 35 km perigee to 4,800 km apogee. Science data acquisition will continue in XM2 until the hydrazine is exhausted, which is expected to occur between August and October 2018. Dawn has successfully completed all science goals for both the primary and XM1 extended missions. To date the IPS has been operated for approximately 51,250 hours, consumed approximately 416 kg of xenon, and provided a delta-V of almost 11.5 km/s, a record for an on-board propulsion system. The IPS performance characteristics are close to the expected performance based on analysis and testing performed pre-launch. Dawn’s IPS continues to be fully operational as of June 2018. This paper provides an overview of Dawn’s mission objectives and the results of Dawn IPS mission operations for XM1, and XM2 through June 2018.

Rayman, Marc D.↗
NASA NTRSDate not supplied

Crucial Role of Thermal Gradients in MMS Fluxgate In-Flight Calibration

To meet the science goals of the Magnetospheric Multiscale (MMS) mission, the Fluxgate Magnetometer (FGM) must measure the ambient magnetic field with an accuracy of 0.1 nT. On a typical MMS orbit, the offsets (or zero levels) of the 3-axis FGM can vary by ~0.5 nT (exclusive of periods in Earth shadow). Previous studies have shown that these variations can be characterized as functions of sensor temperature, TS, and can thus be corrected to within 0.2 nT in the spin plane using in-flight calibration techniques (Bromund, et al., 2016, https://ntrs.nasa.gov/citations/20160014711). In that presentation, we noted two significant observations: A distinct function of TS must be used to characterize offsets during shadow: offsets at a given TScan differ by as much as 2 nT in shadow vs sunlight. Offsets change after maneuvers, without a commensurate change in TS . These changes can be as large as 2 nT. We now note a third, related observation: Offsets increase with proximity to the earth even when TS is constant, resulting in variations of ~0.2 nT at 4-5 Earth radii (RE) These effects are evidence that offsets are a multivariate function of TS and another factor, namely: thermal gradients. Due to the spacecraft spin, the Earth and the Sun each provide a relatively constant thermal input onto one instrument face while the opposite face remains in shadow, thus giving rise to thermal gradients. The thermal gradient depend on the orientation of the spin axis relative to the Earth or Sun. We observe that offsets vary by as much as 0.17 nT/degree as a function of the tilt of the spin axis towards the Sun. Thermal input from the Earth is dominated by Outgoing Longwave Radiation (OLR). Due to the proximity to Earth, the inverse proportion of the distance squared is a significant factor in the thermal gradient attributed to Earth OLR. We find that offsets can be corrected to <0.05 nT accuracy near perigee when accounting for these factors using empirically determined constants of proportionality that account for differences in emissivity of the top and bottom faces of the sensor to OLR (as well as other thermal effects). The changes in offset associated with thermal gradient are of the same order of magnitude as effects that were formerly attributed to sensor temperature alone, and thus both parameters are necessary to characterize the FGM offsets.

Kenneth R Bromund↗