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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 811 records · Page 45
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Cardiac and Vascular Responses to Thigh Cuffs and Respiratory Maneuvers on Crewmembers of the International Space Station

The transition to microgravity eliminates the hydrostatic gradients in the vascular system. The resulting fluid redistribution commonly manifests as facial edema, engorgement of the external neck veins, and a decrease in leg diameter. This experiment examined the responses to modified Valsalva and Mueller maneuvers measured by cardiac and vascular ultrasound (ECHO) in a baseline steady state and during preload reduction introduced with thigh occlusion cuffs used as a counter-measure device (Braslet cuffs) measured by cardiac and vascular ultrasound examinations. Methods: Nine International Space Station crewmember subjects (Expeditions 16 - 20) were examined in 15 experiment sessions 101 +/- 46.days after launch (mean +/- SD; 33 - 185). Twenty Seven cardiac and vascular parameters were obtained with/without respiratory maneuvers before and after tightening of the Braslet cuffs. Results: Non-physicians performed diagnostic-quality cardiac and vascular ultrasound examinations using remote guidance. Three of 27 combinations of maneuvers and Braslet or Braslet alone were identified as being significant changed when compared to baseline. Eleven of 81 differences between combinations of Mueller, Valsalva or baseline were significant and related to cardiac preload reduction or increase in lower extremity venous volume. Conclusions: Acute application of Braslet occlusion cuffs causes lower extremity fluid sequestration and exerts commensurate measurable effects on cardiac performance in microgravity. Ultrasound techniques to measure the hemodynamic effects of thigh cuffs in combination with respiratory maneuvers may serve as an invaluable tool in determining the volume status of the cardiac patient at the 'microgravity bedside'.

Hamilton, Douglas↗
NASA NTRS2012

Integrated CFD and Controls Analysis Interface for High Accuracy Liquid Propellant Slosh Predictions

Coupling computational fluid dynamics (CFD) with a controls analysis tool elegantly allows for high accuracy predictions of the interaction between sloshing liquid propellants and the control system of a launch vehicle. Instead of relying on mechanical analogs which are n0t va lid during all stages of flight, this method allows for a direct link between the vehicle dynamic environments calculated by the solver in the controls analysis tool to the fluid now equations solved by the CFD code. This paper describes such a coupling methodology, presents the results of a series of test cases, and compares said results against equivalent results from extensively validated tools. The coupling methodology, described herein, has proven to be highly accurate in a variety of different cases.

Marsell, Brandon↗
NASA NTRS2010

Lunar Reconnaissance Orbiter (LRO): Observations for Lunar Exploration and Science

The Lunar Reconnaissance Orbiter (LRO) was implemented to facilitate scientific and engineering-driven mapping of the lunar surface at new spatial scales and with new remote sensing methods, identify safe landing sites, search for in situ resources, and measure the space radiation environment. After its successful launch on June 18,2009, the LRO spacecraft and instruments were activated and calibrated in an eccentric polar lunar orbit until September 15, when LRO was moved to a circular polar orbit with a mean altitude of 50 km. LRO will operate for at least one year to support the goals of NASA's Exploration Systems Mission Directorate (ESMD), and for at least two years of extended operations for additional lunar science measurements supported by NASA's Science Mission Directorate (SMD). LRO carries six instruments with associated science and exploration investigations, and a telecommunications/radar technology demonstration. The LRO instruments are: Cosmic Ray Telescope for the Effects of Radiation (CRaTER), Diviner Lunar Radiometer Experiment (DLRE), Lyman-Alpha Mapping Project (LAMP), Lunar Exploration Neutron Detector (LEND), Lunar Orbiter Laser Altimeter (LOLA), and Lunar Reconnaissance Orbiter Camera (LROC). The technology demonstration is a compact, dual-frequency, hybrid polarity synthetic aperture radar instrument (Mini-RF). LRO observations also support the Lunar Crater Observation and Sensing Satellite (LCROSS), the lunar impact mission that was co-manifested with LRO on the Atlas V (401) launch vehicle. This paper describes the LRO objectives and measurements that support exploration of the Moon and that address the science objectives outlined by the National Academy of Science's report on the Scientific Context for Exploration of the Moon (SCEM). We also describe data accessibility by the science and exploration community.

