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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.

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NASA NTRS2010

Autonomous GN and C for Spacecraft Exploration of Comets and Asteroids

A spacecraft guidance, navigation, and control (GN&C) system is needed to enable a spacecraft to descend to a surface, take a sample using a touch-and-go (TAG) sampling approach, and then safely ascend. At the time of this reporting, a flyable GN&C system that can accomplish these goals is beyond state of the art. This article describes AutoGNC, which is a GN&C system capable of addressing these goals, which has recently been developed and demonstrated to a maturity TRL-5-plus. The AutoGNC solution matures and integrates two previously existing JPL capabilities into a single unified GN&C system. The two capabilities are AutoNAV and GREX. AutoNAV is JPL s current flight navigation system, and is fairly mature with respect to flybys and rendezvous with small bodies, but is lacking capability for close surface proximity operations, sampling, and contact. G-REX is a suite of low-TRL algorithms and capabilities that enables spacecraft operations in close surface proximity and for performing sampling/contact. The development and integration of AutoNAV and G-REX components into AutoGNC provides a single, unified GN&C capability for addressing the autonomy, close-proximity, and sampling/contact aspects of small-body sample return missions. AutoGNC is an integrated capability comprising elements that were developed separately. The main algorithms and component capabilities that have been matured and integrated are autonomy for near-surface operations, terrain-relative navigation (TRN), real-time image-based feedback guidance and control, and six degrees of freedom (6DOF) control of the TAG sampling event. Autonomy is achieved based on an AutoGNC Executive written in Virtual Machine Language (VML) incorporating high-level control, data management, and fault protection. In descending to the surface, the AutoGNC system uses camera images to determine its position and velocity relative to the terrain. This capability for TRN leverages native capabilities of the original AutoNAV system, but required advancements that integrate the separate capabilities for shape modeling, state estimation, image rendering, defining a database of onboard maps, and performing real-time landmark recognition against the stored maps. The ability to use images to guide the spacecraft requires the capability for image-based feedback control. In Auto- GNC, navigation estimates are fed into an onboard guidance and control system that keeps the spacecraft guided along a desired path, as it descends towards its targeted landing or sampling site. Once near the site, AutoGNC achieves a prescribed guidance condition for TAG sampling (position/orientation, velocity), and a prescribed force profile on the sampling end-effector. A dedicated 6DOF TAG control then implements the ascent burn while recovering from sampling disturbances and induced attitude rates. The control also minimizes structural interactions with flexible solar panels and disallows any part of the spacecraft from making contact with the ground (other than the intended end-effector).

Carson, John M.↗
NASA NTRS1986

Galileo inflight friction identification for the scan platform control loop

Two types of friction identification algorithms (position-based and rate-based) are investigated in an attempt to better understand cone actuator inflight behavior in the Galileo scan platform pointing system. This investigation involves the measurement of the scan platform position (or rate) with respect to the stator following a platform slew maneuver. Inflight data collection by the spacecraft's sensors and subsequent ground-based data processing comprise the two major steps in the identification experiment. The proposed algorithms were capable of obtaining friction estimates within + or - 10 percent accuracy; hence, it is concluded that this inflight system identification technique can sufficiently tune the Galileo cone controller.

Ianculescu, G. D.↗
NASA NTRS1990

Sensitivity of high-accuracy Tracking and Data Relay Satellite System (TDRSS) user spacecraft orbit determination to tracking schedules

The results of a study to analyze the dependence of TDRSS user spacecraft orbit determination consistencies on varying tracking schedules are presented. In this study, the TDRS-East orbit determination results obtained utilizing Bilateration Ranging Transponder System data were evaluated. Six state parameters, three position and three velocity components and the solar radiation pressure coefficient, are estimated for TDRS-East. It is concluded that, in order to achieve high-precision orbit determination, the tracking coverage should not fall below 10 minutes every two orbits as decreasing it to every four orbits will significantly degrade the accuracy; present state-of-the-art consistency in orbit determination using TDRSS tracking is approximately 15 to 20 meters.

