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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
All sourcesNASA NTRSDOE OSTI
DOE OSTI2022

Early stage growth of amorphous thin film: Average kinetics, nanoscale dynamics, and pressure dependence

We used the coherent grazing incidence small angle x-ray scattering technique to study the average kinetics and nanoscale dynamics during early-stage a-WSi 2 sputter deposition. The kinetic and dynamic properties are examined as a function of pressure, which is known to be a critical factor in determining final surface roughness. Surface growth kinetics and dynamics are characterized by time parameters extracted from the height-height structure factor and correlation functions. The roughness at a given length scale reaches a maximum before relaxing down to a steady state. The lateral length scale dependence and pressure dependence are then compared among the measured kinetics and dynamics time parameters. Surfaces grown at lower pressures are smoother, leading to longer correlation times. The time parameters to reach a dynamic steady state and a kinetic steady state show contrasting pressure dependence. A dynamic steady state is reached earlier than the kinetic steady state at high pressure. A more random deposition direction and lower kinetic energy at higher pressures can explain these phenomena, along with the hypothesis that larger nanoclusters form in vapor before arriving at the surface. Here, a continuum model is applied to simulate the overall behavior with mixed success.

36 MATERIALS SCIENCE↗
DOE OSTI2024

Notes on Amplitude versus Phase Comparison Monopulse Antennas for Radar

Monopulse is a technique for determining the Direction of Arrival (DOA) of a radar echo by comparing the simultaneous signal responses from two or more antenna beams or apertures. Two principal architectures are employed: 1) amplitude-comparison monopulse, and 2) phase-comparison monopulse. For a constrained-size fully and uniformly illuminated aperture, there is no meaningful difference between the DOA angle precision achievable by an amplitude monopulse architecture versus a phase monopulse architecture. DOA angle estimation precision is almost exclusively a function of antenna size, operating wavelength, and SNR, regardless of amplitude versus phase monopulse architectures.

47 OTHER INSTRUMENTATION↗
NASA NTRS2011

Multipoint Observations of Coronal Mass Ejection and Solar Energetic Particle Events on Mars and Earth During November 2001

Multipoint spacecraft observations provide unique opportunities to constrain the propagation and evolution of interplanetary coronal mass ejections (ICMEs) throughout the heliosphere. Using Mars Global Surveyor (MGS) data to study both ICME and solar energetic particle (SEP) events at Mars and OMNI and Geostationary Operational Environmental Satellite (GOES) data to study ICMEs and SEPs at Earth, we present a detailed study of three CMEs and flares in late November 2001. In this period, Mars trailed Earth by 56deg solar longitude so that the two planets occupied interplanetary magnetic field lines separated by only approx.25deg. We model the interplanetary propagation of CME events using the ENLIL version 2.6 3-D MHD code coupled with the Wang-Sheeley-Arge version 1.6 potential source surface model, using Solar and Heliospheric Observatory (SOHO) Large Angle and Spectrometric Coronagraph (LASCO) images to determine CME input parameters. We find that multipoint observations are essential to constrain the simulations of ICME propagation, as two very different ICMEs may look very similar in only one observational location. The direction and width of the CME as parameters essential to a correct estimation of arrival time and amplitude of the ICME signal. We find that these are problematic to extract from the analysis of SOHO/LASCO images commonly used for input to ICME propagation models. We further confirm that MGS magnetometer and electron reflectometer data can be used to study not only ICME events but also SEP events at Mars, with good results providing a consistent picture of the events when combined with near-Earth data.

Falkenberg, T. V.↗
NASA NTRS2016

Outer Radiation Belt Dropout Dynamics Following the Arrival of Two Interplanetary Coronal Mass Ejections

Magnetopause shadowing and wave-particle interactions are recognized as the two primary mechanisms for losses of electrons from the outer radiation belt. We investigate these mechanisms, sing satellite observations both in interplanetary space and within the magnetosphere and particle drift modeling. Two interplanetary shocks sheaths impinged upon the magnetopause causing a relativistic electron flux dropout. The magnetic cloud (C) and interplanetary structure sunward of the MC had primarily northward magnetic field, perhaps leading to a concomitant lack of substorm activity and a 10 day long quiescent period. The arrival of two shocks caused an unusual electron flux dropout. Test-particle simulations have shown 2 to 5 MeV energy, equatorially mirroring electrons with initial values of L 5.5can be lost to the magnetosheath via magnetopause shadowing alone. For electron losses at lower L-shells, coherent chorus wave-driven pitch angle scattering and ULF wave-driven radial transport have been shownto be viable mechanisms.

