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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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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Development of Formation Deployment and Intialization Concepts

NASA's Cross-Cutting Technology Development Program identified formation flying as a key enabler for the next generation Earth and Sciences campaign. It is hoped that this technology will allow a distributed network of autonomous satellites to act collaboratively as a single collective unit paving the way for extensive co-observing campaigns, coordinated multi-point observing programs, improved space-based interferometry, and entirely new approaches to conducting science. APL as a team member with GSFC, funded by the Earth Sciences and Technology Organization (ESTO), investigated formation deployment and initialization concepts which is central to the formation flying concept. This paper presents the analytical approach and preliminary results of the study. The study investigated a simple mission involving the deployment of six micro-satellites, one at a time, from a bus. At the initialization state, the satellites fly in an along-track trajectory separated by nominal spacing. The study entailed the development of a two-body (bus and satellite) relative motion propagator based on Clohessy-Wiltshire (C-W) equations with drag from which the relative motion of the micro-satellites is deduced. This code was used to investigate cluster development characteristics subject to "tip-off' (ejection) conditions. Results indicate that cluster development is very sensitive to the ballistic coefficients of the bus and satellites, and to relative ejection velocity. This information can be used to identify optimum deployment parameters, along with accuracy bounds for a particular mission, and to develop a cluster control strategy minimizing global fuel and cost. A suitable control strategy concept has been identified, however, it needs to be developed further.

Badesha, Surjit S.↗
NASA NTRS1972

Influence of satellite aerodynamics on atmospheric density determination.

Discussion of aerodynamic factors which influence the interpretation of satellite dynamic response. These factors include: (1) the influence of satellite orientation and shape on the drag coefficient; (2) the effect of changes in the gas flow properties with altitude; and (3) the influence of upper atmospheric winds on the interpretation of data. These factors represent the greatest source of error in current data reduction. For this reason, an estimate is made of a possible correction to present density models.

Karr, G. R.↗
NASA NTRS1987

Predicting Spacecraft Trajectories

Calculations include drag, nonsphericity, and other perturbations. Artificial Satellite Analysis Program (ASAP) is general orbit-prediction program that incorporates sufficient orbit-modeling accuracy for mission design, maneuver analysis, and mission planning. Suitable for studying planetary orbit missions with spacecraft trajectories of exploratory (mapping) nature. Sample data included for study of drift cycle of geosynchronous station, strategy for mapping Venus by radar, frozen orbit about Mars, and repeat-ground-trace orbit. ASAP written in FORTRAN 77.

Kwok, J.↗
NASA NTRS2006

A New Empirical Thermospheric Density Model JB2006 using New Solar Indices

A new empirical atmospheric density model is developed using the CIRA72 (Jacchia 71) model as the basis for the diffusion equations. New solar indices based on orbit based sensor data are used for the solar irradiances in the extreme and far ultraviolet wavelengths. New exospheric temperature and semiannual density equations are employed to represent the major thermospheric density variations. Temperature correction equations are also developed for diurnal and latitudinal effects, and finally density correction factors are used for model corrections required at high altitude (1500-4000 km). The new model, Jacchia-Bowman 2006, is validated through comparisons of accurate daily density drag data previously computed for numerous satellite. For 400 km altitude the standard deviation of 16% for the standard Jacchia model is reduced to 10% for the new JB2006 model for periods of low geomagnetic storm activity.

Bowman, Bruce R.↗
NASA NTRS1991

Simulation of GRM drag compensation system

NASA's Geopotential Research Mission (GRM) was proposed in 1986 for globally determining the earth's gravitational and magnetic fields with high precision via full earth coverage polar-orbit satellites. In the GRM system, at least one of the two satellites was required to be stationed in a low-altitude, 160-km orbit subject to atmospheric drag effects that could both corrupt gravity measurements and reduce the satellite's lifetime. The Disturbance Compensation System, 'DISCOS', was incorporated to select a drag-free orbit during the active portion of GRM satellite operations by firing thrusters that offset the effects of disturbances. A drag-free thruster algorithm has been created to simulate the DISCOS system; simulation results are presented.

