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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 307 records · Page 17
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Feasibility of Decentralized Linear-Quadratic-Gaussian Control of Autonomous Distributed Spacecraft

A distributed satellite formation, modeled as an arbitrary number of fully connected nodes in a network, could be controlled using a decentralized controller framework that distributes operations in parallel over the network. For such problems, a solution that minimizes data transmission requirements, in the context of linear-quadratic-Gaussian (LQG) control theory, was given by Speyer. This approach is advantageous because it is non-hierarchical, detected failures gracefully degrade system performance, fewer local computations are required than for a centralized controller, and it is optimal with respect to the standard LQG cost function. Disadvantages of the approach are the need for a fully connected communications network, the total operations performed over all the nodes are greater than for a centralized controller, and the approach is formulated for linear time-invariant systems. To investigate the feasibility of the decentralized approach to satellite formation flying, a simple centralized LQG design for a spacecraft orbit control problem is adapted to the decentralized framework. The simple design uses a fixed reference trajectory (an equatorial, Keplerian, circular orbit), and by appropriate choice of coordinates and measurements is formulated as a linear time-invariant system.

Carpenter, J. Russell↗
NASA NTRS1981

Pursuit/evasion in orbit

Maneuvers available to a spacecraft having sufficient propellant to escape an antisatellite satellite (ASAT) attack are examined. The ASAT and the evading spacecraft are regarded as being in circular orbits, and equations of motion are developed for the ASAT to commence a two-impulse maneuver sequence. The ASAT employs thrust impulses which yield a minimum-time-to-rendezvous, considering available fuel. Optimal evasion is shown to involve only in-plane maneuvers, and begins as soon as the ASAT launch information is gathered and thrust activation can be initiated. A closest approach, along with a maximum evasion by the target spacecraft, is calculated to be 14,400 ft. Further research to account for ASATs in parking orbit and for generalization of a continuous control-modeled differential game is indicated.

Kelley, H. J.↗
NASA NTRS1983

Synthetic aperture radar imaging from an inclined geosynchronous orbit

Images of earth can be produced with an assumed synthetic aperture radar (SAR) on a satellite platform undergoing a nutating relative motion from geosynchronous altitude. From a 50 deg inclined circular orbit, the contiguous United States can be imaged in about 3 h of segmented operation at 100-m resolution with 4-azimuth-look averaging. The 2450-MHz transmitter radiates 1312 W of average power from a steerable 15-m-diam antenna. The SAR can image daily an area bounded longitudinally and latitudinally.

Tomiyasu, K.↗
DOE OSTI2020

Revisiting the Architecture of the KOI-89 System

While high stellar obliquities observed in exoplanetary systems may be attributed to processes that tilt the planetary orbits, it is also possible that they reflect misalignments between protoplanetary disks and stellar spins. This latter hypothesis predicts the presence of coplanar multiplanetary systems misaligned with their central stars. Here we reevaluate the evidence of such an architecture that has been claimed for the KOI-89 system. An early-type star, KOI-89 has one validated transiting planet, KOI-89.01/Kepler-462b (period 84.7 days, radius 3.0 R {sub ⊕}), and one transiting planet candidate, KOI-89.02 (period 207.6 days, radius 4.0 R {sub ⊕}), where the latter exhibits transit timing variations (TTVs). A previous modeling of the stellar gravity-darkening effect in the transit light curves inferred a high stellar obliquity of ≈70°. We perform photodynamical modeling of the Kepler transit light curves and use the resulting constraints on the orbital configuration and transit times to update the gravity-darkened transit model. As a result, we find no firm evidence for the gravity-darkening effect in the transit shapes and conclude that stellar obliquity is not constrained by the data. Given the evidence for low orbital eccentricities from the dynamical analysis, the system architecture can thus be consistent with many other multitransiting systems with flat, near-circular orbits aligned with the stellar spin. We find that the TTVs imparted on its neighbor imply that KOI-89.01 has a mass ≳20 M {sub ⊕}. This would render it one of the densest known sub-Neptunes, mostly composed of a solid core. Lower masses are possible if the TTVs are instead due to an unseen third planet.