Alves, L. R.↗
NASA NTRS2002

Management of Total Pressure Recovery, Distortion and High Cycle Fatigue in Compact Air Vehicle Inlets

It is the purpose of this study to demonstrate the viability and economy of Response Surface Methods (RSM) and Robustness Design Concepts (RDC) to arrive at micro-secondary flow control installation designs that maintain optimal inlet performance over a range of the mission variables. These statistical design concepts were used to investigate the robustness properties of 'low unit strength' micro-effector installations. 'Low unit strength' micro-effectors are micro-vanes set at very low angles-of-incidence with very long chord lengths. They were designed to influence the near wall inlet flow over an extended streamwise distance, and their advantage lies in low total pressure loss and high effectiveness in managing engine face distortion.

Anderson, Bernhard H.↗
NASA NTRS2006

Trajectory Design of the Lunar Impactor Mission Concept

The National Aeronautics and Space Administration (NASA) solicited proposals in 2006 for an opportunity to include a small secondary payload with the launch of the Lunar Reconnaissance Orbiter (LRO) scheduled for October 2008. The cost cap of the proposal was between $50 and $80M, and the mass cap was 1,000 kilograms. JPL proposed a Lunar Impactor (LI) concept for this solicitation. The mission objective of LI was to impact the permanently shadowed region of a South polar crater ultimately to detect the presence of water. The detection of water ice would prove to be an important factor on future lunar exploration. NASA Ames Research Center also proposed a similar concept, the Lunar Crater observation and Sensing Satellite (LCROSS), which was selected by NASA for the mission. However, in this paper, the trajectory design of the LI proposed by JPL is considered. Since the LI spacecraft was to be launched on the LRO launch vehicle as a secondary payload, its initial trajectory must be diverted at some later time from the LRO trans-lunar trajectory for the subsequent impact. Several such trajectories have been considered, where each trajectory option fields some specific values for the mission parameters. The mission parameters include the availability of LRO instruments at the time of impact for the observation by LRO, the mission duration, the impact velocity, the impact angle, etc. It is possible for the LI to be deflected with a relatively low delta-V to impact a South polar crater at a reasonable impact velocity and impact angle directly with no delay. However, the instruments on-board LRO may not be ready for observation. Thus, several delayed trajectory options have been considered further. The lunar phase at the time of impact may also play an important factor for observation, especially from Earth. Several lunar flyby trajectory maneuvers have been identified to arrive at the Moon for impact at the desired lunar phase. By using a combination of these successive lunar flyby maneuvers, the impact lunar phase may be adjusted to the desired location. A few such trajectories have been suggested. Also, some attempts have been made to maximize the impact velocity by converting the impact trajectory into a retrograde orbit with respect to Earth. Since these types of trajectories take advantage of the Sun-Earth three-body region to minimize the delta-V, the mission duration is relatively long. A few such trajectories are suggested. Also, an attempt has been made to adjust the lunar impact within a desired time period for the optimum Earth observation for the above trajectories. The mission parameters resulting from each trajectory option above are considered and weighed against the cost and robustness of the mission in a brief summary.

trajectory design↗
NASA NTRS1980

Comparison of ISEE-1 and -3 interplanetary magnetic field observations

Three hourly correlation coefficients and the lag at maximum correlation are computed using one minute averages of ISEE-1 and -3 magnetometer data during the period in which ISEE-3 moved from the earth to its halo orbit around the libration point. The maximum correlation coefficients are highly variable ranging from close to zero to almost unity. The lags, while on the average approximating the expected corotation delay, have very large departures from this value. These results suggest that the normals to the planes separating fields of differing orientation often make large angles to the ecliptic plane and/or that the interplanetary magnetic field has a significant amount of bending on a scale length the order of 200 earth radii. Furthermore, there often appears to be a significant amount of propagating structure in the IMF. Thus, ISEE-3 magnetic field measurements should be used with caution if precise timing of arrival at the earth or precise directions of the field upon arrival are important.