Vondrak, Richard↗
NASA NTRS2021

Development of Spaceborne SoOp Reflectometry Model for Complex Terrains

Following the launch of multiple global navigation satel-lite system (GNSS) reflectometry (GNSS-R) missions, theSignals of Opportunity (SoOp) method has proven to be apowerful tool for geophysical parameter retrieval for land ap-plications such as soil moisture. Having demonstrated thefeasibility of the SoOp techniques at P- and S-band, the devel-opment of SoOp measurements beyond the GNSS frequencyregime is highly anticipated. The SoOp Coherent Bistatic(SCoBi) model and simulator, developed in 2017 and open-sourced in 2018, has been made available to provide multi-frequency, fully polarimetric SoOp simulations for ground-based applications through the joint use of analytical wavetheory and distorted Borne approximation to evaluate landcontributions from multilayer dielectric profiles composed ofsoil moisture, vegetation, and surface roughness effects. Thispaper describes the advancement of SCoBi from a ground-and airborne-based model to a spaceborne model. This ex-tension allows for fully polarimetric, complex delay-Dopplermap (DDM) simulations through evaluation of the coherentsuperposition of electric fields emerging from a grid of ori-ented facets. The model generates a grid of facets by de-termining the geometry of contributing elements from digitalelevation models, with each element providing its contribu-tion under a flat-earth assumption. This module will enablethe analysis of fully polarimetric scattering from frequenciesavailable across the ultra-high frequency (UHF) regime.

D R Boyd↗
NASA NTRS2021

Case Studies in Verifying Spacecraft Autonomy

Spacecraft, operating with a time-to-effect faster than human command response, or with absent or delayed communication, have depended on autonomy. It is likely that this reliance will continue, and increase further in future missions. Spacecraft autonomy is used when operator intervention is not available or feasible, or when dynamic aspects of a spacecraft situation cannot be predicted in advance. This paper provides case studies of several successfully verified autonomous software systems across multiple spacecraft, flight phases (launch, transit, entry and landing, science utilization), and software classifications, and can serve as examples to future missions in methods of successfully assuring spacecraft autonomy.

Johnson, Stephen B.↗
NASA NTRSDate not supplied

The Transition of RGB Imagery Applied to Fog and Low Clouds from NASA Capabilities within the GOES Proving Ground to the Present Era

A wide range of RGB Imagery products are now available from the GOES ABI instrument, but less than a decade earlier, these false-color imagery products were only seen over Europe and Africa within the suite of the EUMETSAT SEVIRI geostationary instrument. The issue of fog and low clouds as applied to aviation and ground transportation hazards has been a long-standing challenge where satellite imagery provides notable value to visually inspect large spatial areas vs relying on in situ point observations. A legacy channel difference product was the standard satellite imagery tool, but the Nighttime Microphysics RGB would soon be introduced with GOES-R and -S where users had no interpretation experience. As part of the GOES Proving Ground efforts, the SPoRT program utilized existing NASA LEO satellites to demonstrate the future ABI capabilities. Operational users were able to test and adopt the RGBs from NASA LEO instruments prior to the GOES-R launch in order to be ready for the wealth of new channels available from ABI. The transition experience and methods from these proxy RGBs to the start of the GOES-R ABI era to the present day provides insights to the value of these new qualitative products in operations and to how the transition method of such products can be successfully implemented as well as future needs in the RGB Imagery area.

