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We calculate localized gravity/topography admittances for Mars, in order to estimate elastic lithosphere thickness. A finite-amplitude correction to modeled gravity is required to properly interpret admittances in high-relief regions of Mars. Additional information is contained in the original extended abstract.
The GEOSAT Follow-On spacecraft (GFO-l), launched in 1998, began continuous radar altimeter coverage of the oceans in 2000, and after an extensive series of calibration campaigns in 1999 and 2000 was finally accepted by the US Navy on November 29, 2000. By providing high quality altimeter data, GFO can supplement TOPEX/POSEIDON(T/P), ERS-1, and ERS-2 data, providing a different synoptic sampling of the oceans with its 17-day ground track repeat cycle. Altimeter analysis suggests GFO is capable of POSEIDON class altimetry, with orbit error the largest contributor to the GFO altimeter error budget. Satellite laser ranging (SLR), especially in combination with altimeter crossover data, offer the only means of determining high-quality precise orbits. SLR tracking is also augmented by Doppler (Tranet style) beacons. These data were used to tune the gravity field and macromodel (3D representation of the spacecraft geometry and surface properties). Beginning with January 2000 GSFC has produced a 2.5 year span of the Precision Orbit Ephemeris (POE) intended for use on the GFO GDR and for the Pathfinder Project. Using GFO GDR and IGDR altimeter data from NOAA, a two year span of POE orbits are evaluated with respect to the TOPEX orbits by adjusting the GFO data into the TOPEX frame with the help of TOPEX/GFO altimeter crossovers. A 2.5 year span of orbits are also evaluated using tracking data residual analysis, GFO altimeter crossover residual analysis, and direct orbit comparison. In an effort to improve the current POE orbit several ideas are considered, including using recent gravity models which include CHAMP data, the reduced-dynamic approach, and including TOPEX/GFO altimeter crossover data in the orbit solution.
A final release (Version L) of the Mars Orbiter Laser Altimeter (MOLA) Precision Experiment Data Record (PEDR) has been submitted to the Planetary Data System (PDS). Additional gridded data record products are forthcoming. These products have evolved since their original description, owing in part to improved gravity modeling and cartographic reference frames, and in part to refinements in calibration. An additional component, the 1064 nm narrowband radiometry data, is also being archived. These data will be invaluable for future studies by Mars explorers and scientists.
The oblateness of the Earth's gravity field, 52, has long been observed to undergo a slight decrease due to post-glacial rebound of the mantle. Sometime around 1998 this trend reversed quite suddenly. This reversal persisted until 2001, at which point the atmosphere-corrected time series appears to have reversed yet again. Presently, the time series appears to be returning to the value that would nominally have been reached had the anomaly not occurred. This anomaly signifies a large interannual change in global mass distribution whose J2 effect overshadows that of the post-glacial rebound over such timescales. A number of possible causes have been considered, with oceanic mass redistribution as the leading candidate although other effects, such as glacial melting and core effects may be contributing. The amount by which J2 returns to it's nominal value provides a valuable constraint on the separation of the causes, and will be considered. We will present our latest Satellite Laser Ranging and DORIS Doppler derived time series for J2, and various other low-degree harmonic terms, as well as our investigations into the causes. In addition, we will show the comparison of the J2 results with those derived from CHAMP, as computed at NASA GSFC, and the recently released GRACE gravity model.
We have compared GPS-based dynamic and reduced-dynamic TOPEX/POSEIDON orbits over a 10 day period beginning March 10,1993. The results suggest that the prelaunch gravity model (JGM-1) introduces geographically correlated errors which, when projectd on a mean sea survase, have a storng meridional dependence.
The US Navy's GEOSAT Follow-On spacecraft (GFO) primary mission objective is to map the oceans using a radar altimeter. Satellite laser ranging data, especially in combination with altimeter crossover data, offer the only means of determining high-quality precise orbits. Two tuned gravity models, PGS7727 and PGS7777b, were created at NASA GSFC for GFO that reduce the predicted radial orbit through degree 70 to 13.7 and 10.0 mm. A macromodel was developed to model the nonconservative forces and the SLR spacecraft measurement offset was adjusted to remove a mean bias. Using these improved models, satellite-ranging data, altimeter crossover data, and Doppler data are used to compute both daily medium precision orbits with a latency of less than 24 hours. Final precise orbits are also computed using these tracking data and exported with a latency of three to four weeks to NOAA for use on the GFO Geophysical Data Records (GDR s). The estimated orbit precision of the daily orbits is between 10 and 20 cm, whereas the precise orbits have a precision of 5 cm.
