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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 217 records · Page 12
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NASA NTRS2008

Light Echos in Kerr Geometry: A Source of High Frequency QPOs from Random X-ray Bursts

We propose that high frequency quasi-periodic oscillations (HFQPOs) can be produced from randomly-formed X-ray bursts (flashes) by plasma interior to the ergosphere of a rapidly-rotating black hole. We show by direct computation of their orbits that the photons comprising the observed X-ray light curves, if due to a multitude of such flashes, are affected significantly by the black hole's dragging of inertial frames; the photons of each such burst arrive to an observer at infinity in multiple (double or triple), distinct 'bunches' separated by a roughly constant time lag of t/M approximately equal to 14, regardless of the bursts' azimuthal position. We argue that every other such 'bunch' represents photons that follow trajectories with an additional orbit around the black hole at the photon circular orbit radius (a photon 'echo'). The presence of this constant lag in the response function of the system leads to a QPO feature in its power density spectra, even though the corresponding light curve consists of a totally stochastic signal. This effect is by and large due to the black hole spin and is shown to gradually diminish as the spin parameter a decreases or the radial position of the burst moves outside the static limit surface (ergosphere). Our calculations indicate that for a black hole with Kerr parameter of a/M=0.99 and mass of M=10*Msun the QPO is expected at a frequency of approximately 1.3-1.4 kHz. We discuss the plausibility and observational implications of our model/results as well as its limitations.

Fukumura, K.↗
NASA NTRS2008

Light Echoes in Kerr Geometry: A Source of High Frequency QPOs from Random X-ray Bursts

We propose that high frequency quasi-periodic oscillations (HFQPOs) can be produced from randomly-formed X-ray bursts (flashes) by plasma interior to the ergosphere of a rapidly-rotating black hole. We show by direct computation of their orbits that the photons comprising the observed X-ray light curves, if due to a multitude of such flashes, are affected significantly by the black hole's dragging of inertial frames; the photons of each such burst arrive to an observer at infinity in multiple (double or triple), distinct "bunches" separated by a roughly constant time lag of Deltat(t(sub lag))/M approx. 14, regardless of the bursts' azimuthal position. We argue that every other such "bunch" represents photons that follow trajectories with an additional orbit around the black hole at the photon circular orbit radius (a photon "echo"). The presence of this constant lag in the response function of the system leads to a QPO feature in its power density spectra, even though the corresponding light curve consists of a totally stochastic signal. This effect is by and large due to the black hole spin and is shown to gradually diminish as the spin parameter a decreases or the radial position of the burst moves outside the static limit surface (ergosphere). Our calculations indicate that for a black hole with Kerr parameter of a/M = 0.99 and mass of M = 10Stellar Mass the QPO is expected at a frequency of v(sub QPO) approx. 1.3 - 1.4 kHz. We discuss the plausibility and observational implications of our model/results as well as its limitations. Subject headings: accretion, accretion disks - black hole physics - X-rays: galaxies - stars: oscillations

Fukumura, Keigo↗
NASA NTRS2009

High Frequency QPOs due to Black Hole Spin

We present detailed computations of photon orbits emitted by flares at the innermost stable circular orbit (ISCO) of accretion disks around rotating black holes. We show that for sufficiently large spin parameter, i.e. a > 0.94 M, flare a sufficient number of photons arrive at an observer after multiple orbits around the black hole, to produce an "photon echo" of constant lag, i.e. independent of the relative phase between the black hole and the observer, of T approximates 14 M. This constant time delay, then, leads to a power spectrum with a QPO at a frequency nu approximates 1/14M, even for a totally random ensemble of such flares. Observation of such a QPO will provide incontrovertible evidence for the high spin of the black hole and a very accurate, independent, measurement of its mass.