79 ASTRONOMY AND ASTROPHYSICS↗
NASA NTRS1996

Analysis of Apollo 8 Photography and Visual Observations

Apollo 8 was launched from Cape Kennedy, Fla., at 7 :50 a.m., e.s.t., on December 21, 1968. Two hours 50 minutes later, translunar injection was performed; and astronauts Col. Frank Borman, the commander; Capt. James A. Lovell, Jr., the command module pilot; and Maj. (now Lt. Col.) William A. Anders, the lunar module pilot, were on their way to the Moon. The spacecraft was placed in an elliptical lunar orbit at 69 hours 8 minutes after liftoff. After flying two elliptical orbits of 168.5 by 60 nautical miles with an inclination of 12° to the Equator, the spacecraft was placed in a nearly circular orbit of 59.7 by 60.7 nautical miles, in which it remained for eight orbits. At 89 hours 19 minutes, trans earth injection was performed from behind the Moon. A nearly flawless mission was completed on the morning of December 27 when splashdown occurred in the Pacific Ocean after a total elapsed time of 147 hours. Lt. Gen. Sam C. Phillips, the Director of the Apollo Program, announced that such a mission was being considered at a press conference on August 19, 1968. Formal announcement that NASA was preparing Apollo 8 for an orbital flight around the Moon was released to the press on November 12, 1968. The primary purpose of this mission was to further progress toward the goal of landing men on the Moon by gaining operational experience and testing the Apolio systems. However, a great effort was also made to accomplish worthwhile scientific tasks with photography and visual observations by the astronauts. In planning the scientific tasks to be attempted on this mission, it was obvious that one of the prime tasks should be photography of the lunar surface. Such photography would furnish valuable information on the following : 1. Approach topography and landmarks for the early Apollo landings 2. The scientific merit and the roughness of areas for possible follow-on Apollo landings 3. The broad structure and characteristics of the lunar surface During the orbital part of the mission, a major portion of the lunar far side would be in sunlight. Although almost all of the far side of the Moon has been photographed by the automated Lunar Orbiter spacecraft, the photography generally was made with the spacecraft relatively far from the Moon, limiting the Lunar Orbiter photographs to an average resolution of approximately 100 meters. Thus, Apollo photographs of the far side would have much better resolution than existing pictures. Finally, it was recognized that contamination, both as it relates to window fogging (which did occur) and to contamination clouds around t he spacecraft, should be studied for both scientific and operational interests.

Richard J. Allenby↗
NASA NTRS1978

Space Shuttle payloads and data handling accommodations

The Space Transportation System, which includes the Space Shuttle, will reduce the costs of earth orbital operations while improving operational capabilities and flexibility. The Shuttle can deliver payloads up to 65,000 pounds to a 150-nmi circular orbit at a substantially lower operational cost per flight than the Titan III-C system, which has a payload delivery capability of only 30,000 pounds. In addition, the Space Shuttle can return to earth with up to 32,000 pounds of payload, a capability not provided by expendable launch vehicles. Typical payload missions are considered along with aspects of satellite placement and recovery, typical free-flying payloads, attached scientific payloads, and the typical nature of payload data.

Teasdale, W. E.↗
NASA NTRS1987

Nuclear reactor power for an electrically powered orbital transfer vehicle

To help determine the systems requirements for a 300-kWe space nuclear reactor power system, a mission and spacecraft have been examined which utilize electric propulsion and this nuclear reactor power for multiple transfers of cargo between low earth orbit (LEO) and geosynchronous earth orbit (GEO). A propulsion system employing ion thrusters and xenon propellant was selected. Propellant and thrusters are replaced after each sortie to GEO. The mass of the Orbital Transfer Vehicle (OTV), empty and dry, is 11,000 kg; nominal propellant load is 5000 kg. The OTV operates between a circular orbit at 925 km altitude, 28.5 deg inclination, and GEO. Cargo is brought to the OTV by Shuttle and an Orbital Maneuvering Vehicle (OMV); the OTV then takes it to GEO. The OTV can also bring cargo back from GEO, for transfer by OMV to the Shuttle. OTV propellant is resupplied and the ion thrusters are replaced by the OMV before each trip to GEO. At the end of mission life, the OTV's electric propulsion is used to place it in a heliocentric orbit so that the reactor will not return to earth. The nominal cargo capability to GEO is 6000 kg with a transit time of 120 days; 1350 kg can be transferred in 90 days, and 14,300 kg in 240 days. These capabilities can be considerably increased by using separate Shuttle launches to bring up propellant and cargo, or by changing to mercury propellant.