Kevin Fuell↗
NASA NTRS2012

Method to Remove Particulate Matter from Dusty Gases at Low Pressures

Future human exploration of Mars will rely on local Martian resources to reduce the mass, cost, and risk of space exploration launched from Earth. NASA's In Situ Resource Utilization (ISRU) Project seeks to produce mission consumables from local Martian resources, such as atmospheric gas. The Martian atmosphere, however, contains dust particles in the 2-to-10 -micrometer range. These dust particles must be removed before the Martian atmospheric gas can be processed. The low pressure of the Martian atmosphere, at 5 to 10 mbars, prevents the development of large voltages required for a standard electrostatic precipitator. If the voltage is increased too much, the corona transitions into a glow/streamer discharge unsuitable for the operation of a precipitator. If the voltage is not large enough, the dust particles are not sufficiently charged and the field is not strong enough to drive the particles to the collector. A method using electrostatic fields has been developed to collect dust from gaseous environments at low pressures, specifically carbon dioxide at pressures around 5 to 10 mbars. This method, commonly known as electrostatic precipitation, is a mature technology in air at one atmosphere. In this case, the high voltages required for the method to work can easily be achieved. However, in carbon dioxide at low pressures, such as those found on Mars, large voltages are not possible. The innovation reported here consists of two concentric cylindrical electrodes set at specific potential difference that generate an electric field that produces a corona capable of imparting an electrostatic charge to the incoming dust particles. The strength of the field is carefully balanced so as to produce a stable charging corona at 5 to 10 mbars, and is also capable of imparting a force to the particles that drives them to the collecting electrode. There are only two possible ways that dust can be removed from Martian atmospheric gas intakes: with this electrostatic precipitator design, and with the use of filters. However, filters require upstream compression of the gas to be treated because the atmospheric pressure on Mars is too close to vacuum to use a vacuum pump downstream to the filter to draw the gas through the filter. The electrostatic precipitator is the best and more efficient solution for this environment. No other precipitator designs have been developed for the environment of Mars due to the challenges of the low atmospheric pressure. Dust particles are charged using corona generation around the high-voltage discharge electrode, which ionizes gas molecules. Since the atmospheric gas intakes for the ISRU processing chambers will likely be cylindrical, cylindrical precipitator geometry was chosen. The electrostatic precipitator design presented here removes simulated Martian dust particles in the required range in a simulated Martian atmospheric environment. The current-voltage (I-V) characteristic curves taken for the nine precipitator configurations at 9 mbars of pressure showed that a cylindrical collecting electrode 7.0 cm in diameter with a concentric positive high voltage electrode 100 m thick provides the best range of voltage and charging corona current. This precipitator design is effective for the size of the dust particles expected in the Martian atmosphere. Mass determination, as well as microscopic images and particle size distributions of dust collected on a silicon wafer placed directly below the precipitator with the field on and off, showed excellent initial results.

Calle, Carlos↗
NASA NTRS2007

Proceedings of the 20th International Symposium on Space Flight Dynamics

Topics include: Measuring Image Navigation and Registration Performance at the 3-Sigma Level Using Platinum Quality Landmarks; Flight Dynamics Performances of the MetOp A Satellite during the First Months of Operations; Visual Navigation - SARE Mission; Determining a Method of Enabling and Disabling the Integral Torque in the SDO Science and Inertial Mode Controllers; Guaranteeing Pointing Performance of the SDO Sun-Pointing Controllers in Light of Nonlinear Effects; SDO Delta H Mode Design and Analysis; Observing Mode Attitude Controller for the Lunar Reconnaissance Orbiter; Broken-Plane Maneuver Applications for Earth to Mars Trajectories; ExoMars Mission Analysis and Design - Launch, Cruise and Arrival Analyses; Mars Reconnaissance Orbiter Aerobraking Daily Operations and Collision Avoidance; Mars Reconnaissance Orbiter Interplanetary Cruise Navigation; Motion Parameters Determination of the SC and Phobos in the Project Phobos-Grunt; GRAS NRT Precise Orbit Determination: Operational Experience; Orbit Determination of LEO Satellites for a Single Pass through a Radar: Comparison of Methods; Orbit Determination System for Low Earth Orbit Satellites; Precise Orbit Determination for ALOS; Anti-Collision Function Design and Performances of the CNES Formation Flying Experiment on the PRISMA Mission; CNES Approaching Guidance Experiment within FFIORD; Maneuver Recovery Analysis for the Magnetospheric Multiscale Mission; SIMBOL-X: A Formation Flying Mission on HEO for Exploring the Universe; Spaceborne Autonomous and Ground Based Relative Orbit Control for the TerraSAR-X/TanDEM-X Formation; First In-Orbit Experience of TerraSAR-X Flight Dynamics Operations; Automated Target Planning for FUSE Using the SOVA Algorithm; Space Technology 5 Post-Launch Ground Attitude Estimation Experience; Standardizing Navigation Data: A Status Update; and A Study into the Method of Precise Orbit Determination of a HEO Orbiter by GPS and Accelerometer.

