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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Specificity and Transfer in Learning How to Follow Navigation Instructions

We report a series of experiments that use a navigation task in which instructions for navigating in a space displayed as grids on a computer screen are given to subjects who then attempt to follow them by mouse clicking on the grids. The navigation task was broken down into component dimensions (e.g., presentation mode of the instructions, length of the instructions, characteristics of the display, size of the grids, response type). For each task dimension, one condition was used at training and the same or another condition was used at test. Each task dimension was examined in terms of two measures. One measure provided an index of transfer (i.e., better performance at test than at training when test and training involved different conditions), and the other provided an index of specificity (i.e., better performance at test when training and test conditions were the same than when training and test conditions were different). By and large, these two indices were complementary, so there was evidence of either transfer or specificity but not both. For one dimension transfer but no specificity was evident, and for another dimension specificity but no transfer was evident. For the remaining dimensions, however, there was asymmetrical transfer, with transfer evident for some conditions and specificity evident for others. The findings are interpreted within the procedural reinstatement framework. They have practical implications concerning how to optimize training and how much fidelity to the testing situation is necessary when training.

Healy, Alice F.↗
NASA NTRS2009

Three-Dimensional Unsteady Simulation of Aerodynamics and Heat Transfer in a Modern High Pressure Turbine Stage

Unsteady 3-D RANS simulations have been performed on a highly loaded transonic turbine stage and results are compared to steady calculations as well as to experiment. A low Reynolds number k-epsilon turbulence model is employed to provide closure for the RANS system. A phase-lag boundary condition is used in the tangential direction. This allows the unsteady simulation to be performed by using only one blade from each of the two rows. The objective of this work is to study the effect of unsteadiness on rotor heat transfer and to glean any insight into unsteady flow physics. The role of the stator wake passing on the pressure distribution at the leading edge is also studied. The simulated heat transfer and pressure results agreed favorably with experiment. The time-averaged heat transfer predicted by the unsteady simulation is higher than the heat transfer predicted by the steady simulation everywhere except at the leading edge. The shock structure formed due to stator-rotor interaction was analyzed. Heat transfer and pressure at the hub and casing were also studied. Thermal segregation was observed that leads to the heat transfer patterns predicted by steady and unsteady simulations to be different.

Shyam, Vikram↗
NASA NTRS2014

Towards a General Equation for the Survival of Microbes Transferred between Solar System Bodies

It should be possible to construct a general equation describing the survival of microbes transferred between Solar System bodies. Such an equation will be useful for constraining the likelihood of transfer of viable organisms between bodies throughout the lifetime of the Solar System, and for refining Planetary Protection constraints placed on future missions. We will discuss the construction of such an equation, present a plan for definition of pertinent factors, and will describe what research will be necessary to quantify those factors. Description: We will examine the case of microbes transferred between Solar System bodies as residents in meteorite material ejected from one body (the "intial body") and deposited on another (the "target body"). Any microbes transferred in this fashion will experience four distinct phases between their initial state on the initial body, up to the point where they colonize the target body. Each of these phases features phenomena capable of reducing or exterminating the initial microbial population. They are: 1) Ejection: Material is ejected from the initial body, imparting shock followed by rapid desiccation and cooling. 2) Transport: Material travels through interplanetary space to the target body, exposing a hypothetical microbial population to extended desiccation, irradiation, and temperature extremes. 3) Infall: Material is deposited on the target body, diminishing the microbial population through shock, mass loss, and heating. 4) Adaptation: Any microbes which survive the previous three phases must then adapt to new chemophysical conditions of the target body. Differences in habitability between the initial and target bodies dominate this phase. A suitable general-form equation can be assembled from the above factors by defining the initial number of microbes in an ejected mass and applying multiplicitive factors based on the physical phenomena inherent to each phase. It should be possible to present the resulting equation in terms of initial ejection mass, ejection shock magnitude, transfer time, initial microbial load and/or other terms and generate graphs defining the number of surviving microbes. The general form of the equation is: x(sub f) = x(sub i) f(sub1) f(sub 2) f(sub 3) f(sub 4) Where x(sub f) is the final number of microbes to survive transfer, x(sub i) is the initial population prior to ejection, and f(sub 1-4) are mortality factors for the four phases described above. Among other considerations, f(sub 1) will vary with respect to impact shock magnitude and f(sub 2) will be time-dependent. Considerable research has been performed to date to quantify the survival rates of various microbes in response to portions of these four phases, both as vegetative cells and/or spores. Results indicate that many species tend to respond differently to the pertinent mortality factors, especially in the case of extremophiles. Therefore, a complete equation will include species-specific responses to the mortality factors.