Doll, C. E.↗
NASA NTRS2017

Orbit Determination of the Lunar Reconnaissance Orbiter: Status After Seven Years

The Lunar Reconnaissance Orbiter (LRO) has been orbiting the Moon since 2009, obtaining unique and foundational datasets important to understanding the evolution of the Moon and the Solar System. The high-resolution data acquired by LRO benefit from precise orbit determination (OD), limiting the need for geolocation and co-registration tasks. The initial position knowledge requirement (50 m) was met with radio tracking from ground stations, after combination with LOLA altimetric crossovers. LRO-specific gravity field solutions were determined and allowed radio-only OD to perform at the level of 20 m, although secular inclination changes required frequent updates. The high-accuracy gravity fields from GRAIL, with <10 km spatial resolution, further improved the radio-only orbit reconstruction quality (<10 m). However, orbit reconstruction is in part limited by the 0.3-0.5 mm/s measurement noise level in S-band tracking. 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 lunar hemisphere facing the Earth due to tidal locking). Here, we report on the status of the OD effort since the beginning of the mission, a period spanning more than seven years. We describe modeling improvements and the use of new measurements. In particular, the LOLA altimetric data give accurate, uniform, and independent information about LRO's orbit, with a different sensitivity and geometry which includes coverage over the lunar farside and is not tied to ground-based assets. With SLDEM2015 (a combination of the LOLA topographic profiles and the Kaguya Terrain Camera stereo images), another use of altimetry is possible for OD. We extend the 'direct altimetry' technique developed for the ICESat mission to perform OD and adjust spacecraft position to minimize discrepancies between LOLA tracks and SLDEM2015. Comparisons with the radio-only orbits are used to evaluate this new tracking type, of interest for the OD of future lunar orbiters carrying a laser altimeter. LROC NAC images also provide independent accuracy estimation, through the repeated views taken of anthropogenic features for instance.

Erwan Mazarico↗
NASA NTRS2016

Constraints on LISA Pathfinder's Self-Gravity: Design Requirements, Estimates and Testing Procedures

LISA Pathfinder satellite was launched on 3 December 2015 toward the Sun Earth first Lagrangian point (L1) where the LISA Technology Package (LTP), which is the main science payload, will be tested. LTP achieves measurements of differential acceleration of free-falling test masses (TMs) with sensitivity below 3 x 10(exp -14) m s(exp -2) Hz(exp - 1/2) within the 130 mHz frequency band in one dimension. The spacecraft itself is responsible for the dominant differential gravitational field acting on the two TMs. Such a force interaction could contribute a significant amount of noise and thus threaten the achievement of the targeted free-fall level. We prevented this by balancing the gravitational forces to the sub nm s(exp -2) level, guided by a protocol based on measurements of the position and the mass of all parts that constitute the satellite, via finite element calculation tool estimates. In this paper, we will introduce the gravitational balance requirements and design, and then discuss our predictions for the balance that will be achieved in flight.

Armano, M.↗
NASA NTRS1977

Spacecraft-produced neutron fluxes on Skylab

Estimates of neutron fluxes in different energy ranges are reported for the Skylab spacecraft. Detectors composed of uranium, thorium, and bismuth foils with mica as a fission track recorder, as well as boron foils with cellulose acetate as an alpha-particle recorder, were deployed at different positions in the Orbital Workshop. It was found that the Skylab neutron flux was dominated by high energy (greater than 1 MeV) contributions and that there was no significant time variation in the fluxes. Firm upper limits of 7-15 neutrons/sq cm-sec, depending on the detector location in the spacecraft, were established for fluxes above 1 MeV. Below 1 MeV, the neutron fluxes were about an order of magnitude lower. The neutrons are interpreted as originating from the interactions of leakage protons from the radiation belt with the spacecraft.