Woodard, Mark↗
NASA NTRS2013

Reduction of Martian Sample Return Mission Launch Mass with Solar Sail Propulsion

Solar sails have the potential to provide mass and cost savings for spacecraft traveling within the innter solar system. Companies like L'Garde have demonstrated sail manufacturability and various i-space development methods. The purpose of this study was to evaluate a current Mars sample return architecture and to determine how cost and mass would be reduced by incorporating a solar sail propulsion system. The team validated the design proposed by L'Garde, and scaled the design based on a trajectory analysis. Using the solar sail design reduced the required mass, eliminating one of the three launches required in the original architecture.

Russell, Tiffany E.↗
NASA NTRS2013

Reduction of Martian Sample Return Mission Launch Mass with Solar Sail Propulsion

Solar sails have the potential to provide mass and cost savings for spacecraft traveling within the inner solar system. Companies like L'Garde have demonstrated sail manufacturability and various in-space deployment methods. The purpose of this study was to evaluate a current Mars sample return architecture and to determine how cost and mass would be reduced by incorporating a solar sail propulsion system. The team validated the design proposed by L'Garde, and scaled the design based on a trajectory analysis. Using the solar sail design reduced the required mass, eliminating one of the three launches required in the original architecture.

Russell, Tiffany E.↗
NASA NTRS2013

ISRU Propellant Selection for Space Exploration Vehicles

Chemical propulsion remains the only viable solution as technically matured technology for the near term human space transportation to Lunar and Mars. Current mode of space travel requires us to "take everything we will need", including propellant for the return trip. Forcing the mission designers to carry propellant for the return trip limits payload mass available for mission operations and results in a large and costly (and often unaffordable) design. Producing propellant via In-Situ Resource Utilization (ISRU) will enable missions with chemical propulsion by the "refueling" of return-trip propellant. It will reduce vehicle propellant mass carrying requirement by over 50%. This mass reduction can translates into increased payload to enhance greater mission capability, reduces vehicle size, weight and cost. It will also reduce size of launch vehicle fairing size as well as number of launches for a given space mission and enables exploration missions with existing chemical propulsion. Mars remains the ultimate destination for Human Space Exploration within the Solar System. The Mars atmospheric consist of 95% carbon dioxide (CO2) and the presence of Ice (water) was detected on Mars surfaces. This presents a basic chemical building block for the ISRU propellant manufacturing. However, the rationale for the right propellant to produce via ISRU appears to be limited to the perception of "what we can produce" as oppose to "what is the right propellant". Methane (CH4) is often quoted as a logical choice for Mars ISRU propellant, however; it is believed that there are better alternatives available that can result in a better space transportation architecture. A system analysis is needed to determine on what is the right propellant choice for the exploration vehicle. This paper examines the propellant selection for production via ISRU method on Mars surfaces. It will examine propellant trades for the exploration vehicle with resulting impact on vehicle performance, size, and on launch vehicles. It will investigate propellant manufacturing techniques that will be applicable on Mars surfaces and address related issues on storage, transfer, and safety. Finally, it will also address the operability issues associated with the impact of propellant selection on ground processing and launch vehicle integration.

