Engineering topics

Beckman, Mark

Publications and source records attributed to Beckman, Mark.

At least 19 records

NASA NTRS2013

OSIRIS-REx Touch-And-Go (TAG) Mission Design and Analysis

The Origins Spectral Interpretation Resource Identification Security Regolith Explorer (OSIRIS-REx) mission is a NASA New Frontiers mission launching in 2016 to rendezvous with the near-Earth asteroid (101955) 1999 RQ36 in late 2018. After several months in formation with and orbit about the asteroid, OSIRIS-REx will fly a Touch-And-Go (TAG) trajectory to the asteroid s surface to obtain a regolith sample. This paper describes the mission design of the TAG sequence and the propulsive maneuvers required to achieve the trajectory. This paper also shows preliminary results of orbit covariance analysis and Monte-Carlo analysis that demonstrate the ability to arrive at a targeted location on the surface of RQ36 within a 25 meter radius with 98.3% confidence.

Berry, Kevin↗

NASA NTRS2007

Stationkeeping for the Lunar Reconnaissance Orbiter (LRO)

The Lunar Reconnaissance Orbiter (LRO) is scheduled to launch in 2008 as the first mission under NASA's Vision for Space Exploration. Following several weeks in a quasi-frozen commissioning orbit, LRO will fly in a 50 km mean altitude lunar polar orbit. During the one year mission duration, the orbital dynamics of a low lunar orbit force LRO to perform periodic sets of stationkeeping maneuvers. This paper explores the characteristics of low lunar orbits and explains how the LRO stationkeeping plan is designed to accommodate the dynamics in such an orbit. The stationkeeping algorithm used for LRO must meet five mission constraints. These five constraints are to maintain ground station contact during maneuvers, to control the altitude variation of the orbit, to distribute periselene equally between northern and southern hemispheres, to match eccentricity at the beginning and the end of the sidereal period, and to minimize stationkeeping deltaV. This paper addresses how the maneuver plan for LRO is designed to meet all of the above constraints.

Beckman, Mark↗

NASA NTRS2007

Mission Design for the Lunar Reconnaissance Orbiter

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.

Beckman, Mark↗

NASA NTRS2007

Stationkeeping for the Lunar Reconnaissance Orbiter (LRO)

The Lunar Reconnaissance Orbiter (LRO) is scheduled to launch in 2008 as the first mission under NASA's Vision for Space Exploration. Follo wing several weeks in a quasi-frozen commissioning orbit, LRO will fl y in a 50 km mean altitude lunar polar orbit. During the one year mis sion duration, the orbital dynamics of a low lunar orbit force LRO to perform periodic sets of stationkeeping maneuvers. This paper explor es the characteristics of low lunar orbits and explains how the LRO s tationkeeping plan is designed to accommodate the dynamics in such an orbit. The stationkeeping algorithm used for LRO must meet five miss ion constraints. These five constraints are to maintain ground statio n contact during maneuvers, to control the altitude variation of the orbit, to distribute periselene equally between northern and southern hemispheres, to match eccentricity at the beginning and the end of the sidereal period, and to minimize stationkeeping (Delta)V. This pape r addresses how the maneuver plan for LRO is designed to meet all of the above constraints.

Beckman, Mark↗

NASA NTRS2005

Finding Acceptable James Webb Space Telescope Mission Orbits From a Fixed Ariane Flight Profile

The James Webb Space Telescope (JWST) will be launched into orbit about the Sun/Earth L2 libration point. Trajectory design was recently completed which included expected separation states from the Ariane launch vehicle, constraints such as eclipses, maximum orbit size, maximum Sun-Vehicle-Earth/Moon angles, and launch opportunities. The results of this trajectory design give a set of possible trajectories for JWST with bounded stray light zones and provide a complete launch window. This data is also used to design the initial trajectory correction maneuver such that a maneuver towards the Sun is not required.

