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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 469 records · Page 26
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NASA NTRS1974

Autonomous orbital navigation using Kepler's equation

A simple method of determining the six elements of elliptic satellite orbits has been developed for use aboard manned and unmanned spacecraft orbiting the earth, moon, or any planet. The system requires the use of a horizon sensor or other device for determining the local vertical, a precision clock or timing device, and Apollo-type navigation equipment including an inertial measurement unit (IMU), a digital computer, and a coupling data unit. The three elements defining the in-plane motion are obtained from simultaneous measurements of central angle traversed around the planet and elapsed flight time using a linearization of Kepler's equation about a reference orbit. It is shown how Kalman filter theory may also be used to determine the in-plane orbital elements. The three elements defining the orbit orientation are obtained from position angles in celestial coordinates derived from the IMU with the spacecraft vertically oriented after alignment of the IMU to a known inertial coordinate frame.

Boltz, F. W.↗
NASA NTRS1999

Human Mars Mission: SEP Architecture Crew Taxi Propulsion Stage Study and Design and Technology for Reaction and Control System

The Mars exploration is a candidate pathway to expand human presence and useful activities in the solar system. There are several propulsion system options being considered to place the Mars payload on its inter-planetary transfer trajectory. One propulsion option is the use of Solar Electric Propulsion (SEP) to spiral out with the Mars payload from an initial Low Earth Orbit (LEO) to an elliptical High Earth Orbit (HEO). This report, presented in annotated facing page format, describes the work completed on the design of a crew taxi propulsion stage used in conjunction with the SEP. Transportation system/mission analysis topics covered in this report include sub-system analysis, trajectory profile description, mass performance and crew taxi stage sizing, stage configuration, stage cost, and Trans-Mars Injection (TMI) launch window. The high efficiency of SEP is used to provide the major part of the TMI propulsion maneuver. Orbital energy is continuously added over a period of approximately twelve months. The SEP and Mars payload follow a spiral trajectory from an initial LEO to a final elliptical HEO. A small chemical stage is then used to provide the final part of the TMI. The now unloaded SEP returns to LEO to repeat another spiral trajectory with payload to HEO. The spiral phase of the SEP's trajectory takes several months to reach HEO, thus significantly increasing the exposure time of the crew to zero-gravity. In order to minimize the long zero-gravity effects, a high thrust chemical stage delivers the crew to the SEP's HEO. The crew rendezvous with the Mars payload in HEO. After a checkout period the Mars payload with the crew is injected onto a Trans-Mars Trajectory by a small chemical stage.

Young, Archie↗
NASA NTRS1999

Human Mars Mission: SEP Architecture, Crew Taxi Propulsion Stage Study and Design and Technology for Reaction and Control System

The Mars exploration is a candidate pathway to expand human presence and useful activities in the solar system. There are several propulsion system options being considered to place the Mars payload on its interplanetary transfer trajectory. One propulsion option is the use of Solar Electric Propulsion (SEP) to spiral out with the Mars payload from an initial Low Earth Orbit (LEO) to an elliptical High Earth Orbit (HEO). This report, presented in annotated facing page format, describes the work completed on the design of a crew taxi propulsion stage used in conjunction with the SEP. Transportation system/mission analysis topics covered in this report include sub-system analysis, trajectory profile description, mass performance and crew taxi stage sizing, stage configuration, stage cost, and Trans-Mars Injection (TMI) launch window. The high efficiency of SEP is used to provide the major part of the TMI propulsion maneuver. Orbital energy is continuously added over a period of approximately twelve months. The SEP and Mars payload follow a spiral trajectory from an initial LEO to a final elliptical HEO. A small chemical stage is then used to provide the final part of the TMI. The now unloaded SEP returns to LEO to repeat another spiral trajectory with payload to HEO. The spiral phase of the SEP's trajectory takes several months to reach HEO, thus significantly increasing the exposure time of the crew to zero-gravity. In order to minimize the long zero-gravity effects, a high thrust chemical stage delivers the crew to the SEP's HEO. The crew rendezvous with the Mars payload in HEO. After a checkout period the Mars payload with the crew is injected onto a Trans-Mars Trajectory by a small chemical stage.