Kazanas, Demos↗
NASA NTRS2010

Altair Descent and Ascent Reference Trajectory Design and Initial Dispersion Analyses

The Altair Lunar Lander is the linchpin in the Constellation Program (CxP) for human return to the Moon. Altair is delivered to low Earth orbit (LEO) by the Ares V heavy lift launch vehicle, and after subsequent docking with Orion in LEO, the Altair/Orion stack is delivered through translunar injection (TLI). The Altair/Orion stack separating from the Earth departure stage (EDS) shortly after TLI and continues the flight to the Moon as a single stack. Altair performs the lunar orbit insertion (LOI) maneuver, targeting a 100-km circular orbit. This orbit will be a polar orbit for missions landing near the lunar South Pole. After spending nearly 24 hours in low lunar orbit (LLO), the lander undocks from Orion and performs a series of small maneuvers to set up for descending to the lunar surface. This descent begins with a small deorbit insertion (DOI) maneuver, putting the lander on an orbit that has a perilune of 15.24 km (50,000 ft), the altitude where the actual powered descent initiation (PDI) commences. At liftoff from Earth, Altair has a mass of 45 metric tons (mt). However after LOI (without Orion attached), the lander mass is slightly less than 33 mt at PDI. The lander currently has a single descent module main engine, with TBD lb(sub f) thrust (TBD N), providing a thrust-to-weight ratio of approximately TBD Earth g's at PDI. LDAC-3 (Lander design and analysis cycle #3) is the most recently closed design sizing and mass properties iteration. Upgrades for loss of crew (LDAC-2) and loss of mission (LDAC-3) have been incorporated into the lander baseline design (and its Master Equipment List). Also, recently, Altair has been working requirements analyses (LRAC-1). All nominal data here are from the LDAC-3 analysis cycle. All dispersions results here are from LRAC-1 analyses.

Kos, Larry D.↗
NASA NTRS2011

Design Considerations for a Dedicated Gravity Recovery Satellite Mission Consisting of Two Pairs of Satellites

Future satellite missions dedicated to measuring time-variable gravity will need to address the concern of temporal aliasing errors; i.e., errors due to high-frequency mass variations. These errors have been shown to be a limiting error source for future missions with improved sensors. One method of reducing them is to fly multiple satellite pairs, thus increasing the sampling frequency of the mission. While one could imagine a system architecture consisting of dozens of satellite pairs, this paper explores the more economically feasible option of optimizing the orbits of two pairs of satellites. While the search space for this problem is infinite by nature, steps have been made to reduce it via proper assumptions regarding some parameters and a large number of numerical simulations exploring appropriate ranges for other parameters. A search space originally consisting of 15 variables is reduced to two variables with the utmost impact on mission performance: the repeat period of both pairs of satellites (shown to be near-optimal when they are equal to each other), as well as the inclination of one of the satellite pairs (the other pair is assumed to be in a polar orbit). To arrive at this conclusion, we assume circular orbits, repeat groundtracks for both pairs of satellites, a 100-km inter-satellite separation distance, and a minimum allowable operational satellite altitude of 290 km based on a projected 10-year mission lifetime. Given the scientific objectives of determining time-variable hydrology, ice mass variations, and ocean bottom pressure signals with higher spatial resolution, we find that an optimal architecture consists of a polar pair of satellites coupled with a pair inclined at 72deg, both in 13-day repeating orbits. This architecture provides a 67% reduction in error over one pair of satellites, in addition to reducing the longitudinal striping to such a level that minimal post-processing is required, permitting a substantial increase in the spatial resolution of the gravity field products. It should be emphasized that given different sets of scientific objectives for the mission, or a different minimum allowable satellite altitude, different architectures might be selected.