Jaffe, L.↗
NASA NTRS1988

Nuclear-electric reusable orbital transfer vehicle

To help determine the systems requirements for a 300-kWe space nuclear reactor power system, a mission and spacecraft have been examined that utilize electric propulsion supported by the nuclear reactor's power for multiple transfers of cargo between low earth orbit (LEO) and geosynchronous earth orbit (GEO). A propulsion system employing ion thrusters and xenon propellant was selected. Propellant and thrusters are replaced after each sortie to GEO. The mass of the orbital transfer vehicle (OTV), empty and dry, is 11,000 kg; nominal propellant load is 5000 kg. The OTV operates between a circular orbit at 925-km altitude, 28.5-deg inclination, and GEO. Cargo is brought to the OTV by Shuttle and an orbital maneuvering vehicle (OMV); the OTV then takes it to GEO. The OTV can also bring cargo back from GEO for transfer by OMV to the Shuttle. OTV propellant is resupplied, and the ion thrusters are replaced, by the OMV before each sortie to GEO. At the end of mission life, the OTV's electric propulsion is used to place it in a heliocentric orbit so that the reactor will not return to earth. The nominal cargo capability to GEO is 6000 kg, with a transit time of 120 days; 1350 kg can be transferred in 90 days, and 14,300 kg in 240 days. These capabilities can be considerably increased by using separate Shuttle launches to bring up propellant and cargo or by changing to mercury propellant.

Jaffe, Leonard D.↗
NASA NTRS1994

Orbital maneuvering and reaction control systems

Orbital maneuvering systems (OMS) and reaction control systems (RCS) provide capabilities to spacecraft that include orbit circularization, rendezvous maneuvers, attitude control, and re-entry delta velocity. The mission and vehicle requirements can place severe demands on the orbital maneuvering and reaction control systems. In order to perform proper trade studies and to design these systems, the mission and vehicle configuration must be well defined. In the absence of a clearly defined mission and vehicle configuration, the research and development of basic technologies must support future design efforts by providing a range of options and data from which to select. This paper describes the key OMS and RCS requirements and technology.

Hurlbert, Eric↗
NASA NTRS2011

Understanding Geospace on a Grand Scale: The Global Ionoshphere/Thermosphere Constellation

We present the concept of a constellation of polar orbiting satellites equally spaced in longitude (local time) to systematically sample both the neutral and ionzed gas components of the Earth in circular orbits near 350 km, including their density, temperature, and velocities. The instrumentation would include techniques to measure the height of the ionospheric "F-peak" and its variations along the orbit. The number of satellites (l2? 24? 48?) and their configuration would be determined from modeling analysis and expected geophysical phenomena, including their drivers and characteristic time scales. Together with imaging data from separate satellites, the array of satellites with in situ probes would be expected to provide a new picture of (1) high latitude electrodynamics and atmospheric processes and associated coupling with magnetospheric mass and momentum input, (2) the response of the global ionosphere and thermosphere to magnetic storms, and (3) global neutral wind circulation patterns, neutral density structure, tides, planetary waves, and gravity waves. The comprehensive measurements gathered by the IT-Constellation envisioned here would provide a major leap forward in each of these areas, addressing global physical processes and providing fundamental, new knowledge of Geospace. In particular, by its very nature, the constellation addresses "system science", revealing how the ionosphere-thermosphere connects globally to the magnetosphere above and the troposphere below. We present this concept as the next logical step in observing the "whole" space environment using in situ probes in conjunction with imagers. We invite modelers to not only comment on this concept but also to become actively engaged in helping to define it.