Young, Archie↗
NASA NTRS1972

Applications technology satellites advanced mission study

Three spacecraft configurations were designed for operation as a high powered synchronous communications satellite. Each spacecraft includes a 1 kw TWT and a 2 kw Klystron power amplifier feeding an antenna with multiple shaped beams. One of the spacecraft is designed to be boosted by a Thor-Delta launch vehicle and raised to synchronous orbit with electric propulsion. The other two are inserted into a elliptical transfer orbit with an Atlas Centaur and injected into final orbit with an apogee kick motor. Advanced technologies employed in the several configurations include tubes with multiple stage collectors radiating directly to space, multiple-contoured beam antennas, high voltage rollout solar cell arrays with integral power conditioning, electric propulsion for orbit raising and on-station attitude control and station-keeping, and liquid metal slip rings.

Gould, L. M.↗
NASA NTRS2013

Orbital Transfer Techniques for Round-Trip Mars Missions

The human exploration of Phobos and Deimos or the retrieval of a surface sample launched to low-Mars orbit presents a highly constrained orbital transfer problem. In general, the plane of the target orbit will not be accessible from the arrival or departure interplanetary trajectories with an (energetically optimal) tangential burn at periapsis. The orbital design is further complicated by the addition of a high-energy parking orbit for the relatively massive Deep Space Vehicle to reduce propellant expenditure, while the crew transfers to and from the target orbit in a smaller Space Exploration Vehicle. The proposed strategy shifts the arrival and departure maneuvers away from periapsis so that the apsidal line of the parking orbit lies in the plane of the target orbit, permitting highly efficient plane change maneuvers at apoapsis of the elliptical parking orbit. An apsidal shift during the arrival or departure maneuver is approximately five times as efficient as maneuvering while in Mars orbit, thus significantly reducing the propellant necessary to transfer between the arrival, target, and departure orbits.

human missions↗
NASA NTRS1981

Aerobraking techniques for planetary missions

Aerobraking techniques are applied to trajectory design for the VOIR (Venus Orbiting Imaging Radar) mission to enhance overall mission performance. The approach can be used for other planetary missions such as a Titan orbiter or a Mars sample return to earth. The orientation of the elliptical insertion orbit must be chosen in such a way that the combined effects of solar gravity, central-body harmonics, and drag on periapsis altitude are minimized to avoid frequent maneuvers. The selection of the base and width of a corridor for periapsis altitude depends upon the temperature and integrated heat of the aerobrake shield, the atmospheric density profile, the duration of aerobraking, the number of maneuvers, the time available between maneuvers, and uncertainties in the central-body gravity field and atmospheric density. Flight path profiles for an aerodynamically stable spacecraft and an inertially fixed spacecraft passing through the free molecular flow regime, as well as the transitional flow regime, are illustrated

Tang, C. C. H.↗
NASA NTRS2012

The Role of Cis-Lunar Space in Future Global Space Exploration

Cis-lunar space offers affordable near-term opportunities to help pave the way for future global human exploration of deep space, acting as a bridge between present missions and future deep space missions. While missions in cis-lunar space have value unto themselves, they can also play an important role in enabling and reducing risk for future human missions to the Moon, Near-Earth Asteroids (NEAs), Mars, and other deep space destinations. The Cis-Lunar Destination Team of NASA's Human Spaceflight Architecture Team (HAT) has been analyzing cis-lunar destination activities and developing notional missions (or "destination Design Reference Missions" [DRMs]) for cis-lunar locations to inform roadmap and architecture development, transportation and destination elements definition, operations, and strategic knowledge gaps. The cis-lunar domain is defined as that area of deep space under the gravitational influence of the earth-moon system. This includes a set of earth-centered orbital locations in low earth orbit (LEO), geosynchronous earth orbit (GEO), highly elliptical and high earth orbits (HEO), earth-moon libration or "Lagrange" points (E-ML1 through E-ML5, and in particular, E-ML1 and E-ML2), and low lunar orbit (LLO). To help explore this large possibility space, we developed a set of high level cis-lunar mission concepts in the form of a large mission tree, defined primarily by mission duration, pre-deployment, type of mission, and location. The mission tree has provided an overall analytical context and has helped in developing more detailed design reference missions that are then intended to inform capabilities, operations, and architectures. With the mission tree as context, we will describe two destination DRMs to LEO and GEO, based on present human space exploration architectural considerations, as well as our recent work on defining mission activities that could be conducted with an EML1 or EML2 facility, the latter of which will be an emphasis of this paper, motivated in part by recent interest expressed at the Global Exploration Roadmap Stakeholder meeting. This paper will also explore the links between this HAT Cis-Lunar Destination Team analysis and the recently released ISECG Global Exploration Roadmap and other potential international considerations, such as preventing harmful interference to radio astronomy observations in the shielded zone of the moon.