Quist, T. C.↗
NASA NTRS2011

Attitude Estimation in Fractionated Spacecraft Cluster Systems

An attitude estimation was examined in fractioned free-flying spacecraft. Instead of a single, monolithic spacecraft, a fractionated free-flying spacecraft uses multiple spacecraft modules. These modules are connected only through wireless communication links and, potentially, wireless power links. The key advantage of this concept is the ability to respond to uncertainty. For example, if a single spacecraft module in the cluster fails, a new one can be launched at a lower cost and risk than would be incurred with onorbit servicing or replacement of the monolithic spacecraft. In order to create such a system, however, it is essential to know what the navigation capabilities of the fractionated system are as a function of the capabilities of the individual modules, and to have an algorithm that can perform estimation of the attitudes and relative positions of the modules with fractionated sensing capabilities. Looking specifically at fractionated attitude estimation with startrackers and optical relative attitude sensors, a set of mathematical tools has been developed that specify the set of sensors necessary to ensure that the attitude of the entire cluster ( cluster attitude ) can be observed. Also developed was a navigation filter that can estimate the cluster attitude if these conditions are satisfied. Each module in the cluster may have either a startracker, a relative attitude sensor, or both. An extended Kalman filter can be used to estimate the attitude of all modules. A range of estimation performances can be achieved depending on the sensors used and the topology of the sensing network.

Hadaegh, Fred Y.↗
NASA NTRS2004

Improved Estimate of Phobos Secular Acceleration from MOLA Observations

We report on new observations of the orbital position of Phobos, and use them to obtain a new and improved estimate of the rate of secular acceleration in longitude due to tidal dissipation within Mars. Phobos is the inner-most natural satellite of Mars, and one of the few natural satellites in the solar system with orbital period shorter than the rotation period of its primary. As a result, any departure from a perfect elastic response by Mars in the tides raised on it by Phobos will cause a transfer of angular momentum from the orbit of Phobos to the spin of Mars. Since its discovery in 1877, Phobos has completed over 145,500 orbits, and has one of the best studied orbits in the solar system, with over 6000 earth-based astrometric observations, and over 300 spacecraft observations. As early as 1945, Sharpless noted that there is a secular acceleration in mean longitude, with rate (1.88 + 0.25) 10(exp -3) degrees per square year. In preparation for the 1989 Russian spacecraft mission to Phobos, considerable work was done compiling past observations, and refining the orbital model. All of the published estimates from that era are in good agreement. A typical solution (Jacobson et al., 1989) yields (1.249 + 0.018) 10(exp -3) degrees per square year. The MOLA instrument on MGS is a laser altimeter, and was designed to measure the topography of Mars. However, it has also been used to make observations of the position of Phobos. In 1998, a direct range measurement was made, which indicated that Phobos was slightly ahead of the predicted position. The MOLA detector views the surface of Mars in a narrow field of view, at 1064 nanometer wavelength, and can detect shadows cast by Phobos on the surface of Mars. We have found 15 such serendipitous shadow transit events over the interval from xx to xx, and all of them show Phobos to be ahead of schedule, and getting progressively farther ahead of the predicted position. In contrast, the cross-track positions are quite close to the predicted values. Assuming that the along-track discrepancy is small enough that we can linearize the corrections, we model the mean orbital longitude as a quadratic function of time, and solve for corrections to the constant, linear, and quadratic terms. The time span of the recent observations is insufficient to properly resolve this issue alone, but when the 127 years of prior observations are added, we find a solution which reduces misfit to the new observations considerably, and makes no significant change to the fit to earlier observations. Our estimate for the secular acceleration term is (1.367 + 0.006) degrees per square year. The corresponding rate of energy dissipation is 3.34 MW. From a geophysical perspective, a more interesting parameter than the secular acceleration itself is the tidal lag angle, or tidal quality factor Q, for Mars. Unfortunately, the limiting error source in that determination is remaining uncertainty in the tidal Love numbers at harmonic degrees 2,3, and even 4. Until those parameters are better constrained, improvement in the orbital model of Phobos will not provide corresponding benefits for understanding the interior of Mars.