Beckman, Mark↗

NASA NTRS2003

Navigation Concepts for the James Webb Space Telescope

This paper evaluates the performance that can be achieved using candidate ground and onboard navigation approaches for operation of the James Webb Space Telescope, which will be in an orbit about the Sun-Earth L2 libration point. The ground navigation approach processes standard range and Doppler measurements from the Deep Space Network The onboard navigation approach processes celestial object measurements and/or ground-to- spacecraft Doppler measurements to autonomously estimate the spacecraft s position and velocity and Doppler reference frequency. Particular attention is given to assessing the absolute position and velocity accuracy that can be achieved in the presence of the frequent spacecraft reorientations and momentum unloads planned for this mission. The ground navigation approach provides stable navigation solutions using a tracking schedule of one 30-minute contact per day. The onboard navigation approach that uses only optical quality celestial object measurements provides stable autonomous navigation solutions. This study indicates that unmodeled changes in the solar radiation pressure cross-sectional area and modeled momentum unload velocity changes are the major error sources. These errors can be mitigated by modeling these changes, by estimating corrections to compensate for the changes, or by including acceleration measurements.

Long, Anne↗

NASA NTRS2003

Use of Invariant Manifolds for Transfers Between Three-Body Systems

The Lunar L1 and L2 libration points have been proposed as gateways granting inexpensive access to interplanetary space. To date, only individual solutions to the transfer between three-body systems have been found. The methodology to solve the problem for arbitrary three-body systems and entire families of orbits does not exist. This paper presents the initial approaches to solve the general problem for single and multiple impulse transfers. Two different methods of representing and storing 7-dimensional invariant manifold data are presented. Some particular solutions are presented for the transfer problem, though the emphasis is on developing methodology for solving the general problem.

Beckman, Mark↗

NASA NTRS2003

James Webb Space Telescope Station-keeping

The James Webb Space Telescope (JWST) is planned to be launched in 2011 to the Sun- Earth L2 libration point. The resultant delta-Vs (dV) from momentum unloads will perturb the orbit and necessitate frequent station-keeping maneuvers. The station-keeping dV budget is highly sensitive to the direction of the resultant dV vector. A simple spacecraft reorientation prior to each momentum unload will allow some control over the direction of the resultant dV vector. For each inertial momentum vector direction, an optimum spacecraft attitude is determined which gives a resultant dV vector that requires the least amount of station-keeping dV. Using this procedure, the station-keeping dV budget for JWST can be reduced by 60%.

Beckman, Mark↗

NASA NTRS2002

Libration Orbit Mission Design: Applications of Numerical & Dynamical Methods

Sun-Earth libration point orbits serve as excellent locations for scientific investigations. These orbits are often selected to minimize environmental disturbances and maximize observing efficiency. Trajectory design in support of libration orbits is ever more challenging as more complex missions are envisioned in the next decade. Trajectory design software must be further enabled to incorporate better understanding of the libration orbit solution space and thus improve the efficiency and expand the capabilities of current approaches. The Goddard Space Flight Center (GSFC) is currently supporting multiple libration missions. This end-to-end support consists of mission operations, trajectory design, and control. It also includes algorithm and software development. The recently launched Microwave Anisotropy Probe (MAP) and upcoming James Webb Space Telescope (JWST) and Constellation-X missions are examples of the use of improved numerical methods for attaining constrained orbital parameters and controlling their dynamical evolution at the collinear libration points. This paper presents a history of libration point missions, a brief description of the numerical and dynamical design techniques including software used, and a sample of future GSFC mission designs.

Bauer, Frank↗

NASA NTRS2002

Orbit Determination Issues for Libration Point Orbits

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

Beckman, Mark↗

NASA NTRS2002

Orbit Determination Issues for Libration Point Orbits

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

Beckman, Mark↗

NASA NTRS2002

Servicing and Deployment of National Resources in Sun-Earth Libration Point Orbits

Spacecraft travel between the Sun-Earth system, the Earth-Moon system, and beyond has received extensive attention recently. The existence of a connection between unstable regions enables mission designers to envision scenarios of multiple spacecraft traveling cheaply from system to system, rendezvousing, servicing, and refueling along the way. This paper presents examples of transfers between the Sun-Earth and Earth-Moon systems using a true ephemeris and perturbation model. It shows the (Delta)V costs associated with these transfers, including the costs to reach the staging region from the Earth. It explores both impulsive and low thrust transfer trajectories. Additionally, analysis that looks specifically at the use of nuclear power in libration point orbits and the issues associated with them such as inadvertent Earth return is addressed. Statistical analysis of Earth returns and the design of biased orbits to prevent any possible return are discussed. Lastly, the idea of rendezvous between spacecraft in libration point orbits using impulsive maneuvers is addressed.