Wiese, D. N.↗
NASA NTRS2020

Q4 – A CubeSat Mission to Demonstrate Omnidirectional Optical Communications

We are proposing a technology demonstration mission for JPL’s Inter-Satellite Optical Communicator (ISOC). The ISOC has the potential to enable up to 1 Gbps data rates over a distance up to 200 km in free space. Key features of the ISOC include full sky coverage and the ability to maintain multiple links simultaneously. The Q4 mission consists of (4) 6U CubeSats furnished with ISOCs to demonstrate high data rate communications among the spacecraft. More specifically, we seek to demonstrate the omnidirectionality of the ISOC by providing simultaneous optical links between one spacecraft and the other three. Q4 will be a demonstrator for small spacecraft swarm capabilities. Following a proposed deployment from the ISS, the CubeSats will utilize their onboard thrusters to enter their desired orbit. The orbit consists of one “leader” spacecraft in a circular orbit with a 400 km LEO altitude and three “follower” spacecraft in slightly elliptical orbits rotating azimuthally about the leader. The main mission will consist of three consecutive phases: First, the spacecraft will demonstrate pointing, acquisition, and tracking using newly developed protocols. Next, link establishment and different channel access methods (CDMA and TDMA) will be tested. Finally, Delay Tolerant Network protocols will be implemented to allow fast data sharing among the spacecraft. The distance between the communicating spacecraft will be increased as the mission progresses to characterize the link quality as a function of distance. All communications over the course of the mission will be recorded and analyzed on the ground using bit error rate as a metric for success. The current design for each the four Q4 CubeSats contains an XACT advanced ADCS system for precision beam pointing. Other components currently under consideration include MiPS cold gas thrusters for orbital maneuvering and gimbaled eHaWK 84W solar arrays. In this paper we present design considerations for the Q4 CubeSats, link budget calculations, results from preliminary mission analysis, and expected results.

Velazco, Jose E↗
NASA NTRSDate not supplied

Assessing Utility of Different Orbits for a Lunar PNT Constellation

"As the Artemis program establishes a long-term presence on the lunar surface, lunar surface users will need an accurate source of Position, Navigation, and Timing (PNT). Under the LunaNet framework, NASA has proposed the Lunar Navigation System (LNS), a constellation of satellites in lunar orbit which will broadcast LunaNet’s Augmented Forward Signal (AFS) [1] to the lunar surface. LNS will have the objective of providing user position accuracies of 10 m and timing accuracies of 10 microseconds on the lunar surface and up to 200 km above the surface. Several classes of orbits are possibilities for the LNS constellation, including circular orbits, eccentric orbits, and various families of 3-body orbits. Each of these orbit classes have both advantages and disadvantages in their utility as PNT constellations. In this study we examine a wide variety of potential LNS constellations and evaluate their performance against several metrics, including: a. Dilution of Precision (DOP) performance over the lunar service volume and the South Pole region, during both the constellation buildup and after the constellation is completed. b. Number of satellites required to achieve required DOP performance. c. Relative dynamics of LNS satellites with respect to surface users. d. Signal in space error (SISE) performance, assuming orbit determination that includes the Global Positioning System (GPS), crosslinks, and monitor sites on the surface of the moon. e. Constellation insertion, station keeping, and disposal costs. f. Specific Concept of Operations (CONOPS) and engineering challenges associated with establishing and operating the constellations. (E.g., frequent station keeping, variable signal strength, divergent CONOPS for different planes of the constellation.) The system’s performance is considered both independently and in combination with other proposed systems, such as NASA’s Lunar Communication Relay and Navigation System (LCRNS) and ESA’s Moonlight constellation. The study concludes that a 3-plane Walker constellation at 6 lunar radii (slightly higher than GPS’s 4.2 earth radii from a geometry perspective, but significantly closer in an absolute sense) is a good candidate for the LNS constellation, providing the user with good SISE and DOP performance, while not requiring excessive station keeping or complex CONOPS. However, other viable concepts exist and could be chosen if early South Pole coverage is a driving factor in the system design."

navigation↗
NASA NTRS1977

Orbital evolution

The orbital evolution of a large satellite is governed primarily by tidal interactions between the satellite and the planet it orbits. Tides raised on a planet by a satellite transfer energy and angular momentum to the satellite orbit; this changes the semimajor axes of satellite orbits, increasing the size of those orbits where the satellite mean motion is smaller than the planetary angular velocity, and decreasing those where the opposite is true. Substantial changes caused by such tides for satellites of the terrestrial planets may explain the absence of satellites about Mercury and Venus. For Jovian and Saturnian satellites, such tides probably are only important in bringing about some of the observed orbital resonances. Tides raised on satellites generally cause decreasing orbital eccentricities, indicating why close satellites always have nearly circular orbits. Different processes of orbital evolution dominate for small bodies; their effects probably are critical in positioning material in the primordial dust cloud so that satellite coagulation may occur. A qualitative description is given of the orbital results of gas drag, radiation pressure, Poynting-Robertson drag and electromagnetic forces.