Bills, Bruce↗
NASA NTRS1973

En route position and time control of aircraft using Kalman filtering of radio aid data

Fixed-time-of-arrival (FTA) guidance and navigation is investigated as a possible technique capable of operation within much more stringent en route separation standards and offering significant advantages in safety, higher traffic densities, and improved scheduling reliability, both en route and in the terminal areas. This study investigated the application of FTA guidance previously used in spacecraft guidance. These FTA guidance techniques have been modified and are employed to compute the velocity corrections necessary to return an aircraft to a specified great-circle reference path in order to exercise en route time and position control throughout the entire flight. The necessary position and velocity estimates to accomplish this task are provided by Kalman filtering of data from Loran-C, VORTAC/TACAN, Doppler radar, radio or barometric altitude,and altitude rate. The guidance and navigation system was evaluated using a digital simulation of the cruise phase of supersonic and subsonic flights between San Francisco and New York City, and between New York City and London.

Mcgee, L. A.↗
NASA NTRS1994

Calibration of the TOPEX altimeter using a GPS buoy

The use of a spar buoy equipped with a Global Positioning System (GPS) antenna to calibrate the height measurement of the TOPEX radar altimeter is described. In order to determine the height of the GPS antenna phase center above the ocean surface, the buoy was also equipped with instrumentation to measure the instantaneous location of the waterline, and tilt of the bouy from vertical. The experiment was conducted off the California coast near the Texaco offshore oil platform, Harvest, during cycle 34 of the TOPEX/POSEIDON observational period. GPS solutions were computed for the bouy position using two different software packages, K&RS and GIPSY-OASIS II. These solutions were combined with estimates of the waterline location on the bouy to yield the height of the ocean surface. The ocean surface height in an absolute coordinate system combined with knowledge of the spacecraft height from tracking data provides a computed altimeter range measurement. By comparing this computed value to the actual altimeter measurement, the altimeter bias can be calibrated. The altimeter height bias obtained with the buoy using K&RS was -14.6 +/- 4 cm, while with GIPSY-OASIS II it was -13.1 +/- 4 cm. These are 0.1 cm and 1.6 cm different from the -14.7 +/- 4 cm result obtained for this flight overflight with the tide gauge instruments located on Platform Harvest.

Born, G. H.↗
NASA NTRS2007

On Target to Mars

This viewgraph presentation reviews the use of Descent Image Motion Estimation System (DIMES) for the descent of a spacecraft onto the surface of Mars. In the past this system was used to assist in the landing of the MER spacecraft. The overall algorithm is reviewed, and views of the hardware, and views from Spirit's descent are shown. On Spirit, had DIMES not been used, the impact velocity would have been at the limit of the airbag capability and Spirit may have bounced into Endurance Crater. By using DIMES, the velocity was reduced to well within the bounds of the airbag performance and Spirit arrived safely at Mars. Views from Oppurtunity's descent are also shown. The system to avoid and detect hazards is reviewed next. Landmark Based Spacecraft Pinpoint Landing is also reviewed. A cartoon version of a pinpoint landing and the various points is shown. Mars s surface has a large amount of craters, which are ideal landmarks . According to literatures on Martian cratering, 60 % of Martian surface is heavily cratered. The ideal (craters) landmarks for pinpoint landing will be between 1000 to 50 meters in diagonal The ideal altitude for position estimation should greater than 2 km above the ground. The algorithms used to detect and match craters are reviewed.