Fries, M.↗
NASA NTRS2014

Lunar Cube Transfer Trajectory Options

Contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these designs can be restricted by the selection of the Cubesat subsystem design such as propulsion or communication. Nonetheless, many trajectory options can be designed with have a wide range of transfer durations, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several design options including deployment into low Earth orbit (LEO), geostationary transfer orbits (GTO), and higher energy direct lunar transfer orbits. In addition to direct transfer options from these initial orbits, we also investigate the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory. In this article we examine several design options that meet the above limited deployment and subsystem drivers. We study ways that both impulsive and low-thrust Solar Electric Propulsion (SEP) engines can be used to place the Cubesat first into a highly eccentric Earth orbit, enter the Moon's Sphere of Influence, and finally achieve a highly eccentric lunar orbit. We show that such low-thrust transfers are feasible with a realistic micro-thruster model, assuming that the Cubesat can generate sufficient power for the SEP. Two examples are shown here: (1) A Cubestat injected by Exploration Mission 1 (EM-1) then employing low thrust; and (2) a CubSat deployed in a GTO, then employing impulsive maneuvers. For the EM-1 injected initial design, we increase the EM-1 targeted lunar flyby distance to reduce the energy of the lunar flyby to match that of a typical lMoon system heteroclinic manifold. Figure 1 presents an option that encompasses the similar dynamics as that of the ARTEMIS mission design. Low-thrust maneuvers are used along the manifold trajectory to raise perigee to that of a lunar orbit, adjust the timing with respect to the Moon, rotate the line of apsides, and target a ballistic lunar encounter. In this design a second flyby decreases the orbital energy with respect to the Moon, so that C3 -0.1 km2s2. Another design, shown in Figure 2 emanates from a GTO then uses impulsive maneuvers to phase onto a local Earth-Moon manifold, which then transfers the CubeSat to a lunar encounter.

Dynamical System↗
NASA NTRS2011

Targeting Low-Energy Transfers to Low Lunar Orbit

A targeting scheme is presented to build trajectories from a specified Earth parking orbit to a specified low lunar orbit via a low-energy transfer and up to two maneuvers. The total transfer delta V (velocity) is characterized as a function of the Earth parking orbit inclination and the departure date for transfers to each given low lunar orbit. The transfer delta V (velocity) cost is characterized for transfers constructed to low lunar polar orbits with any longitude of ascending node and for transfers that arrive at the Moon at any given time during a month.

mission designs↗
NASA NTRS2009

Low-Energy Ballistic Transfers to Lunar Halo Orbits

Recent lunar missions have begun to take advantage of the benefits of low-energy ballistic transfers between the Earth and the Moon rather than implementing conventional Hohmann-like lunar transfers. Both Artemis and GRAIL plan to implement low-energy lunar transfers in the next few years. This paper explores the characteristics and potential applications of many different families of low-energy ballistic lunar transfers. The transfers presented here begin from a wide variety of different orbits at the Earth and follow several different distinct pathways to the Moon. This paper characterizes these pathways to identify desirable low-energy lunar transfers for future lunar missions.

Three-Body Halo Orbits↗
NASA NTRS2017

Low Thrust Cis-Lunar Transfers Using a 40 kW-Class Solar Electric Propulsion Spacecraft

This paper captures trajectory analysis of a representative low thrust, high power Solar Electric Propulsion (SEP) vehicle to move a mass around cis-lunar space in the range of 20 to 40 kW power to the Electric Propulsion (EP) system. These cis-lunar transfers depart from a selected Near Rectilinear Halo Orbit (NRHO) and target other cis-lunar orbits. The NRHO cannot be characterized in the classical two-body dynamics more familiar in the human spaceflight community, and the use of low thrust orbit transfers provides unique analysis challenges. Among the target orbit destinations documented in this paper are transfers between a Southern and Northern NRHO, transfers between the NRHO and a Distant Retrograde Orbit (DRO) and a transfer between the NRHO and two different Earth Moon Lagrange Point 2 (EML2) Halo orbits. Because many different NRHOs and EML2 halo orbits exist, simplifying assumptions rely on previous analysis of orbits that meet current abort and communication requirements for human mission planning. Investigation is done into the sensitivities of these low thrust transfers to EP system power. Additionally, the impact of the Thrust to Weight ratio of these low thrust SEP systems and the ability to transit between these unique orbits are investigated.