The Lunar Reconnaissance Orbiter (LRO) will be the first mission under NASA's Vision for Space Exploration. LRO will fly in a low 50 km mean altitude lunar polar orbit. LRO will utilize a direct minimum energy lunar transfer and have a launch window of three days every two weeks. The launch window is defined by lunar orbit beta angle at times of extreme lighting conditions. This paper will define the LRO launch window and the science and engineering constraints that drive it. After lunar orbit insertion, LRO will be placed into a commissioning orbit for up to 60 days. This commissioning orbit will be a low altitude quasi-frozen orbit that minimizes stationkeeping costs during commissioning phase. LRO will use a repeating stationkeeping cycle with a pair of maneuvers every lunar sidereal period. The stationkeeping algorithm will bound LRO altitude, maintain ground station contact during maneuvers, and equally distribute periselene between northern and southern hemispheres. Orbit determination for LRO will be at the 50 m level with updated lunar gravity models. This paper will address the quasi-frozen orbit design, stationkeeping algorithms and low lunar orbit determination.
This slide presentation reviews the planned sequence of events and the rationale for all lunar missions, and the flight experiences and lessons learned for the lunar orbit activities from a trajectory perspective. Shown are trajectories which include the moon's position at the various stages in the complete trip from launch, to the return and reentry. Included in the presentation are objectives and the sequence of events,for the Apollo 8, and Apollo 10. This is followed by a discussion of Apollo 11, including: the primary mission objective, the sequence of events, and the flight experience. The next mission discussed was Apollo 12. It reviews the objectives, the ground tracking, procedure changes, and the sequence of events. The aborted Apollo 13 mission is reviewed, including the objectives, and the sequence of events. Brief summaries of the flight experiences for Apollo 14-16 are reviewed. The flight sequence of events of Apollo 17 are discussed. In summary each mission consistently performing precision landings required that Apollo lunar orbit activities devote considerable attention to: (1) Improving fidelity of lunar gravity models, (2) Maximizing availability of ground tracking, (3) Minimizing perturbations on the trajectory, (4) Maximizing LM propellant reserves for hover time. Also the use of radial separation maneuvers (1) allows passive re-rendezvous after each rev, but ... (2) sensitive to small dispersions in initial sep direction
Ultrahigh energy cosmic rays that produce giant extensive showers of charged particles and photons when they interact in the Earth's atmosphere provide a unique tool to search for new physics. Of particular interest is the possibility of detecting a very small violation of Lorentz invariance such as may be related to the structure of space-time near the Planck scale of approximately 10 (exp -35) m. We discuss here the possible signature of Lorentz invariance violation on the spectrum of ultrahigh energy cosmic rays as compared with present observations of giant air showers. We also discuss the possibilities of using more sensitive detection techniques to improve searches for Lorentz invariance violation in the future. Using the latest data from the Pierre Auger Observatory, we derive a best fit to the LIV parameter of 3 .0 + 1.5 - 3:0 x 10 (exp -23) ,corresponding to an upper limit of 4.5 x 10-23 at a proton Lorentz factor of approximately 2 x 10(exp 11) . This result has fundamental implications for quantum gravity models.
In order to support the anytime abort requirements of a manned mission to the Moon, the vehicle abort capabilities for the translunar and circumlunar phases of the mission must be studied. Depending on the location of the abort maneuver, the maximum return time to Earth and the available propellant, two different kinds of return trajectories can be calculated: direct and fly-by. This paper presents a new method to compute these return trajectories in a deterministic and fast way without using numerical optimizers. Since no simplifications of the gravity model are required, the resulting trajectories are very accurate and can be used for both mission design and operations. This technique has been extensively used to evaluate the abort capabilities of the Orion/Altair vehicles in the Constellation program for the translunar phase of the mission.