Antreasian, Peter G.↗
NASA NTRS2018

Drag De-Orbit Device (D3): A Retractable Device for CubeSat Attitude and Orbit Control using Aerodynamic Forces

The increasing number of CubeSats being launched has raised concerns about orbital debris since most of these satellites have no means of active orbit control. Some technologies exist to increase the surface area of a CubeSat and expedite de-orbit due to aerodynamic drag in low Earth orbit, but most of these devices cannot be retracted and hence cannot be used for orbital maneuvering. This paper discusses the De-Orbit Drag Device (D3) module that is capable of de-orbiting a 12U, 15kg CubeSat from a 700 km circular orbit in under 25 years and can be deployed and retracted to modulate the aerodynamic drag force experienced by the satellite. This facilitates orbital maneuvering using aerodynamic drag and the active targeting of a de-orbit location. In addition, the geometry of this drag device provides 3-axis attitude stabilization of the host CubeSat using aerodynamic and gravity gradient torques which is useful for many missions and provides a predictable aerodynamic profile for use in orbital maneuvering algorithms.

algorithm↗
NASA NTRS1958

Three-Dimensional Orbits of Earth Satellites, Including Effects of Earth Oblateness and Atmospheric Rotation

The principal purpose of the present paper is to present sets of equations which may be used for calculating complete trajectories of earth satellites from outer space to the ground under the influence of air drag and gravity, including oblateness effects, and to apply these to several examples of entry trajectories starting from a circular orbit. Equations of motion, based on an "instantaneous ellipse" technique, with polar angle as independent variable, were found suitable for automatic computation of orbits in which the trajectory consists of a number of revolutions. This method is suitable as long as the trajectory does not become nearly vertical. In the terminal phase of the trajectories, which are nearly vertical, equations of motion in spherical polar coordinates with time as the independent variable were found to be more suitable. In the first illustrative example the effects of the oblateness component of the earth's gravitational field and of atmospheric rotation were studied for equatorial orbits. The satellites were launched into circular orbits at a height of 120 miles, an altitude sufficiently high that a number of revolutions could be studied. The importance of the oblateness component of the earth's gravitational field is shown by the fact that a satellite launched at circular orbital speed, neglecting oblateness, has a perigee some 67,000 feet lower when oblateness forces are included in the equations of motion than when they are not included. Also, the loss in altitude per revolution is double that of a satellite following an orbit not subject to oblateness. The effect of atmospheric rotation on the loss of altitude per revolution was small. As might be surmised, the regression of the line of nodes as predicted by celestial mechanics is unchanged when drag is included. It is clear that the inclination of the orbital plane to the equator will be relatively unaffected by drag for no atmospheric rotation since the drag lies in the orbital plane in this case. With the inclusion of atmospheric rotation it was found that the inclination of the plane changed about one-millionth of a radian per revolution. Thus the prediction of the position of the orbital plane of an earth satellite is not complicated by the introduction of drag. The line of apsides, which without drag but with oblateness moves slowly in space, tends to move with the satellite when drag is included in the calculations. As a results, the usual linearized solutions based on oblateness alone must be basically altered when drag is included to take into account the rapid movement of the line of apsides. In the second illustrative example the final revolution was calculated to impact for a number of trajectories in an orbital plane inclined at 650 to the equator. Of particular interest is the large effect the oblateness gravitational field and atmospheric rotation can have on the impact point. For a value of CDA/m of unity, and for an initial downward angle at 80 miles altitude of 0.01 radian, such as might be utilized for manned re-entry, oblateness had an influence of about 300 miles in the impact point, and atmospheric rotation had about a 150-mile influence.

Nielsen, Jack N.↗
NASA NTRS1976

IMS/Satellite Situation Center report: Orbit plots and bar charts for Prognoz 4, days 1-91 1976

Orbit plots for the Prognoz 4 satellite for the time period January to March 1976 are given. This satellite was identified as a possible important contributor to the International Magnetospheric Study project. The orbits were based on an element epoch of December 26, 1975, 3h 8min and 17s. In view of the low perigee of this satellite, the Satellite Situation Center (SSC) considered that the effect of atmospheric drag precludes orbit predictions for the length of time normally used by the SSC for high-altitude satellites. Consequently, orbit data are shown for the first 3 months of 1976 only. The orbit generated for this report was based on the earlier epoch, and it positions the satellite within 30s of the ascending node at the later epoch. Therefore, within the accuracy of the plots shown in this report, the orbit used was regarded as an achieved orbit. The orbit information is displayed graphically in four ways: bar charts, geocentric solar ecliptic plots, boundary plots, and solar magnetic latitude versus local time plots. The most concise presentation is the bar charts. The bar charts give the crude three-dimensional position of the satellite for each magnetospheric region.