Mars↗
NASA NTRSDate not supplied

Solar Sail Torque Model Characterization for the Near Earth Asteroid Scout Mission

Near Earth Asteroid Scout (NEA Scout) was a mission to test solar sail propulsion for orbital transfer from cislunar space to flyby and image an asteroid. Had it succeeded, one of the mission goals was to characterize the solar torque on the sail to ensure successful attitude control for the orbit transfer and imaging the asteroid. The simulation used to develop the flight attitude control software uses the generalized model for solar sails, a tensor equation of the forces and torques on sails of arbitrary shape. Rios-Reyes and Scheeres developed a general process to update the torque tensor coefficients using estimates of sail torque over a range of directions to the sun. Their process was adapted and implemented for the specific case of NEA Scout using spacecraft telemetry collected during sail characterization maneuvers in combination with simulation models and parameters. The NEA Scout maneuvers were limited to the operating range of the mission and constraints of the control hardware and allowed safe testing of each attitude before proceeding to the next. The NEA Scout reaction wheel speeds are used to measure accumulated momentum, while the Active Mass Translator (AMT) position is used to subtract out the torque from the center of mass crossed with the sail force and isolate the torque from only the sail shape. The process was tested by running attitude control simulations of the characterization maneuvers, generating simulated telemetry, estimating the solar torques, then using a least squares estimating the solar torque coefficients using least-squares and then performing a least-squares fit to the solar torque tensor coefficients. These estimated coefficients were tested by evaluating the solar torques under the same conditions as the simulated telemetry and comparing to the true simulated torques. Solar force model updates can be performed separately by observing the effect of the sail on the trajectory, and the torque model can be refined using those solar force updates. This process met the needs of the NEA Scout mission and can be adapted to characterize the solar torque for other missions with different sails.

solar sail↗
NASA NTRSDate not supplied

Solar Sail Torque Model Characterization for the Near Earth Asteroid Scout Mission

Near Earth Asteroid Scout (NEA Scout) was a mission to test solar sail propulsion for orbital transfer from cislunar space to flyby and image an asteroid. One of the goals of the mission was to characterize the solar torque on the sail to ensure successful attitude control for the orbit transfer and imaging the asteroid. The simulation used to develop the flight attitude control software uses the generalized model for solar sails, a tensor equation of the forces and torques on sails of arbitrary shape. Rios-Reyes and Scheeres developed a general process to update the torque tensor coefficients using estimates of sail torque over a range of directions to the sun. Their process was adapted and implemented for the specific case of NEA Scout using spacecraft telemetry collected during sail characterization maneuvers in combination with simulation models and parameters. The NEA Scout maneuvers were limited to the operating range of the mission and constraints of the control hardware and allowed safe testing of each attitude before proceeding to the next. The NEA Scout reaction wheel speeds are used to measure accumulated momentum, while the Active Mass Translator (AMT) position is used to subtract out the torque from the center of mass crossed with the sail force and isolate the torque from only the sail shape. The process was tested by running attitude control simulations of the characterization maneuvers, generating simulated telemetry, estimating the solar torques, then using a least squares estimating the solar torque coefficients using least-squares and then performing a least-squares fit to the solar torque tensor coefficients. These estimated coefficients were tested by evaluating the solar torques under the same conditions as the simulated telemetry and comparing to the true simulated torques. Solar force model updates can be performed separately by observing the effect of the sail on the trajectory, and the torque model can be refined using those solar force updates. This process met the needs of the NEA Scout mission and can be adapted to characterize the solar torque for other missions with different sails.

solar sail↗
NASA NTRS1972

Nonlinear and adaptive estimation in reentry.

The problem of real-time estimation of a lifting reentry vehicle trajectory of the shuttle orbiter type is considered. Simulations feature large position and velocity uncertainties at radar acquisition and realistic model errors in lift, drag and other model parameters. Radar tracking and accelerometer data are simulated. Significant nonlinearities are found to exist on spacecraft acquisition. An iterated nonlinear filter is shown to perform optimally during the radar acquisition phase. An adaptive filter is shown to track time-varying model errors, such as errors in the lift and drag coefficients, down to the noise level. Such real-time model tracking (identification) is frequently required for guidance and control implementation.