Aerocapture is an orbit insertion maneuver in which a spacecraft flies through a planetary atmosphere one time using drag force to decelerate and effect a hyperbolic to elliptical orbit change. Aerocapture employs a feedback Guidance, Navigation, and Control (GN&C) system to deliver the spacecraft into a precise postatmospheric orbit despite the uncertainties inherent in planetary atmosphere knowledge, entry targeting and aerodynamic predictions. Only small amounts of propellant are required for attitude control and orbit adjustments, thereby providing mass savings of hundreds to thousands of kilograms over conventional all-propulsive techniques. The Analytic Predictor Corrector (APC) guidance algorithm has been developed to steer the vehicle through the aerocapture maneuver using bank angle control. Through funding provided by NASA's In-Space Propulsion Technology Program, the operation of an aerocapture GN&C system has been demonstrated in high-fidelity simulations that include real-time hardware in the loop, thus increasing the Technology Readiness Level (TRL) of aerocapture GN&C. First, a non-real-time (NRT), 6-DOF trajectory simulation was developed for the aerocapture trajectory. The simulation included vehicle dynamics, gravity model, atmosphere model, aerodynamics model, inertial measurement unit (IMU) model, attitude control thruster torque models, and GN&C algorithms (including the APC aerocapture guidance). The simulation used the vehicle and mission parameters from the ST-9 mission. A 2000 case Monte Carlo simulation was performed and results show an aerocapture success rate of greater than 99.7%, greater than 95% of total delta-V required for orbit insertion is provided by aerodynamic drag, and post-aerocapture orbit plane wedge angle error is less than 0.5 deg (3-sigma). Then a real-time (RT), 6-DOF simulation for the aerocapture trajectory was developed which demonstrated the guidance software executing on a flight-like computer, interfacing with a simulated IMU and simulated thrusters, with vehicle dynamics provided by an external simulator. Five cases from the NRT simulations were run in the RT simulation environment. The results compare well to those of the NRT simulation thus verifying the RT simulation configuration. The results of the above described simulations show the aerocapture maneuver using the APC algorithm can be accomplished reliably and the algorithm is now at TRL-6. Flight validation is the next step for aerocapture technology development.
A portable comprehensive navigational system has been developed that both robotic and human explorers can use to determine their location, attitude, and heading anywhere on the lunar surface independent of external infrastructure (needs no Lunar satellite network, line of sight to the Sun or Earth, etc.). The system combines robust processing power with an extensive topographical database to create a real-time atlas (GIS Geospatial Information System) that is able to autonomously control and monitor both single unmanned rovers and fleets of rovers, as well as science payload stations. The system includes provisions for teleoperation and tele-presence. The system accepts (but does not require) inputs from a wide range of sensors. A means was needed to establish a location when the search is taken deep in a crater (looking for water ice) and out of view of Earth or any other references. A star camera can be employed to determine the user's attitude in menial space and stellar map in body space. A local nadir reference (e.g., an accelerometer that orients the nadir vector in body space) can be used in conjunction with a digital ephemeris and gravity model of the Moon to isolate the latitude, longitude, and azimuth of the user on the surface. That information can be used in conjunction with a Lunar GIS and advanced navigation planning algorithms to aid astronauts (or other assets) to navigate on the Lunar surface.
The Gravity Recovery and Interior Laboratory (GRAIL) spacecraft conducted the mapping of the gravity field of the Moon from March 1, 2012 to May 29, 2012. The twin spacecraft acquired highly precise K Band range-rate (KBRR) intersatellite ranging data and Deep Space Network (DSN) data during this prime mission phase from altitudes of 15 to 75 km above the lunar surface over three lunar months. We have processed these data using the NASA GSFC GEODYN orbit determination and geodetic parameter estimation program, and we have determined gravity fields up to degree and order 420 in spherical harmonics. The new gravity solutions show improved correlations with LOLA-derived topography to high degree and order and resolve many lunar features in the geopotential with a resolution of less than 30 km, including for example the central peak of the crater Tycho. We discuss the methodology used for the processing of the GRAIL data, the quality of the orbit determination on the GRAIL satellites and the derivation of the solutions, and their evaluation with independent data, including Lunar Prospector. We show that with these new GRAIL gravity solutions, we can now fit the low altitude, extended mission Lunar Prospector tracking data better than with any previous gravity model that included the LP data.