Russell, C. T.↗
NASA NTRS1980

Radar estimation of slant path rain attenuation at frequencies above 10 GHz and comparisons with measured multi-season results

Techniques and results pertaining to estimating Earth satellite path rain attenuation events and statistics using radar at nonattenuating wavelengths are presented. The radar formulation and beam scanning methods are described and the procedure for relating the radar measured Rayleigh reflectivity to the high frequency Mie attenuation coefficient are given. Examples of radar derived single terminal statistics and estimation criteria as they relate to path angle and frequency are reviewed. Radar derived space diversity statistics and their dependence on terminal spacing and frequency are described. Site diversity performance curves obtained by radar and radiometry are compared with each other demonstrating the utility of radar methods. Results of a multi-year experiment to test, refine, and establish accuracies of radar methods for arriving at estimates of rain attenuation along an Earth-satellite path are discussed. Comparisons of measured and radar estimated fade events are presented and found to be good. Comparisons of cumulative fade distributions show agreement to be excellent giving an rms deviation of 1 dB.

Goldhirsh, J.↗
NASA NTRS2006

Low-Cost Propellant Launch From a Tethered Balloon

A document presents a concept for relatively inexpensive delivery of propellant to a large fuel depot in low orbit around the Earth, for use in rockets destined for higher orbits, the Moon, and for remote planets. The propellant is expected to be at least 85 percent of the mass needed in low Earth orbit to support the NASA Exploration Vision. The concept calls for the use of many small ( 10 ton) spin-stabilized, multistage, solid-fuel rockets to each deliver 250 kg of propellant. Each rocket would be winched up to a balloon tethered above most of the atmospheric mass (optimal altitude 26 2 km). There, the rocket would be aimed slightly above the horizon, spun, dropped, and fired at a time chosen so that the rocket would arrive in orbit near the depot. Small thrusters on the payload (powered, for example, by boil-off gases from cryogenic propellants that make up the payload) would precess the spinning rocket, using data from a low-cost inertial sensor to correct for small aerodynamic and solid rocket nozzle misalignment torques on the spinning rocket; would manage the angle of attack and the final orbit insertion burn; and would be fired on command from the depot in response to observations of the trajectory of the payload so as to make small corrections to bring the payload into a rendezvous orbit and despin it for capture by the depot. The system is low-cost because the small rockets can be mass-produced using the same techniques as those to produce automobiles and low-cost munitions, and one or more can be launched from a U.S. territory on the equator (Baker or Jarvis Islands in the mid-Pacific) to the fuel depot on each orbit (every 90 minutes, e.g., any multiple of 6,000 per year).

Wilcox, Brian↗
NASA NTRS2004

Aerocapture Guidance Performance for the Neptune Orbiter

A performance evaluation of the Hybrid Predictor corrector Aerocapture Scheme (HYPAS) guidance algorithm for aerocapture at Neptune is presented in this paper for a Mission to Neptune and the Neptune moon Triton'. This mission has several challenges not experienced in previous aerocapture guidance assessments. These challengers are a very high Neptune arrival speed, atmospheric exit into a high energy orbit about Neptune, and a very high ballistic coefficient that results in a low altitude acceleration capability when combined with the aeroshell LD. The evaluation includes a definition of the entry corridor, a comparison to the theoretical optimum performance, and guidance responses to variations in atmospheric density, aerodynamic coefficients and flight path angle for various vehicle configurations (ballistic numbers). The benefits of utilizing angle-of-attack modulation in addition to bank angle modulation to improve flight performance is also discussed. The results show that despite large sensitivities in apoapsis targeting, the algorithm performs within the allocated AV budget for the Neptune mission bank angle only modulation. The addition of angle-of-attack modulation with as little as 5 degrees of amplitude significantly improves the scatter in final orbit apoapsis. Although the angle-of-attack modulation complicates the vehicle design, the performance enhancement reduces aerocapture risk and reduces the propellant consumption needed to reach the high energy target orbit for a conventional propulsion system.