cislunar↗
NASA NTRS2018

Analysis of Cislunar Transfers from a Near Rectilinear Halo Orbit with High Power Solar Electric Propulsion

As government and commercial interest in the exploration of the Moon and cislu- nar space has grown, Near Rectilinear Halo Orbits (NRHOs) have shown to be of particular interest as staging orbits for human exploration of the Moon. Once in such staging orbits, low thrust solar electric propulsion (SEP) can enable efficient transfer to other orbits in cislunar space. This paper captures ongoing analysis to design efficient transfers of a massive spacecraft from a L2 Southern NRHO to a Distant Retrograde Orbit, L1 Northern NRHO, and Flat L2 Halo Orbit using low thrust SEP. For each transfer type, reference transfer is designed for an assumed 39 t spacecraft with 26.6 kW SEP system. For each reference transfer, analysis is completed to understand the sensitivity of the transfer to changes in initial mass and SEP power and identify the optimal number of thrusters to use for a given combination of mass and power.

SEP↗
NASA NTRS2018

Analysis of Cislunar Transfers from a Near Rectilinear Halo Orbit with High Power Solar Electric Propulsion

As government and commercial interest in the exploration of the Moon and cislunar space has grown, Near Rectilinear Halo Orbits (NRHOs) have shown to be of particular interest as staging orbits for human exploration of the Moon. Once in such staging orbits, low thrust solar electric propulsion (SEP) can enable efficient transfer to other orbits in cislunar space. This paper captures ongoing analysis to design efficient transfers of a massive spacecraft from a L2 Southern NRHO to a Distant Retrograde Orbit, L1 Northern NRHO, and Flat L2 Halo Orbit using low thrust SEP. For each transfer type, reference transfer is designed for an assumed 39 t spacecraft with 26.6 kW SEP system. For each reference transfer, analysis is completed to understand the sensitivity of the transfer to changes in initial mass and SEP power and identify the optimal number of thrusters to use for a given combination of mass and power.

Near Rectilinear Halo Orbits (NRHO)↗
NASA NTRS2020

Missed Thrust Analysis and Design for Low Thrust Cislunar Transfers

This paper details the analysis of the missed thrust problem for low thrust cislunar transfers. Missed thrust analysis is completed for a reference NRHO to DRO transfer that is designed to maximize the final mass without consideration of robustness to unexpected loss of thrust. A new missed thrust design method is presented to include the robustness of the transfer as part of the optimization problem by including a minimal number of branching trajectories tied to the start of thrust arcs in the reference transfer and optimizing them all simultaneously. Results from variations of this approach are presented for and compared to the reference transfer designed without consideration for missed thrust. The results show that this new method can reduce the additional propellant required for recovery from an unexpected 7-day outage by up to 90% without significant increase to the propellant required to complete the reference transfer.

Grebow, Daniel J.↗
NASA NTRS2022

Cryogenic Spray Quenching of A Simulated Propellant Storage Tank Wall With Heat Transfer Enhancement By A Thin-Film Coating and Flow Pulsing in Microgravity