Chen, Timothy T.↗
NASA NTRSDate not supplied

An Interim Set of TNT Curves for LOX/LNG Explosions

Space launch companies are actively developing, or in some cases have already developed, new vehicles that use large quantities of liquid oxygen (LOX) and liquid natural gas (LNG) propellants. This propellant mixture currently lacks sound/verified explosive safety standards for use in explosive siting and flight safety analysis. In the U.S., multiple government and commercial organizations have conducted limited testing, or are preparing to conduct tests, intended to provide a sound technical basis for LOX/LNG explosive standards applicable to space launch vehicle ground and flight operations. The limited testing and analysis conducted so far indicates the potential to produce extremely energetic explosions due to the miscibility of LOX/LNG (methane), a unique feature relative to previously used propellant mixtures such as hydrogen/oxygen or kerosene/oxygen. MOX is a homogenous mixture of methane and oxygen that is possible because of methane’s 100% miscibility in LOX. Preliminary testing by N2L, Inc. described in this paper shows that it is possible to form MOX with a variety of mixing methods, and that MOX is a sensitive high explosive capable of producing overpressures greater than comparable masses of C-4. The limited amount and quality of large-scale LOX/LNG explosive test data, and the potential formation of high-explosive MOX, create significant unknowns in the determination of hazard areas (cleared of the public) for ground operations (such as a static fire test) and flights of launch vehicles with LOX/LNG propellant. This paper will review the methods used by NASA to develop and validate the LOX/LH2 blast model using large-scale explosive test programs such as Project PYRO, the Hydrogen-Oxygen Vertical Impact (HOVI), and the Large-Scale Hydrogen-Oxygen Explosion (LSHOE) tests. This paper will describe how the same process was used to prepare an interim LOX/LNG blast model that applies to various LV failure scenarios and conservatively accounts for potential MOX formation. This paper also summarizes past and future testing and modeling efforts funded by a consortium of NASA, the US Federal Aviation Administration (FAA), and the US Space Force (USSF). These test programs should be concluded within the next 3 years and are intended to provide empirical data for model verification and validation.

Liquid Oxygen↗
NASA NTRSDate not supplied

An Interim Set of TNT Curves for LOX/LNG Explosions

Space launch companies are actively developing, or in some cases have already developed, new vehicles that use large quantities of liquid oxygen (LOX) and liquid natural gas (LNG) propellants. This propellant mixture currently lacks sound/verified LOX/LNG explosive safety standards for use in explosive siting and flight safety analysis. In the U.S., multiple government and commercial organizations have conducted limited testing, or are preparing to conduct tests, intended to provide a sound technical basis for explosive standards applicable to space launch vehicle ground and flight operations. The limited testing and analysis conducted so far indicates the potential to produce extremely energetic explosions due to the miscibility of LOX/LNG (methane), a unique feature relative to previously used propellant mixtures such as hydrogen/oxygen or kerosene/oxygen. MOX is a homogenous mixture of methane and oxygen that is possible because of methane’s 100% miscibility in LOX. Preliminary testing by N2L, Inc. described in this paper shows that it is possible to form MOX with a variety of mixing methods, and that MOX is a sensitive high explosive capable of producing overpressures greater than comparable masses of C-4. The limited amount and quality of large-scale LOX/LNG explosive test data, and the potential formation of high-explosive MOX, create significant unknowns in the determination of hazard areas (cleared of the public) for ground operations (such as a static fire test) and flights of launch vehicles with LOX/LNG propellant. This paper will review the methods used by NASA to develop and validate the LOX/LH2 blast model using large-scale explosive test programs such as Project PYRO, the Hydrogen-Oxygen Vertical Impact (HOVI), and the Large-Scale Hydrogen-Oxygen Explosion (LSHOE) tests. This paper will describe how the same process was used to prepare an interim LOX/LNG blast model that applies to various LV failure scenarios and conservatively accounts for potential MOX formation. This paper also summarizes past and future testing and modeling efforts funded by a consortium of NASA, the US Federal Aviation Administration (FAA), and the US Space Force (USSF). These test programs should be concluded within the next 3 years and are intended to provide empirical data for model verification and validation.

Liquid Oxygen↗
NASA NTRS1992

A hybrid approach to near-optimal launch vehicle guidance

This paper evaluates a proposed hybrid analytical/numerical approach to launch-vehicle guidance for ascent to orbit injection. The feedback-guidance approach is based on a piecewise nearly analytic zero-order solution evaluated using a collocation method. The zero-order solution is then improved through a regular perturbation analysis, wherein the neglected dynamics are corrected in the first-order term. For real-time implementation, the guidance approach requires solving a set of small dimension nonlinear algebraic equations and performing quadrature. Assessment of performance and reliability are carried out through closed-loop simulation for a vertically launched 2-stage heavy-lift capacity vehicle to a low earth orbit. The solutions are compared with optimal solutions generated from a multiple shooting code. In the example the guidance approach delivers over 99.9 percent of optimal performance and terminal constraint accuracy.