Bobskill, Marianne R.↗
NASA NTRS1976

Effects of atomic oxygen on graphite ablation

It is shown that a previously derived semiempirical equation for describing observed ablation rates of isotropic graphites cannot be applied to low-density flows containing dissociated oxygen. Experimentally determined reaction probabilities of isotropic graphites to molecular and atomic oxygen are used to calculate heat-transfer rates and stagnation-point ablation rates for typical conditions. Integrated mass losses are computed for a group of flight trajectories which start from geosynchronous orbit and enter earth's atmosphere in a skipping motion following near-elliptic decaying orbits. A comparison of the results with those obtained by the equation under question shows excellent agreement for steep trajectories, but large discrepancies for shallow trajectories. The differences are attributed to surface oxidation by atomic oxygen.

Park, C.↗
NASA NTRS1980

Targeting with fixed propellant load

The inertial upper stage (IUS) for the space shuttle operation employs solid rocket stages with fixed, propellant loadings. This means that, if for a given mission the satellite weight is less than the maximum, the IUS will deliver higher delta V's than required. One way to waste the excess capability in order to achieve the desired orbit would be to execute a nonoptimal transfer which would require higher than maximum delta V's. An algorithm is presented which defines takeoff points on the parking orbit and the injection points on the orbit for which transfer orbits require a fixed delta V sub 1 and a fixed delta V sub 2 (defined by the satellite weight). It is assumed that both the parking and the target orbit are elliptical. This allows the use of the same algorithm for guidance. The transfer orbit achieved by the erroneous delta V sub 1 is regarded as a new parking orbit and the new transfer problem is solved by assuming a delta, delta V sub 1. The maximum value of delta, delta V sub 1 is the delta V sub 1 variation and its minimum value is the one which still yields a solution by the algorithm. The delta V sub errors are regarded as orbit injection errors and compensated in the usual way.

Gedeon, G. S.↗
NASA NTRS1986

Aero-assisted orbital plane change using an elliptic drag polar

A three-impulse, aero-assisted maneuver is used to change the plane of a circular orbit. The guidance law used during the atmospheric phase is based on the repetitive use of an approximate optimal control law. The approximations include Loh's term being constant, an exponential atmosphere, and a constant-coefficient drag polar. Whereas the true optimal trajectory is flown at maximum lift-to-drag ratio, simulation results for a parabolic drag polar show the angle of attack becoming large near the end of the trajectory. Here, the approximate optimal control rule is developed for an elliptic polar which has maximum lift coefficient and higher values of the drag coefficient at high values of the lift coefficient than the parabolic polar. Simulation results show that the angle of attack still increases to high values near the end of the trajectory and that, overall, the parabolic drag polar produces better results. Finally, it is shown that Loh's term is not constant over the ascent portion of the atmospheric turn and is the probable cause of high angles of attack. However, the guidance laws developed by assuming Loh's term to be constant work well.

Hull, D. G.↗
NASA NTRS1978

New concepts for Mercury orbiter missions

The next logical step in the exploration of Mercury is an orbiter mission. A conflict exists between those in the field of planetary sciences who desire a mission with a low circular orbit, and scientists in the fields and particles disciplines, who generally prefer a highly elliptical spacecraft orbit. The thermal environment imposed by the sun and planet render the low orbit intolerable for spacecraft using previous thermal control methods. A thermal control concept and a spacecraft mission concept have been developed which resolve these problems and promise a scientifically significant mission for the mid-1980s.