Human space exploration to the Moon, Mars, and possibly asteroids is NASA’s biggest challenge for the new millennium. One of the critical elements to this mission is the effective, sufficient, and reliable supply of cryogenic propellant fluids. Future lower-earth-orbiting (LEO) propellant fuel depots and human-carrying orbital transfer spacecraft flying to the moon and Mars will have to utilize the high thrust and high efficiency of liquid cryogenic chemical propulsion or nuclear thermal propulsion. Efficient in-space tank-to-tank propellant transfer (propellant fuel depot to orbital transfer spacecraft) of cryogenic propellants is an enabling technology for the planned Crewed Mars Surface Mission. The transfer of cryogenic propellants in space, however, has yet to be accomplished, solely due to the unavailability of cryogenic quenching heat transfer data during chilldown (quenching) and filling of the propellant receiver tank in reduced gravity and microgravity as liquid propellant cannot be stored in a required liquid state until the tank is quenched down to the liquid temperature. Therefore, highly energy efficient thermal-fluid management breakthrough concepts to conserve and minimize the cryogen consumption during propellant transfer have become the focus of research and engineering development, especially for the deep-space mission to Mars. In this paper, we introduce such concepts and demonstrate their feasibility for cryogenic storage tank chilldown in parabolic flights under a simulated space microgravity condition. In order to maximize the storage tank chilldown efficiency for the least amount of cryogen consumption, the technology adopted included cryogenic spray cooling, Teflon thin-film coating of the simulated tank surface, and spray flow pulsing. The completed flight experiments successfully demonstrated that spray cooling is the most efficient cooling method for the tank chilldown in microgravity. In microgravity, Teflon coating alone can improve the efficiency up to 72% and the efficiency can be improved up to 59% by flow pulsing alone. However, Teflon coating together with flow pulsing was found to substantially enhance the chilldown efficiency in microgravity for up to 113%.

spray↗
NASA NTRSDate not supplied

Human Factors Considerations for Pressurized Crew Transfer Between Ascent or Lander Spacecraft and Surface Assets

An option for pressurized crew transfer between surface assets and lander or ascent spacecraft offers multiple advantages. The most important are the loss of life risks due to crew member injuries or surface spacesuit failures. However, even in nominal scenarios pressurized crew transfer is advantageous, protecting lander and ascent spacecraft as well as launch and entry spacesuits from dust exposure and aiding in adaptation to gravity following prolonged periods of microgravity. Several examples of lunar and Mars lander or ascent spacecraft have been released publicly and can serve as examples of the range of potential vehicles that may benefit from pressurized crew transfer. There are also examples of surface assets that can be used to provide crew transfer, including both current surface mobility or crew transfer concepts and terrestrial examples that could be adapted to use on the Moon or Mars. In all cases, the resulting passageway must be acceptable for use and protected from itself experiencing dust contamination. Key design considerations will be discussed.

Artemis↗
NASA NTRS2014

Modified Displacement Transfer Functions for Deformed Shape Predictions of Slender Curved Structures with Varying Curvatives

To eliminate the need to use finite-element modeling for structure shape predictions, a new method was invented. This method is to use the Displacement Transfer Functions to transform the measured surface strains into deflections for mapping out overall structural deformed shapes. The Displacement Transfer Functions are expressed in terms of rectilinearly distributed surface strains, and contain no material properties. This report is to apply the patented method to the shape predictions of non-symmetrically loaded slender curved structures with different curvatures up to a full circle. Because the measured surface strains are not available, finite-element analysis had to be used to analytically generate the surface strains. Previously formulated straight-beam Displacement Transfer Functions were modified by introducing the curvature-effect correction terms. Through single-point or dual-point collocations with finite-elementgenerated deflection curves, functional forms of the curvature-effect correction terms were empirically established. The resulting modified Displacement Transfer Functions can then provide quite accurate shape predictions. Also, the uniform straight-beam Displacement Transfer Function was applied to the shape predictions of a section-cut of a generic capsule (GC) outer curved sandwich wall. The resulting GC shape predictions are quite accurate in partial regions where the radius of curvature does not change sharply.

displacement transfer functions↗
NASA NTRS2015

Variable-Domain Displacement Transfer Functions for Converting Surface Strains into Deflections for Structural Deformed Shape Predictions

Variable-Domain Displacement Transfer Functions were formulated for shape predictions of complex wing structures, for which surface strain-sensing stations must be properly distributed to avoid jointed junctures, and must be increased in the high strain gradient region. Each embedded beam (depth-wise cross section of structure along a surface strain-sensing line) was discretized into small variable domains. Thus, the surface strain distribution can be described with a piecewise linear or a piecewise nonlinear function. Through discretization, the embedded beam curvature equation can be piece-wisely integrated to obtain the Variable-Domain Displacement Transfer Functions (for each embedded beam), which are expressed in terms of geometrical parameters of the embedded beam and the surface strains along the strain-sensing line. By inputting the surface strain data into the Displacement Transfer Functions, slopes and deflections along each embedded beam can be calculated for mapping out overall structural deformed shapes. A long tapered cantilever tubular beam was chosen for shape prediction analysis. The input surface strains were analytically generated from finite-element analysis. The shape prediction accuracies of the Variable- Domain Displacement Transfer Functions were then determined in light of the finite-element generated slopes and deflections, and were fofound to be comparable to the accuracies of the constant-domain Displacement Transfer Functions