Pfaff, Robert↗
NASA NTRS2000

An Overview of the Topography of Mars from the Mars Orbiter Laser Altimeter (MOLA)

The Mars Global Surveyor (MGS) spacecraft has now completed more than half of its one-Mars-year mission to globally map Mars. During the MGS elliptical and circular orbit mapping phases, the Mars Orbiter Laser Altimeter (MOLA), an instrument on the MGS payload, has collected over 300 million precise elevation measurements. MOLA measures the range from the MGS spacecraft to the Martian surface and to atmospheric reflections. Range is converted to topography through knowledge of the MGS spacecraft orbit. Ranges from MOLA have resulted in a precise global topographic map of Mars. The instrument has also provided measurements of the width of the backscattered optical pulse and of the 1064 nm reflectivity of the Martian surface and atmosphere. The range resolution of the MOLA instrument is 37.5 cm and the along-track resolution of MOLA ground shots is approx. 300 m; the across-track spacing depends on latitude and time in the mapping orbit. The best current topographic grid has a spatial resolution of approx. 1/16 deg and vertical accuracy of approx. one meter. Additional information is contained in the original extended abstract.

Smith, David E.↗
NASA NTRS1989

Electromagnetic launchers for space applications

An electromagnetic launcher (EML) was designed for NASA-Langley to boost large models to hypervelocity for flight evaluation. Two different concepts were developed using railgun and coilgun principles. A coilgun was designed to accelerate a 14-kg mass to 6 km/s and, by adding additional equipment, to accelerate a 10-kg mass to 11 km/s. The railgun system was designed to accelerate only 14 kg to 6 km/s. Of significance in this development is the opportunity to use the launcher for aeroballistic research of the upper atmosphere, eventually placing packages in low earth orbit using a small rocket. The authors describe the railgun and coilgun launch designs and suggest a reconfiguration for placement of 150-kg parcels into low earth orbit for aeroballistic studies and possible space lab support. Each design is detailed along with the performance adjustments which would be required for circular orbit payload placement.

Schroeder, J. M.↗
NASA NTRS2003

Simulated Fuel Usage for Drag Compensated Spacecraft in Low Earth Orbit

Atmospheric drag causes the greatest uncertainty in the spacecraft flight dynamics equation for spacecraft in Low Earth Orbit (LEO). The continuously varying atmospheric density levels require increased spacecraft tracking to accurately predict spacecraft location. In addition, periodic propulsive maneuvers must be designed and performed to counteract the effects of drag on the spacecraft orbit. If atmospheric drag effects can be continuously and autonomously counteracted through the use of a drag-free control system, they will essentially be eliminated from the spacecraft flight dynamics equation. The main perturbations on the spacecraft will then be those due to the Earth s gravitational field, which are easily predicted. Although theoretically beneficial from a flight dynamics perspective, the costs associated with drag-free control must be determined and weighed against those benefits before the feasibility of drag-free control of spacecraft in LEO can be determined. Through the use of a MATLAEV Satellite Tool Kit simulation, this paper attempts to quantify one such cost associated with drag-free control: the amount of fuel needed for continuous drag compensation as measured by cumulative orbital velocity changes, or AV. This AV cost can then be compared to the fuel required for more traditional, periodic orbit raising maneuvers of every two to four weeks. The simulation considers various sized spacecraft, as measured by the spacecraft ballistic coefficient, in circular orbits of various altitudes and inclinations. The time between orbit raising maneuvers is allowed to decrease until it approaches near continuous. In this manner, the trend in AV cost can be seen as drag compensation moves from periodic to near continuous.

Fleck, Melissa E.↗
NASA NTRS2005

Propellantless AOCS Design for a 160-m, 450-kg Sailcraft of the Solar Polar Imager Mission