Burns, J. A.↗
NASA NTRS1989

Ascent, abort, and entry capability assessment of a Space Station rescue and personnel/logistics vehicle

The ascent, abort, and entry capability of a vehicle for rescue of the Space Station crew or ferry of personnel and supplies to and from the Space Station are analyzed. The configurations of the Space Station rescue vehicle and the Space Station personnel/logistics vehicle (SSPLV) are discussed and illustrated. The nominal ascent trajectory for the SSPLV delivered to orbit on a Titan II is presented. The ascent abort modes from launch to orbital injection are evaluated. It is shown that five landing sites with runways longer than 10,000 ft could provide a landing opportunity from every orbit of the Space Station in a 220-nmi circular orbit with a 28.5 deg inclination.

Naftel, J. C.↗
NASA NTRS1995

Synthetic-Aperture Coherent Imaging From A Circular Path

Imaging algorithms based on exact point-target responses. Developed for use in reconstructing image of target from data gathered by radar, sonar, or other transmitting/receiving coherent-signal sensory apparatus following circular observation path around target. Potential applications include: Wide-beam synthetic-aperture radar (SAR) from aboard spacecraft in circular orbit around target planet; SAR from aboard airplane flying circular course at constant elevation around central ground point, toward which spotlight radar beam pointed; Ultrasonic reflection tomography in medical setting, using one transducer moving in circle around patient or else multiple transducers at fixed positions on circle around patient; and Sonar imaging of sea floor to high resolution, without need for large sensory apparatus.

Jin, Michael Y.↗
NASA NTRS2010

Finding a 24 Day Orbital Period for the X-Ray Binary 1A 1118-616

We report the first determination of the binary period and the orbital ephemeris of the Be X-ray binary containing the pulsar IA 1118-616 (35 years after the discovery of the source). The orbital period is found to be P(sub orb) = 24.0+/-0.4 days. The source was observed by RXTE during its last big X-ray outburst in January 2009, peaking at MJD 54845.4. This outburst was sampled by taking short observations every few days, covering an elapsed time comparable to the orbital period. Using the phase connection technique, pulse arrival time delays could be measured and an orbital solution determined. The data are consistent with a circular orbit, the time of 90 degrees longitude was found to be T,/2 = MJD 54845.37(10), coincident with the peak X-ray flux.

Staubert, R.↗
NASA NTRS2011

Michigan Orbital DEbris Survey Telescope Observations of the Geosynchronous Orbital Debris Environment. Observing Years: 2007-2009

NASA uses the Michigan Orbital DEbris Survey Telescope (MODEST), the University of Michigan's 0.61-m aperture Curtis-Schmidt telescope at the Cerro Tololo Inter-American Observatory in Chile, to help characterize the debris environment in geosynchronous orbit; this began in February 2001 and continues to the present day. Detected objects that are found to be on the U.S. Space Surveillance Network cataloged objects list are termed correlated targets (CTs), while those not found on the list are called uncorrelated targets (UCTs). This Johnson Space Center report provides details of observational and data-reduction processes for the entire MODEST dataset acquired in calendar years (CYs) 2007, 2008, and 2009. Specifically, this report describes the collection and analysis of 36 nights of data collected in CY 2007, 43 nights of data collected in CY 2008, and 43 nights of data collected in CY 2009. MODEST is equipped with a 2048 x 2048-pixel charged coupled device camera with a 1.3 by 1.3 deg field of view. This system is capable of detecting objects fainter than 18th magnitude (R filter) using a 5-s integration. This corresponds to a 20-cm diameter, 0.175-albedo object at 36,000 km altitude assuming a diffuse Lambertian phase function. The average number of detections each night over all 3 years was 26. The percentage of this number that represented the UCT population ranged from 34% to 18%, depending on the observing strategy and the field center location. Due to the short orbital arc over which observations are made, the eccentricity of the object s orbit is extremely difficult to measure accurately. Therefore, a circular orbit was assumed when calculating the orbital elements. A comparison of the measured inclination (INC), right ascension of ascending node (RAAN), and mean motion to the quantities for CTs from the U.S. Space Surveillance Network shows acceptable errors. This analysis lends credibility to the determination of the UCT orbital distributions. Figure 1 shows the size distribution of 3,143 objects detected in the data processed for CYs 2007, 2008, and 2009. The actual peak of the absolute magnitude distribution for the functional correlated targets is 10th magnitude, whereas the peak was 11th magnitude in 2002 2003 and 10th magnitude for 2004-2006. An absolute magnitude of 10.5 corresponds to objects with average diameters of 6.3 m, assuming an albedo of 0.175 and a diffuse Lambertian phase function. This result generally agrees with the known sizes of intact satellites. The absolute magnitude distribution for the UCTs is broad, but starts to roll off near 25 cm diameter or 17.5 magnitude. This roll off in the distribution reflects the detection capability of MODEST, not the true nature of the population. The true population is believed to continue at the same slope through fainter magnitudes based on comparisons with the LEO break-up law.