The Gravity Recovery and Interior Laboratory (GRAIL) [1], NASA s eleventh Discovery mission, successfully executed its Primary Mission (PM) in lunar orbit between March 1, 2012 and May 29, 2012. GRAIL s Extended Mission (XM) initiated on August 30, 2012 and was successfully completed on December 14, 2012. The XM provided an additional three months of gravity mapping at half the altitude (23 km) of the PM (55 km), and is providing higherresolution gravity models that are being used to map the upper crust of the Moon in unprecedented detail.
Recent interest at NASA for the application of solar electric propulsion for the transfer of significant payloads in cislunar space has led to the development of high-fidelity simulations of such missions. With such transfers involving transfer times on the order of months, simulation time can be significant. In the past, the examination of such missions typically began with the use of lower-fidelity trajectory optimization tools such as SEPSPOT to develop and tune guidance laws which delivered optimal or near- optimal trajectories, where optimal is generally defined as minimizing propellant expenditure or time of flight. The transfer of these solutions to a high-fidelity simulation is typically an iterative process whereby the initial solution may nearly, but not precisely, meet mission objectives. Further tuning of the guidance algorithm is typically necessary when accounting for high-fidelity perturbations such as those due to more detailed gravity models, secondary-body effects, solar radiation pressure, etc. While trajectory optimization is a useful method for determining optimal performance metrics, algorithms which deliver nearly optimal performance with minimal tuning are an attractive alternative.
The irregular nature of the lunar gravity field will severely affect the orbit lifetime and behavior of future lunar small satellite missions. These spacecraft need stable orbits that do not require large deltaV budgets for station-keeping maneuvers. The initial classical elements of any lunar orbit are critical to address its stability and to comply with mission requirements. This publication identifies stable regions according to different initial conditions at the time of lunar orbit insertion (LOI). High fidelity numerical simulations with two different gravity models were performed. We focus in low altitude orbits where the dominant force in orbit propagation is the existence of unevenly distributed lunar mass concentrations. These orbits follow a periodic oscillation in some of the classical elements that is particularly useful for mission design. A set of orbital maintenance strategies for various mission concepts is presented.
As a part of the NASA initiative to collect a boulder off of an asteroid and return it to Lunar orbit, the Satellite Servicing Capabilities Office (SSCO) and NASA GSFC are developing an on-board relative terrain imaging navigation algorithm for the Asteroid Redirect Robotic Mission (ARRM). After performing several flybys and dry runs to verify and refine the shape, spin, and gravity models and obtain centimeter level imagery, the spacecraft will descend to the surface of the asteroid to capture a boulder and return it to Lunar Orbit. The algorithm implements Stereophotoclinometry methods to register landmarks with images taken onboard the spacecraft, and use these measurements to estimate the position and orientation of the spacecraft with respect to the asteroid. This paper will present an overview of the ARRM GN&C system and concept of operations as well as a description of the algorithm and its implementation. These techniques will be demonstrated for the descent to the surface of the proposed asteroid of interest, 2008 EV5, and preliminary results will be shown.
We describe the results of the Laser Ranging (LR) experiment carried out from June 2009 to September 2014 in order to make one-way time-of-flight measurements of laser pulses between Earth-based laser ranging stations and the Lunar Reconnaissance Orbiter (LRO) orbiting the Moon. Over 4,000 hours of successful LR data are obtained from 10 international ground stations. The 20-30 centimeter precision of the full-rate LR data is further improved to 5-10 centimeter after conversion into normal points. The main purpose of LR is to utilize the high accuracy normal point data to improve the quality of the LRO orbits, which are nomi- nally determined by the radiometric S-band tracking data. When independently used in the LRO precision orbit determination process with the high-resolution GRAIL (Gravity Recovery and Interior Laboratory) gravity model, LR data provide good orbit solutions, with an average difference of approximately 50 meters in total position, and approximately 20 centimeters in radial direction, compared to the definitive LRO trajectory. When used in combination with the S-band tracking data, LR data help to improve the orbit accuracy in the radial direction to approximately 15 centimeters. In order to obtain highly accurate LR range measurements for precise orbit determination results, it is critical to closely model the behavior of the clocks both at the ground stations and on the spacecraft. LR provides a unique data set to calibrate the spacecraft clock. The LRO spacecraft clock is characterized by the LR data to a timing knowledge of 0.015 milliseconds over the entire 5 years of LR operation. We here present both the engineering setup of the LR experiments and the detailed analysis results of the LR data.