An attitude and orbit control system (AOCS) is developed for a 160-m, 450-kg solar sail spacecraft of the Solar Polar Imager (SPI) mission. The SPI mission is one of several Sun- Earth Connections solar sail roadmap missions currently envisioned by NASA. A reference SPI sailcraft consists of a 160-m, 150-kg square solar sail, a 250-kg spacecraft bus, and 50-kg science payloads, The 160-m reference sailcraft has a nominal solar thrust force of 160 mN (at 1 AU), an uncertain center-of-mass/center-of-pressure offset of +/- 0.4 m, and a characteristic acceleration of 0.35 mm/sq s. The solar sail is to be deployed after being placed into an earth escaping orbit by a conventional launch vehicle such as a Delta 11. The SPI sailcraft first spirals inwards from 1 AU to a heliocentric circular orbit at 0.48 AU, followed by a cranking orbit phase to achieve a science mission orbit at a 75-deg inclination, over a total sailing time of 6.6 yr. The solar sail will be jettisoned after achieving the science mission orbit. This paper focuses on the solar sailing phase of the SPI mission, with emphasis on the design of a reference AOCS consisting of a propellantless primary ACS and a microthruster-based secondary (optional) ACS. The primary ACS employs trim control masses running along mast lanyards for pitch/yaw control together with roll stabilizer bars at the mast tips for quadrant tilt (roll) control. The robustness and effectiveness of such a propellantless primary ACS would be enhanced by the secondary ACS which employs tip-mounted, lightweight pulsed plasma thrusters (PPTs). The microPPT-based ACS is mainly intended for attitude recovery maneuvers from off-nominal conditions. A relatively fast, 70-deg pitch reorientation within 3 hrs every half orbit during the orbit cranking phase is shown to be feasible, with the primary ACS, for possible solar observations even during the 5-yr cranking orbit phase.

Wie, Bong↗
NASA NTRS1976

Requirements and capabilities for planetary missions. Volume 2: Mars polar orbiter penetrator 1981

The Mars Polar Orbiter/Penetrator 1981 mission, intended to investigate the manner in which Mars has evolved, and which surveys its geochemistry, performs climatological investigations, and attempts to determine the planet's gravitational field, was described. The spacecraft, modified from the Viking Orbiter design, carries a new remote-sensing payload and six penetrators. The penetrators are released from a 2.46-h, 1000-km sun synchronous circular orbit and interrogated daily throughout the 2-year orbital mission. X-band telemetry is used to increase data return.

Ball, G. G.↗
NASA NTRS1990

Ionizing radiation exposure of LDEF

The Long Duration Exposure Facility (LDEF) was launched into orbit by the Space Shuttle 'Challenger' mission 41C on 6 April 1984 and was deployed on 8 April 1984. The original altitude of the circular orbit was 258.5 nautical miles (479 km) with the orbital inclination being 28.5 degrees. The 21,500 lb NASA Langley Research Center satellite, having dimensions of some 30x14 ft was one of the largest payloads ever deployed by the Space Shuttle. LDEF carried 57 major experiments and remained in orbit five years and nine months (completing 32,422 orbits). It was retrieved by the Shuttle 'Columbia' on January 11, 1990. By that time, the LDEF orbit had decayed to the altitude of 175 nm (324 km). The experiments were mounted around the periphery of the LDEF on 86 trays and involved the representation of more than 200 investigators, 33 private companies, 21 universities, seven NASA centers, nine Department of Defense laboratories and eight foreign countries. The experiments covered a wide range of disciplines including basic science, electronics, optics, materials, structures, power and propulsion. The data contained in the LDEF mission represents an invaluable asset and one which is not likely to be duplicated in the foreseeable future. The data and the subsequent knowledge which will evolve from the analysis of the LDEF experiments will have a very important bearing on the design and construction of the Space Station Freedom and indeed on other long-term, near-earth orbital space missions. A list of the LDEF experiments according to experiment category and sponsor is given, as well as a list of experiments containing radiation detectors on LDEF including the LDEF experiment number, the title of the experiment, the principal investigator, and the type of radiation detectors carried by the specific experiment.

Benton, E. V.↗
NASA NTRS2003

Mars orbits with daily repeating ground traces

This paper derives orbits at Mars with ground traces that repeat at the same times every solar day (sol). A relay orbiter in such an orbit would pass over insitu probes at the same times every sol, ensuring consistent coverage and simplifying mission design and operations. 42 orbits in five classes are characteried: 14 cicular equatorial prograde orbits; 14 circular equatorial retrograde orbits; 11 circular sun synchrounous orbits; 2 eccentroc equatorial orbits; 1 eccentric critcally inclined orbit. the paper reports on the performance of a relay orbiter in some of the orbits.

Mars↗