Leung, Martin S. K.↗
NASA NTRS1990

Optimal guidance law development for an advanced launch system

A regular perturbation analysis is presented. Closed-loop simulations were performed with a first order correction including all of the atmospheric terms. In addition, a method was developed for independently checking the accuracy of the analysis and the rather extensive programming required to implement the complete first order correction with all of the aerodynamic effects included. This amounted to developing an equivalent Hamiltonian computed from the first order analysis. A second order correction was also completed for the neglected spherical Earth and back-pressure effects. Finally, an analysis was begun on a method for dealing with control inequality constraints. The results on including higher order corrections do show some improvement for this application; however, it is not known at this stage if significant improvement will result when the aerodynamic forces are included. The weak formulation for solving optimal problems was extended in order to account for state inequality constraints. The formulation was tested on three example problems and numerical results were compared to the exact solutions. Development of a general purpose computational environment for the solution of a large class of optimal control problems is under way. An example, along with the necessary input and the output, is given.

Calise, Anthony J.↗
NASA NTRS1990

A transient response analysis of the space shuttle vehicle during liftoff

A proposed transient response method is formulated for the liftoff analysis of the space shuttle vehicles. It uses a power series approximation with unknown coefficients for the interface forces between the space shuttle and mobile launch platform. This allows the equation of motion of the two structures to be solved separately with the unknown coefficients at the end of each step. These coefficients are obtained by enforcing the interface compatibility conditions between the two structures. Once the unknown coefficients are determined, the total response is computed for that time step. The method is validated by a numerical example of a cantilevered beam and by the liftoff analysis of the space shuttle vehicles. The proposed method is compared to an iterative transient response analysis method used by Martin Marietta for their space shuttle liftoff analysis. It is shown that the proposed method uses less computer time than the iterative method and does not require as small a time step for integration. The space shuttle vehicle model is reduced using two different types of component mode synthesis (CMS) methods, the Lanczos method and the Craig and Bampton CMS method. By varying the cutoff frequency in the Craig and Bampton method it was shown that the space shuttle interface loads can be computed with reasonable accuracy. Both the Lanczos CMS method and Craig and Bampton CMS method give similar results. A substantial amount of computer time is saved using the Lanczos CMS method over that of the Craig and Bampton method. However, when trying to compute a large number of Lanczos vectors, input/output computer time increased and increased the overall computer time. The application of several liftoff release mechanisms that can be adapted to the proposed method are discussed.

Brunty, J. A.↗
NASA NTRS2011

Using Innovative Techniques for Manufacturing Rocket Engine Hardware

Many of the manufacturing techniques that are currently used for rocket engine component production are traditional methods that have been proven through years of experience and historical precedence. As we enter into a new space age where new launch vehicles are being designed and propulsion systems are being improved upon, it is sometimes necessary to adopt new and innovative techniques for manufacturing hardware. With a heavy emphasis on cost reduction and improvements in manufacturing time, manufacturing techniques such as Direct Metal Laser Sintering (DMLS) are being adopted and evaluated for their use on J-2X, with hopes of employing this technology on a wide variety of future projects. DMLS has the potential to significantly reduce the processing time and cost of engine hardware, while achieving desirable material properties by using a layered powder metal manufacturing process in order to produce complex part geometries. Marshall Space Flight Center (MSFC) has recently hot-fire tested a J-2X gas generator discharge duct that was manufactured using DMLS. The duct was inspected and proof tested prior to the hot-fire test. Using the Workhorse Gas Generator (WHGG) test setup at MSFC?s East Test Area test stand 116, the duct was subject to extreme J-2X gas generator environments and endured a total of 538 seconds of hot-fire time. The duct survived the testing and was inspected after the test. DMLS manufacturing has proven to be a viable option for manufacturing rocket engine hardware, and further development and use of this manufacturing method is recommended.

Betts, Erin M.↗