Folta, David C.↗

NASA NTRS1999

Lunar Prospector Extended Mission

The National Aeronautics and Space Administration (NASA) selected Lunar Prospector (LP) as one of the discovery missions to conduct solar system exploration science investigations. The mission is NASA's first lunar voyage to investigate key science objectives since Apollo and was launched in January 1998. In keeping with discovery program requirements to reduce total mission cost and utilize new technology, Lunar Prospector's mission design and control focused on the use of innovative and proven trajectory analysis programs. As part of this effort, the Ames Research Center and the Goddard Space Flight Center have become partners in the Lunar Prospector trajectory team to provide the trajectory analysis, maneuver planning, orbit determination support, and product generation. At the end of 1998, Lunar Prospector completed its one-year primary mission at 100 km altitude above the lunar surface. On December 19, 1998, Lunar Prospector entered the extended mission phase. Initially the mission orbit was lowered from 100 km to a mean altitude of 40 km. The altitude of Lunar Prospector varied between 25 and 55 km above the mean lunar geode due to lunar potential effects. After one month, the lunar potential model was updated based upon the new tracking data at 40 km. On January 29, 1999, the altitude was lowered again to a mean altitude of 30 km. This altitude varies between 12 and 48 km above the mean lunar geode. Since the minimum altitude is very close to the mean geode, various approaches were employed to get accurate lunar surface elevation including Clementine altimetry and line of sight analysis. Based upon the best available terrain maps, Lunar Prospector will reach altitudes of 8 km above lunar mountains in the southern polar and far side regions. This extended mission phase of six months will enable LP to obtain science data up to 3 orders of magnitude better than at the mission orbit. This paper details the trajectory design and orbit determination planning and actual results of the Lunar Prospector extended mission including maneuver design, eccentricity & argument of perigee evolution, and lunar potential modeling.

Folta, David↗

NASA NTRS1999

Lunar Prospector Extended Mission

The National Aeronautics and Space Administration (NASA) selected Lunar Prospector as one of the discovery missions to conduct solar system exploration science investigations. The mission is NASA's first lunar voyage to investigate key science objectives since Apollo and was launched in January 1998. In keeping with discovery program requirements to reduce total mission cost and utilize new technology, Lunar Prospector's mission design and control focused on the use of innovative and proven trajectory analysis programs. As part of this effort, the Ames Research Center and the Goddard Space Flight Center have become partners in the Lunar Prospector trajectory team to provide the trajectory analysis, maneuver planning, orbit determination support, and product generation. At the end of 1998, Lunar Prospector completed its one-year primary mission at 100 km altitude above the lunar surface. On December 19, 1998, Lunar Prospector entered the extended mission phase. Initially the mission orbit was lowered from 100 km to a mean altitude of 40 km. The altitude of Lunar Prospector varied between 25 and 55 km above the mean lunar geode due to lunar potential effects. After one month, the lunar potential model was updated based upon the new tracking data at 40 km. On January 29, 1999, the altitude was lowered again to a mean altitude of 30 km. This altitude varies between 12 and 48 km above the mean lunar geode. Since the minimum altitude is very close to the mean geode, various approaches were employed to get accurate lunar surface elevation including Clementine altimetry and line of sight analysis. Based upon the best available terrain maps, Lunar Prospector will reach altitudes of 8 km above lunar mountains in the southern polar and far side regions. This extended mission phase of six months will enable LP to obtain science data up to 3 orders of magnitude better than at the mission orbit. This paper details the trajectory design and orbit determination planning, and actual results of the the Lunar Prospector extended mission including maneuver design, eccentricity & argument of perigee evolution, and lunar potential modeling.