Masciarelli, James P.↗
NASA NTRS2014

Mars Sample Return Using Commercial Capabilities: Propulsive Entry, Descent, and Landing of a Capsule Form Vehicle

This paper describes a critical portion of the work that has been done at NASA, Ames Research Center regarding the use of the commercially developed Dragon capsule as a delivery vehicle for the elements of a high priority Mars Sample Return mission. The objective of the investigation was to determine entry and landed mass capabilities that cover anticipated mission conditions. The "Red Dragon", Mars configuration, uses supersonic retro-propulsion, with no required parachute system, to perform Entry, Descent, and Landing (EDL) maneuvers. The propulsive system proposed for use is the same system that will perform an abort, if necessary, for a human rated version of the Dragon capsule. Standard trajectory analysis tools are applied to publically available information about Dragon and other legacy capsule forms in order to perform the investigation. Trajectory simulation parameters include entry velocity, flight path angle, lift to drag Ratio (L/D), landing site elevation, atmosphere density, and total entry mass, in addition engineering assumptions for the performance of the propulsion system are stated. Mass estimates for major elements of the overall proposed architecture are coupled to this EDL analysis to close the overall architecture. Three synodic launch opportunities, beginning with the 2022 opportunity, define the arrival conditions. Results state the relations between the analysis parameters as well as sensitivities to those parameters. The EDL performance envelope includes landing altitudes between 0 and -4 km referenced to the Mars Orbiter Laser Altimeter datum as well as minimum and maximum atmosphere density. Total entry masses between 7 and 10 mt are considered with architecture closure occurring between 9.0 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the effect of the entry conditions on the Thermal Protection System (TPS) currently in use for Dragon missions shows no significant stressors. A useful payload mass of 2.0 mt is provided and includes mass and grow allowance for a Mars Ascent Vehicle (MAV), Earth Return Vehicle (ERV), and mission unique equipment. The useful payload supports an architecture that receives a sample from another surface asset and sends it directly back to Earth for recovery in a high Earth orbit. The work shows that emerging commercial capabilities as well as previously studied EDL methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that will also use propulsive EDL techniques

EDL↗
NASA NTRS2012

A Robust Initialization Scheme for a Lateral Trajectory Optimization Problem with Time of Arrival Windows

We present a robust initialization scheme that estimates parameter values for the numerical solution of a two-point boundary value problem. The two-point boundary value problem formulation stems from the optimization of a cost functional subject to the dynamics of a simplified lateral aircraft model and other constraints. Leveraging regular perturbation methods, initial parameter estimates are analytically determined and used to initialize a gradient descent optimization routine which is shown to rapidly converge over a range of initial aircraft positions and heading angles. Additionally, the velocity of the aircraft is optimized to ensure the trajectory of the aircraft terminates within a desired region in both time and space.

lateral trajectory optimization↗
DOE OSTI2023

First estimation of high-resolution solar photovoltaic resource maps over China with Fengyun-4A satellite and machine learning

Fengyun-4A (FY-4A), which is the latest-generation Chinese geostationary meteorological satellite, measures solar reflection and thermal emission with high temporal, spatial, and spectral resolutions. It is expected to be highly beneficial for solar resource assessment and forecasting in China. This study is the first to estimate, using FY-4A and a random forest model, the global horizontal irradiance (GHI) at a 4-km–15-min spatio-temporal resolution over China, as a means to arrive at a solar photovoltaic (PV) resource map. In terms of GHI estimates, the root mean square error and mean bias error between hourly measured and retrieved values are 147.02 (35.2%), –5.64 W/m 2 (–1.4%), respectively, whereas the values of daily estimates are 29.20 (18.0%), –2.97 W/m 2 (–1.3%). The retrieval accuracy is found much better for instances with solar zenith angles smaller than 60°. Relatively larger errors are found at locations in the Sichuan Basin and northeastern China, which can be attributed to bright surfaces and/or strong cloud transients. With the retrieved irradiance, PV resource is derived through a physical model chain. The annual mean PV resource map suggests that, over most of the west areas, the annual mean effective irradiance exceeds 1700 kWh/m 2 , with the highest value found in Tibet (around 2000 kWh/m 2 per annum). Eastern China has an annual effective irradiance of only 1300–1500 kWh/m 2 . In conclusion, the region with poorest solar resource is the Sichuan Basin (less than 1100 kWh/m 2 per annum).