Source record↗
NASA NTRSDate not supplied

Solar Cycle 25 Model Prediction Comparisons: an Engineering Perspective

The solar cycle 25 progression is ramping up, prompting satellite operators to provide estimates of propellent usage over the next several years. Drag due to atmospheric heating from solar activity is the main driver to satellite lifetime decay, especially for those in low-Earth orbit destinations. In this poster, several model predictions of the sunspot number are compared and discussed with comments on applicability to engineering programs.

Space Weather↗
NASA NTRS1976

Investigation of environmental perturbations on passive asymmetric satellite

The effects of environmental perturbations on the attitude of a slow tumbling earth-oriented satellite are investigated. The environmental perturbations considered were aerodynamic drag, gravity-gradient, solar radiation pressure, and magnetic torques. The Euler attitude equations were solved numerically for the Skylab spacecraft. Results are presented for both torque-free motion and for cases in which aerodynamic and gravity-gradient torques are acting in a slow tumble mode. Simulations show gravity-gradient effects on satellite momentum to be cyclic and to increase the precession rate of the angular momentum vector about the radius vector. This also tends to align the minor axis along the radius vector. Aerodynamic drag initially decreases angular momentum, slowly precesses the momentum vector about the radius vector, and finally drives the satellite into an unstable mode. Combined gravity-gradient and aerodynamic torques reduce angular momentum and energy, and induce a steady precession rate of the momentum vector about the radius vector.

Tate, V.↗
NASA NTRS1993

Orbital simulations of satellite escape/capture and the origin of satellites such as Triton

We investigate satellite escape/capture in the context of the restricted, circular three body problem as applied to the Sun, Neptune, and Triton. We have computed a large number of coplanar prograde and retrograde orbital simulations over a range of initial distances and velocities. The satellite starts at superior conjunction within approximately 2 Hill radii of Neptune and has a velocity orthogonal to the Sun-planet line. Orbits with these initial conditions can be reflected with respect to time, so an escape is simply the reverse of a capture. We numerically integrate the equations of motion to compute the satellite's position until it escapes, collides with Neptune, or after 100 planetary years fails to escape, when computations cease. The initial distance x and velocity v in the restricted problem uniquely define the Jacobi constant C, a conserved energy-like quantity. Plots of the simulation outcomes in the prograde and retrograde C, x phase spaces reveal distinct zones in which temporary satellites approach the planet closely enough that permanent capture can be effected by gas drag with a protoplanetary nebula or by collision with a pre-existing satellite. Single and double close-flybys constitute the most common possible capture orbits. Long term multiple flyby orbits occur near the stability limits between bound and unbound orbits, and are more common among retrograde captures.

Benner, Lance A. M.↗
NASA NTRS2012

Recent Simulations of the Late Stages Growth of Jupiter

Presented by Lissauer et al. (2009, Icarus 199, 338) are used to test the model of capture of Jupiter's irregular satellites within proto-Jupiter's distended and thermally-supported envelope. We find such capture highly unlikely, since the envelope shrinks too slowly for a large number of moons to be retained, and many of those that would be retained would orbit closer to the planet than do the observed Jovian irregulars. Our calculations do not address (and therefore do not exclude) the possibility that the irregular satellites were captured as a result of gas drag within a circumjovian disk. Support for this research from NASA Outer Planets Research Program is gratefully acknowledged.

circumjovian disk↗
NASA NTRSDate not supplied

TechEdSat 7, 10, 13, 15: Exo-brake Experiments on the ISS, First Virgin Orbit, and First Firefly-Alpha Test Flights.

The TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has been studying cube satellite re-entry technologies with a focus on Exo-Brake drag device research. An exo-atmospheric braking device, the Exo-Brake uses the tenuous upper atmosphere acting in a free molecular flow regime to enable controllable adjustment of the drag profile of the host spacecraft, currently enabling rapid disposal of small spacecraft after mission conclusion, and eventual targeted de-orbit to a desired entry point at the Von Karman line, or approximately 100km in altitude. The TES-7 mission, flown on the first successful Virgin Orbit flight January 17, 2021, was injected into a 500km, 61-degree inclination orbit, with an expected orbital lifetime of approximately ten years. After successful deployment of a unique hydrogen gas-cell inflation design ‘disposal’-type Mylar Exo-Brake intended to rapidly de-orbit the spacecraft, the maximum expected orbital lifetime of the TES-7 spacecraft was successfully reduced from ten years to 1.3 years, an 87% reduction in orbit lifetime via a passive, fuel-less system. The subsequent TES-13 mission utilized a new, less complex ‘disposal’-type Exo-Brake design made from collapsible rigid struts rather than inflated struts to reduce the engineering, construction, and safety complexities introduced by the prior hydrogen gas inflation design. The rigid strut Exo-Brake design is stowed via rotational compression and deploys using only stored spring energy once released via electronic actuator. This new rotation compression storage design has become the new Exo-Brake design standard used on TES missions thanks to its simplicity. Following on the success of this new Exo-Brake design demonstrated by TES-13, the design was further refined to enabled active manipulation of the Exo-Brake effective drag area though a winch-like system. Both the TES-10 and TES-15 spacecraft were equipped with geometrically similar Exo-Brake devices capable of such active drag modulation, with the intent of studying the operational impact of having such devices on cube satellite missions. To increase the survival time of the Exo-Brake at low altitudes and thus gain more experiment time and better guidance capability, TES-15 was equipped with a modulated Exo-Brake constructed from 3M™ Nextel™ 440 ceramic oxide fiber, the same fabric material used in the Space Shuttle TPS system, rather than Mylar. After launching on Alpha Flight 2 October 1st, 2022, Firefly’s first successful launch, TES-15 met the same modulation restrictions as TES-10, with modulation prohibited above 200km to avoid possible collision with other spacecraft. In this case, a low deployment orbit caused an extremely short mission life, which, coupled with delayed spacecraft identification, exacerbated collision concerns and caused performance data of the high-temperature Exo-Brake design to be inconclusive. As such, the Nano Orbital Workshop has taken a new approach of only conducting modulated Exo-Brake research on missions deploying below ISS altitude to avoid collision avoidance restrictions, and is working to improve spacecraft identification and location reporting techniques to increase experiment durations. Upcoming TES mission will therefore use one of two classes of Exo-Brake depending on their target altitude, as to be described in the submission.

M Murbach↗
NASA NTRSDate not supplied

TechEdSat 7, 10, 13, 15: Exo-brake Experiments on the ISS, First Virgin Orbit, and First Firefly-Alpha Test Flights

The TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has been studying cube satellite re-entry technologies with a focus on Exo-Brake drag device research. An exo-atmospheric braking device, the Exo-Brake uses the tenuous upper atmosphere acting in a free molecular flow regime to enable controllable adjustment of the drag profile of the host spacecraft, currently enabling rapid disposal of small spacecraft after mission conclusion, and eventual targeted de-orbit to a desired entry point at the Von Karman line, or approximately 100km in altitude. The TES-7 mission, flown on the first successful Virgin Orbit flight January 17, 2021, was injected into a 500km, 61-degree inclination orbit, with an expected orbital lifetime of approximately ten years. After successful deployment of a unique hydrogen gas-cell inflation design ‘disposal’-type Mylar Exo-Brake intended to rapidly de-orbit the spacecraft, the maximum expected orbital lifetime of the TES-7 spacecraft was successfully reduced from ten years to 1.3 years, an 87% reduction in orbit lifetime via a passive, fuel-less system. The subsequent TES-13 mission utilized a new, less complex ‘disposal’-type Exo-Brake design made from collapsible rigid struts rather than inflated struts to reduce the engineering, construction, and safety complexities introduced by the prior hydrogen gas inflation design. The rigid strut Exo-Brake design is stowed via rotational compression and deploys using only stored spring energy once released via electronic actuator. This new rotation compression storage design has become the new Exo-Brake design standard used on TES missions thanks to its simplicity. Following on the success of this new Exo-Brake design demonstrated by TES-13, the design was further refined to enabled active manipulation of the Exo-Brake effective drag area though a winch-like system. Both the TES-10 and TES-15 spacecraft were equipped with geometrically similar Exo-Brake devices capable of such active drag modulation, with the intent of studying the operational impact of having such devices on cube satellite missions. To increase the survival time of the Exo-Brake at low altitudes and thus gain more experiment time and better guidance capability, TES-15 was equipped with a modulated Exo-Brake constructed from 3M™ Nextel™ 440 ceramic oxide fiber, the same fabric material used in the Space Shuttle TPS system, rather than Mylar. After launching on Alpha Flight 2 October 1st, 2022, Firefly’s first successful launch, TES-15 met the same modulation restrictions as TES-10, with modulation prohibited above 200km to avoid possible collision with other spacecraft. In this case, a low deployment orbit caused an extremely short mission life, which, coupled with delayed spacecraft identification, exacerbated collision concerns and caused performance data of the high-temperature Exo-Brake design to be inconclusive. As such, the Nano Orbital Workshop has taken a new approach of only conducting modulated Exo-Brake research on missions deploying below ISS altitude to avoid collision avoidance restrictions, and is working to improve spacecraft identification and location reporting techniques to increase experiment durations. Upcoming TES mission will therefore use one of two classes of Exo-Brake depending on their target altitude, as to be described in the submission.