Folta, David↗

NASA NTRS1999

The Lunar Prospector Mission: Final Results of Trajectory Design, Quasi-Frozen Orbits, Extended Mission Targeting, and Lunar Topography and Potential Models

The National Aeronautics and Space Administration (NASA) selected Lunar Prospector (LP) as one of the discovery missions to conduct solar system exploration science investigations. The mission was NASA's first lunar voyage to investigate key science objectives since Apollo and was launched in January 1998. In keeping with discovery program requirements to reduce total mission cost and utilize new technology, Lunar Prospector's mission design and control will focus on the use of innovative and proven trajectory analysis programs. As part of this effort, the Goddard Space Flight Center and the Ames Research Center became partners in the Lunar Prospector trajectory team to provide the trajectory analysis and orbit determination support. At the end of 1998, Lunar Prospector completed its one-year primary mission at 100-km altitude above the lunar surface. On December 19, 1998, Lunar Prospector entered its extended mission phase. The mission orbit was lowered from 100 km to a mean altitude of 40 km. Due to lunar potential effects, the altitude of Lunar Prospector varied from 25 to 55 km above the mean lunar geoid. After one month at 40 km, the lunar potential model was updated based upon the new tracking data at 40 km. On January 15, 1999, the altitude was lowered again to a mean altitude of 30 km. The spherical altitude varied between 15 km and 45 km above the mean lunar geoid while the topographical altitude varied between 10 km and 50 km. Various means were employed to get accurate lunar surface elevation including Clementine altimetry and LOS analysis. Based upon the best available terrain maps, Lunar Prospector reached actual altitudes of 8 km above lunar mountains in the southern polar region. This extended mission phase of six months will enable LP to obtain science data up to 3 orders of magnitude better than at the mission orbit. At the end of the operations mission, LP was targeted for impact at a chosen location that allowed optical observation of the lunar ejecta as LP ended its mission at 1.6 km/sec. This paper details the trajectory design and orbit determination planning and actual results of the Lunar Prospector nominal and extended mission including maneuver design, eccentricity vs. argument of perigee evolution, topographical altitude estimation, and lunar potential modeling. This paper provides understanding of the quasi-frozen orbit design of the LP mission, the optimization process of lunar orbit targets, the impacts that the selected lunar potential models play, and discusses the feasibility of meeting the mission goals. Observed evolution of the Keplerian orbit elements are compared to the theoretical predictions using the latest lunar potential model available which incorporates the Lunar Prospector Doppler data. Mapping orbit maintenance maneuver design along with results of the actual maneuvers to maintain the orbital requirements are also presented.

Folta, David↗

NASA NTRS1998

Lunar prospector mission design and trajectory support

The Lunar Prospector mission is the first dedicated NASA lunar mapping mission since the Apollo Orbiter program which was flown over 25 years ago. Competitively selected under the NASA Discovery Program, Lunar Prospector was launched on January 7, 1998 on the new Lockheed Martin Athena 2 launch vehicle. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with three scheduled orbit correction maneuvers to remove launch and cislunar injection errors prior to lunar insertion. At lunar encounter, a series of three lunar orbit insertion maneuvers and a small circularization burn were executed to achieve a 100 km altitude polar mapping orbit. This paper will present the design of the Lunar Prospector transfer, lunar insertion and mapping orbits, including maneuver and orbit determination strategies in the context of mission goals and constraints. Contingency plans for handling transfer orbit injection and lunar orbit insertion anomalies are also summarized. Actual flight operations results are discussed and compared to pre-launch support analysis.

Lozier, David↗

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 NTRS1997

Lunar Prospector Mission Design

The National Aeronautics and Space Administration (NASA) has selected Lunar Prospector as one of the discovery missions to conduct solar system exploration science investigations. The mission is NASA's first lunar voyage to investigate key science objectives since Apollo and will be launched in September 1997. In keeping with discovery program requirements to reduce total mission cost and utilize new technology, Lunar Prospector's mission design and control will focus on the use of innovative and proven trajectory analysis programs. As part of this effort, the Ames Research Center and the Goddard Space Flight Center have become partners in the Lunar Prospector trajectory team to provide the trajectory analysis and orbit determination support. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with an injection into a circular 100 km altitude polar mapping orbit. The mission design will incorporate the optimization of the reference trajectory for a new launch vehicle. An extended mission includes a periselene altitude approaching 10 km. As part of the navigation, a new lunar potential model will be used that incorporates data from the recent Clementine mission. This paper provides understanding of the design of the Lunar Prospector mission and the optimization process of the nominal trajectory, orbit determination, launch windows, and lunar orbit targets, and discusses the feasibility of meeting the mission goals.

Folta, David↗