Abercromby, K. J.↗
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 NTRS1974

A second-order theory of an artificial satellite under the influence of the oblateness of the earth

The general theory of the method of averaging is used to study the effect of the oblateness of the earth on the motion of an artificial satellite. The first-order harmonic J2 and the second-order harmonics J3 and J4 are included in the analysis. The first-order, periodic variations together with the second-order, averaged and secular variations are derived for all the six orbital elements. Alternative orbital elements and necessary relations are introduced to deal with circular orbits, orbits at critical inclination and equatorial orbits. Essential transformations and initialization procedures are also outlined.

Liu, J. J. F.↗
NASA NTRS1991

Orbit crowding of molecular gas at a bar-spiral arm transition zone in M83

The southwestern bar-spiral arm transition zone in M83 is been studied in CO, H-alpha, H I, red light, and the radio continuum. A massive molecular gas complex in the heart of the transition zone is composed or two principal components which have the morphology and kinematics expected from orbit crowding, where gas on highly elliptical orbits form the bar region converges with gas on more circular orbits from the spiral arm region. Three mechanisms for the origin of the orbit crowding are investigated, and it is proposed that the crowding is due primarily to density wave streaming motions caused by the bar and spiral arms. The inner CO component is partially coincident with a region of highly polarized radio continuum emission which precedes the H-alpha spiral arm by 15-25 arcsec, indicating that it lies on or just downstream from a shock front. This suggests that the bar gas approaching the transition zone is shocked and explains the ridge of dense gas seen upstream from the spiral arm.

Kenney, Jeffrey D. P.↗
NASA NTRS1986

Orbit Transfer Programs

Collection of computer programs developed that solve problem of transfer between noncoplanar circular orbits for spacecraft with chemical propulsion systems. Two basic programs given. First, referred to as "exact solution," gives complete, exact time histories of transfers. Second, or "approximate solution," program gives approximate information on transfer time and fuel cost but provides no detail of trajectory.

Breakwell, J. V.↗
NASA NTRS2016

Fast, Safe, Propellant-Efficient Spacecraft Motion Planning Under Clohessy-Wiltshire-Hill Dynamics

This paper presents a sampling-based motion planning algorithm for real-time and propellant-optimized autonomous spacecraft trajectory generation in near-circular orbits. Specifically, this paper leverages recent algorithmic advances in the field of robot motion planning to the problem of impulsively actuated, propellant- optimized rendezvous and proximity operations under the Clohessy-Wiltshire-Hill dynamics model. The approach calls upon a modified version of the FMT* algorithm to grow a set of feasible trajectories over a deterministic, low-dispersion set of sample points covering the free state space. To enforce safety, the tree is only grown over the subset of actively safe samples, from which there exists a feasible one-burn collision-avoidance maneuver that can safely circularize the spacecraft orbit along its coasting arc under a given set of potential thruster failures. Key features of the proposed algorithm include 1) theoretical guarantees in terms of trajectory safety and performance, 2) amenability to real-time implementation, and 3) generality, in the sense that a large class of constraints can be handled directly. As a result, the proposed algorithm offers the potential for widespread application, ranging from on-orbit satellite servicing to orbital debris removal and autonomous inspection missions.

spacecraft relative motio↗