Marcus Murbach↗
NASA NTRS2011

Solar Radiation Pressure Binning for the Geosynchronous Orbit

Orbital maintenance parameters for individual satellites or groups of satellites have traditionally been set by examining orbital parameters alone, such as through apogee and perigee height binning; this approach ignored the other factors that governed an individual satellite's susceptibility to non-conservative forces. In the atmospheric drag regime, this problem has been addressed by the introduction of the "energy dissipation rate," a quantity that represents the amount of energy being removed from the orbit; such an approach is able to consider both atmospheric density and satellite frontal area characteristics and thus serve as a mechanism for binning satellites of similar behavior. The geo-synchronous orbit (of broader definition than the geostationary orbit -- here taken to be from 1300 to 1800 minutes in orbital period) is not affected by drag; rather, its principal non-conservative force is that of solar radiation pressure -- the momentum imparted to the satellite by solar radiometric energy. While this perturbation is solved for as part of the orbit determination update, no binning or division scheme, analogous to the drag regime, has been developed for the geo-synchronous orbit. The present analysis has begun such an effort by examining the behavior of geosynchronous rocket bodies and non-stabilized payloads as a function of solar radiation pressure susceptibility. A preliminary examination of binning techniques used in the drag regime gives initial guidance regarding the criteria for useful bin divisions. Applying these criteria to the object type, solar radiation pressure, and resultant state vector accuracy for the analyzed dataset, a single division of "large" satellites into two bins for the purposes of setting related sensor tasking and orbit determination (OD) controls is suggested. When an accompanying analysis of high area-to-mass objects is complete, a full set of binning recommendations for the geosynchronous orbit will be available.

Hejduk, M. D.↗
NASA NTRS1994

Tracking system options for future altimeter satellite missions

Follow-on missions to provide continuity in the observation of the sea surface topography once the successful TOPEX/POSEIDON (T/P) oceanographic satellite mission has ended are discussed. Candidates include orbits which follow the ground tracks of T/P GEOSAT or ERS-1. The T/P precision ephemerides, estimated to be near 3 cm root-mean-square, demonstrate the radial orbit accuracy that can be achieved at 1300 km altitude. However, the radial orbit accuracy which can be achieved for a mission at the 800 km altitudes of GEOSAT and ERS-1 has not been established, and achieving an accuracy commensurate with T/P will pose a great challenge. This investigation focuses on the radial orbit accuracy that can be achieved for a mission in the GEOSAT orbit. Emphasis is given to characterizing the effects of force model errors on the estimated radial orbit accuracy, particularly those due to gravity and drag. The importance of global, continuous tracking of the satellite for reduction in these sources of orbit error is demonstrated with simulated GPS tracking data. A gravity tuning experiment is carried out to show how the effects of gravity error may be reduced. Assuming a GPS flight receiver with a full-sky tracking capability, the simulation results indicate that a 5 cm radial orbit accuracy for an altimeter satellite in GEOSAT orbit should be achievable during low-drag atmospheric conditions and after an acceptable tuning of the gravity model.

Davis, G. W.↗
NASA NTRS1982

The dynamics of planetary rings

The physical processes that occur in planetary rings are discussed. The theoretical arguments leading to the conclusion that Saturn's rings are solid particles in nonuniform rotation are summarized, and the optical depth, thickness, and particle size of the rings are discussed. The influence of nearby satellites on the rings is analyzed, and asymmetries in the rings are briefly discussed. What is known of the rings of Uranus and of Jupiter's ring is summarized. Some of the dynamical processes and influences that are expected to be incorporated in more advanced theories of planetary rings are reviewed in detail, including radiation drag, plasma drag, interparticle collisions and viscosity, resonances with external satellites, shepherd satellites and moonlets. Finally, the orbital evolution of the shepherd satellites caused by the rings is estimated.

Goldreich, P.↗