French, J. R.↗
NASA NTRS1988

LLOFX earth orbit to lunar orbit delta V estimation program user and technical documentation

The LLOFX computer program calculates in-plane trajectories from an Earth-orbiting space station to Lunar orbit in such a way that the journey requires only two delta V burns (one to leave Earth circular orbit and one to circularize into Lunar orbit). The program requires the user to supply the Space Station altitude and Lunar orbit altitude (in km above the surface), and the desired time of flight for the transfer (in hours). It then determines and displays the trans-Lunar injection (TLI) delta V required to achieve the transfer, the Lunar orbit insertion (LOI) delta V required to circularize the orbit around the Moon, the actual time of flight, and whether the transfer orbit is elliptical or hyperbolic. Return information is also displayed. Finally, a plot of the transfer orbit is displayed.

Source record↗
NASA NTRS1990

Optimal aeroassisted orbital transfer with plane change using collocation and nonlinear programming

The fuel optimal control problem arising in the non-planar orbital transfer employing aeroassisted technology is addressed. The mission involves the transfer from high energy orbit (HEO) to low energy orbit (LEO) with orbital plane change. The basic strategy here is to employ a combination of propulsive maneuvers in space and aerodynamic maneuvers in the atmosphere. The basic sequence of events for the aeroassisted HEO to LEO transfer consists of three phases. In the first phase, the orbital transfer begins with a deorbit impulse at HEO which injects the vehicle into an elliptic transfer orbit with perigee inside the atmosphere. In the second phase, the vehicle is optimally controlled by lift and bank angle modulations to perform the desired orbital plane change and to satisfy heating constraints. Because of the energy loss during the turn, an impulse is required to initiate the third phase to boost the vehicle back to the desired LEO orbital altitude. The third impulse is then used to circularize the orbit at LEO. The problem is solved by a direct optimization technique which uses piecewise polynomial representation for the state and control variables and collocation to satisfy the differential equations. This technique converts the optimal control problem into a nonlinear programming problem which is solved numerically. Solutions were obtained for cases with and without heat constraints and for cases of different orbital inclination changes. The method appears to be more powerful and robust than other optimization methods. In addition, the method can handle complex dynamical constraints.

Shi, Yun. Y.↗
NASA NTRS1995

Orbiting meteoroid and debris counting experiment

The Orbiting Meteoroid and Debris Counting Experiment (OMDC) flew for approximately 90 days in a highly elliptical earth orbit onboard the Clementine Interstage Adapter (ISA) Spacecraft. This experiment obtained data on the impact flux of natural micrometeoroids and it provided limited information on the population of small mass man-made debris as a function of altitude in near earth space. The flight of the OMDC experiment on the ISA spacecraft also demonstrated that the ultra-lightweight, low-power, particle impact detector system that was used is a viable system for flights on future spacecraft to monitor the population of small mass man-made debris particles and to map the cosmic dust environment encountered on interplanetary missions. An overview of the ISA spacecraft mission, the approach to the OMDC experiment, and the data obtained by the experiment are presented.