29 ENERGY PLANNING, POLICY, AND ECONOMY↗
NASA NTRS2004

Combining Satellite Data, Trajectory Modeling and Surface Insolation Measurements to Deduce the Direct Radiative Effect of Smoke Aerosol

In this paper, we have introduced a method of inferring the radiative effect of smoke aerosols using a technique that combines satellite remote sensing with trajectory modeling. The results shown here clearly show large flux biases between theoretical and measured radiative fluxes correlate with the arrival of smoke aerosol to the area. Further analysis is required to convincingly demonstrate that the reason for these differences is the radiative effect of the smoke aerosols. To do this, the estimated fluxes taken from the ERA-15 will be recomputed every 3 hours using International Satellite Cloud Climatology Project (ISCCP) data set entitled DX gridded to a 1o equal angle resolution (see paper 7B.2 for details). Surface radiometric and ancillary data for several more Canadian surface sites are being obtained at minute temporal resolution. The ultimate purpose of this research is to derive aerosol smoke maps for fire events such as this to be included in an aerosol climatology and be incorporated in the computation of the earth's surface radiation budget to better understand the radiative effect of aerosols.

Stackhouse, Paul W., Jr.↗
NASA NTRS2007

Entry, Descent, and Landing Operations Analysis for the Genesis Entry Capsule

On September 8, 2004, the Genesis spacecraft returned to Earth after spending 29 months about the sun-Earth libration point (L1) collecting solar wind particles. Four hours prior to Earth arrival, the sample return capsule containing the samples was released for entry and subsequent landing at the Utah Test and Training Range. This paper provides an overview of the entry, descent, and landing trajectory analysis that was performed during the mission operations phase leading up to final approach to Earth. The final orbit determination solution produced an inertial entry flight-path angle of -8.002 deg (which was the desired nominal value) with a 3-sigma error of +/-0.0274 deg (a third of the requirement). The operations effort accurately delivered the entry capsule to the desired landing site. The final landing location was 8.3 km from the target, and was well within the allowable landing area. Overall, the Earth approach operation procedures worked well and there were no issues (logistically or performance based) that arose. As a result, the process of targeting a capsule from deep space and accurately landing it on Earth was successfully demonstrated.

Desai, Prasun N.↗
NASA NTRS1993

Whistlers and plasmaspheric hiss - Wave directions and three-dimensional propagation

Wave propagation directions are determined on the basis of wave data from the DE 1 satellite showing simultaneously nonducted whistlers and hiss. Hiss wave normal angles are determined as about 70 and 77 deg for f = 3.5 and 2.5 kHz, respectively, with the wave vector being almost perpendicular to the meridional plane. A novel approximate analytical formulation of 3D propagation of whistler waves is developed and used to model the drift of magnetospherically reflected whistlers in azimuth. It is shown that depending on initial parameters, the time of arrival of whistler rays at a fixed observation point can differ by 10-20 s, with signals from different magnetospherically reflected whistlers overlapping to evolve into a hisslike signal. The total azimuthal drift of whistler rays is found to not exceed about 30 deg, so that plasmaspheric hiss may be produced by nonducted whistlers at longitudes correlated with the location of thunderstorm activity.

Draganov, A. B.↗
NASA NTRS1993

Single-input/single-output transfer functions for a flexible slewing link

Truncated linear models based on constrained and unconstrained mode expansions for a flexible link with torque actuation at its hub led us to examine the behavior of the open-loop-zeros of five commonly used SISO transfer functions. We arrive at new conclusions with regard to minimum- and non-minimum-phase zeros, and passivity of the transfer functions. The flexible-to-rigid inertia (FRI) ratio plays an important role in the determination of these characteristics. This paper illustrates how the zero locations of some of the transfer functions are dramatically affected by the FRI ratio. We also investigate the performance of a hub angle, hub rate, and relative tip position static feedback controller based on a reduced-order constrained mode model when applied to both unconstrained and constrained mode models of the same dimension. Simulations indicate that the constrained model can be stable while the unconstrained model is not for high FRI ratios. Our findings can provide useful information to designers in cases where the controller is sensitive to the 1ocation of the plant's zeros.

Barbieri, Enrique↗
NASA NTRS1988

Surveying System

Sunrise Geodetic Surveys are setting up their equipment for a town survey. Their equipment differs from conventional surveying systems that employ transit rod and chain to measure angles and distances. They are using ISTAC Inc.'s Model 2002 positioning system, which offers fast accurate surveying with exceptional signals from orbiting satellites. The special utility of the ISTAC Model 2002 is that it can provide positioning of the highest accuracy from Navstar PPS signals because it requires no knowledge of secret codes. It operates by comparing the frequency and time phase of a Navstar signal arriving at one ISTAC receiver with the reception of the same set of signals by another receiver. Data is computer processed and translated into three dimensional position data - latitude, longitude and elevation.

Source record↗