Jazwinski, A. H.↗
NASA NTRS2002

The StarLight interferometer architecture and operational concepts

The StarLight flight project was designed to demonstrate the key technologies of spacebome long-baseline stellar interferometry and precision formation flying for potential use on the Terrestrial Planet Finder (TPF) and other future astrophysics missions. Interferometer performance validation could be achieved over a 6-12 month period by obtaining several hundred fringe visibility amplitude measurements for stars in the band 600-1000 nm for a variety of stellar visibilities (0.2-1.0), magnitudes (Mv = 2-5), and baselines (B = 30-125 meters). Interferometry could be performed both in a 1 meter fixed-baseline combiner-only mode and in a two-spacecraft formation mode. In formation mode, the combiner spacecraft would remain at the focus of a virtual parabola, while the collector spacecraft assumed various positions along the parabola such that the two arms of the interferometer remained equal over a variety of separations and bearing angles. Challenges to be encountered in flight include high-bandwidth inter-spacecraft stellar and metrology pointing control, alignment and shear correction, delay and delay-rate estimation, visibility calibration, and robust fringe tracking in the presence of local and inter-spacecraft dynamics. This paper is based on the StarLight project design-capture of March 2002 and will describe the StarLight Interferometer System architecture and selected operational concepts (both of which have relevance to the on-going TPF Technology Program).

stellar↗
NASA NTRS1996

GPS Attitude Determination Using Deployable-Mounted Antennas

The primary objective of this investigation is to develop a method to solve for spacecraft attitude in the presence of potential incomplete antenna deployment. Most research on the use of the Global Positioning System (GPS) in attitude determination has assumed that the antenna baselines are known to less than 5 centimeters, or one quarter of the GPS signal wavelength. However, if the GPS antennas are mounted on a deployable fixture such as a solar panel, the actual antenna positions will not necessarily be within 5 cm of nominal. Incomplete antenna deployment could cause the baselines to be grossly in error, perhaps by as much as a meter. Overcoming this large uncertainty in order to accurately determine attitude is the focus of this study. To this end, a two-step solution method is proposed. The first step uses a least-squares estimate of the baselines to geometrically calculate the deployment angle errors of the solar panels. For the spacecraft under investigation, the first step determines the baselines to 3-4 cm with 4-8 minutes of data. A Kalman filter is then used to complete the attitude determination process, resulting in typical attitude errors of 0.50.

Osborne, Michael L.↗
NASA NTRS1982

Visible light observations of a dense plasmoid associated with a moving Type IV solar radio burst

A coronal transient rising above the eastern limb of the sun was observed simultaneously on April 27, 1980 by the Coronagraph/Polarimeter (C/P) aboard the Solar Maximum Mission spacecraft and by the Culgoora radioheliograph (CRH). The C/P observed an outward-moving loop transient followed by a plasmoid of dense ionized material, while the CRH observed several types of radio bursts, including a moving Type IV burst at 80 MHz whose position coincided with the bright plasmoid. An estimate of the reasonable lower limit of the electron density in the moving Type IV burst indicates that this burst could have resulted from emission at either the fundamental or the second harmonic of the plasma frequency. This is a new result because in the past it was thought that the density in the moving Type IV source region would be too low for plasma radiation.

Stewart, R. T.↗
NASA NTRS2014

Free-Flight Experiments in LISA Pathfinder

The LISA Pathfinder mission will demonstrate the technology of drag-free test masses for use as inertial references in future space-based gravitational wave detectors. To accomplish this, the Pathfinder spacecraft will perform drag-free flight about a test mass while measuring the acceleration of this primary test mass relative to a second reference test mass. Because the reference test mass is contained within the same spacecraft, it is necessary to apply forces on it to maintain its position and attitude relative to the spacecraft. These forces are a potential source of acceleration noise in the LISA Pathfinder system that are not present in the full LISA configuration. While LISA Pathfinder has been designed to meet it's primary mission requirements in the presence of this noise, recent estimates suggest that the on-orbit performance may be limited by this 'suspension noise'. The drift-mode or free-flight experiments provide an opportunity to mitigate this noise source and further characterize the underlying disturbances that are of interest to the designers of LISA-like instruments. This article provides a high-level overview of these experiments and the methods under development to analyze the resulting data.

The technique of drag-free↗