Kinard, William H.↗
NASA NTRS2014

Lunar Flashlight: Mapping Lunar Surface Volatiles Using a Cubesat

Water ice and other volatiles may be located in the Moon's polar regions, with sufficient quantities for in situ extraction and utilization by future human and robotic missions. Evidence from orbiting spacecraft and the LCROSS impactor suggests the presence of surface and/or nearsurface volatiles, including water ice. These deposits are of interest to human exploration to understand their potential for use by astronauts. Understanding the composition, quantity, distribution, and form of water/H species and other volatiles associated with lunar cold traps is identified as a NASA Strategic Knowledge Gap (SKG) for Human Exploration. These polar volatile deposits could also reveal important information about the delivery of water to the Earth- Moon system, so are of scientific interest. The scientific exploration of the lunar polar regions was one of the key recommendations of the Planetary Science Decadal Survey. In order to address NASA's SKGs, the Advanced Exploration Systems (AES) program selected three lowcost 6-U CubeSat missions for launch as secondary payloads on the first test flight (EM1) of the Space Launch System (SLS) scheduled for 2017. The Lunar Flashlight mission was selected as one of these missions, specifically to address the SKG associated with lunar volatiles. Development of the Lunar Flashlight CubeSat concept leverages JPL's Interplanetary Nano- Spacecraft Pathfinder In Relevant Environment (INSPIRE) mission, MSFC's intimate knowledge of the Space Launch System and EM-1 mission, small business development of solar sail and electric propulsion hardware, and JPL experience with specialized miniature sensors. The goal of Lunar Flashlight is to determine the presence or absence of exposed water ice and its physical state, and map its concentration at the kilometer scale within the permanently shadowed regions of the lunar south pole. After being ejected in cislunar space by SLS, Lunar Flashlight deploys its solar panels and solar sail and maneuvers into a low-energy transfer to lunar orbit. The solar sail and attitude control system work to bring the satellite into an elliptical polar orbit spiraling down to a perilune of 30-10 km above the south pole for data collection. Lunar Flashlight uses its solar sail to shine reflected sunlight into permanently shadowed regions, measuring surface albedo with a four-filter point spectrometer at 1.1, 1.5 1.9, and 2.0 microns. Water ice will be distinguished from dry regolith from these measurements in two ways: 1) spatial variations in absolute reflectance (water ice is much brighter in the continuum channels), and 2) reflectance ratios between absorption and continuum channels. Derived reflectance and reflectance ratios will be mapped onto the lunar surface in order to distinguish the composition of the PSRs from that of the sunlit terrain. Lunar Flashlight enables a low-cost path to in-situ resource utilization (ISRU) by identifying operationally useful deposits (if there are any), which is a game-changing capability for expanded human exploration.

Cohen, B. A.↗
NASA NTRS2015

Lunar Flashlight

The Lunar Flashlight is a Jet Propulsion Laboratory project, with NASA Marshall Space Flight Center (MSFC) serving as the principal investigator and providing the solar sail propulsion system. The goal of Lunar Flashlight is to determine the presence and abundance of exposed lunar water ice within permanently shadowed regions (PSRs) at the lunar south pole, and to map its concentration at the 1-2 kilometer scale to support future exploration and use. After being ejected in cis-lunar space by the launch vehicle, Lunar Flashlight deploys solar panels and an 85-square-meter solar sail and maneuvers into a low-energy transfer to lunar orbit. The solar sail and attitude control system work to bring the satellite into an elliptical polar orbit, spiraling down over a period of 18 months to a perilune of 30-10 kilometers above the south pole for data collection. Lunar Flashlight uses its solar sail to shine reflected sunlight onto the lunar surface, measuring surface reflectance with a four-filter point spectrometer. The spectrometer measures water ice absorption features (1.5, 1.95 microns) and the continuum between them (1.1, 1.9 microns). The ratios of water ice bands to the continuum will provide a measure of the abundance of surface frost and its variability across PSRs. Water ice abundance will be correlated with other data from previous missions, such as the Lunar Reconnaissance Orbiter and Lunar Crater Observation and Sensing Satellite, to provide future human and robotic explorers with a map of potential resources. The mission is enabled by the use of an 85-square-meter solar sail being developed by MSFC.

Baker, John↗
NASA NTRS1987

On meteor stream spatial structure theory

The classical spatial representation of meteor streams is an elliptical torus with variable cross section. The position of this torus in space is determined by the mean orbit elements that may be obtained directly from observations of individual meteor stream particles when crossed by the Earth. Since the orbits of individual particles of a stream differ from each other, the distance between them on a plane normal to the mean orbit of elliptical torus forms some area, i.e., a cross section. The size and form of these cross sections change with the change of the direction among the mean orbit and are completely defined by the dispersion values of the orbit elements in a stream. An attempt was made to create an analytical method that would permit description of the spatial and time parameters of meteor streams, i.e., the form and size of their cross section, density of incident flux and their variations along the mean orbit and in time. In this case, the stream is considered as a continuous flux rather than a set of individual particles.

Andreev, G. V.↗