displacement transfer functions↗
NASA NTRS2019

A Fast Vector Radiative Transfer Model for the Atmosphere-Ocean Coupled System

To infer atmospheric and oceanic constituent properties from polarimetric observations, an efficient and accurate retrieval algorithm is desirable. In-line radiative transfer calculations are indispensable if a large state vector, including both atmospheric profiles and surface properties, is used to improve retrieval accuracy. However, in-line radiative transfer calculations are usually not computationally efficient for remote sensing applications. Therefore, there is a pressing need to develop an accurate and fast vector radiative transfer model (RTM) to fully utilize satellite polarimetric observations. This paper reports on a fast vector RTM, referred to as TAMU-VRTM, in support of polarimetric remote sensing, which is capable of simulating the Stokes vector values observed at the top of the atmosphere and at the surface by fully considering absorption, scattering, and emission in the atmosphere and ocean. Gaseous absorption is parameterized with respect to gas concentration, temperature, and pressure, by using a regression method applicable to an inhomogeneous atmospheric path. An efficient two-component approach combining the small angle approximation and the adding-doubling method is utilized to solve the vector radiative transfer equation (RTE). The thermal emission component of the RTE solution is obtained by an efficient doubling process. The air-sea interface is treated as a wind-ruffled rough surface in the model to mimic a realistic ocean surface. Several oceanic optical property models are introduced to model ocean inherent optical properties. To demonstrate the applicability of the TAMUVRTM, simulations are compared with satellite observations, and results from other vector radiative transfer methods including benchmarks.

radiative transfer in coupled atmosphere-ocean sys↗
NASA NTRSDate not supplied

Numerical Prediction of Heat Transfer Coefficients during Xenon Tank Fill in Microgravity

NASA’s Gateway will serve as an orbital outpost to enable sustained human lunar exploration and facilitate access to destinations beyond the Moon. To maintain sustainable deep space missions, on-orbit propellant transfer and refueling must be realized. One of Gateway’s critical components, the Power and Propulsion Element (PPE), will use solar electric propulsion to perform attitude control maneuvers and orbit transfers. Xenon will be used as the propellant for the PPE and stored in a composite overwrapped pressure vessel (COPV). During the propellant refueling process, xenon will compress causing the fluid and tank walls to warm. There is concern that the bond between the tank liner and composite may degrade at elevated temperatures, along with the possibility of the tank exceeding design pressure. While on-orbit testing is being proposed, numerical models are being used to predict the thermal response and reduce risk during a refueling operation in microgravity where natural convection is diminished. Accurate numerical models may be used to inform decisions for follow-on testing, or design and operation of refueling architectures. Challenges of modeling xenon tank fill in microgravity include non-ideal gas behavior, operating near the fluid’s critical point, and lack of experimental flight data. This study presents a computational fluid dynamics (CFD) model with conjugate heat transfer that is used to predict averaged heat transfer coefficients during tank filling. At the anticipated injection flow rates, the CFD results showed that forced convection dominates in microgravity. Initial results from a multi-node Thermal Desktop model that assumed a purely conducting fluid (Nu=1) over-predicted the CFD fluid temperature and pressure, while a similar model allowing for natural convection under-predicted. The heat transfer coefficients predicted by CFD were implemented into the multi-node model, and the COPV thermal response from the two computational approaches were compared with good agreement.

Heat Transfer↗
NASA NTRS1959

Measurements of Heat-Transfer and Friction Coefficients for Helium Flowing in a Tube at Surface Temperatures up to 5900 Deg R

Measurements of average heat transfer and friction coefficients and local heat transfer coefficients were made with helium flowing through electrically heated smooth tubes with length-diameter ratios of 60 and 92 for the following range of conditions: Average surface temperature from 1457 to 4533 R, Reynolds numbe r from 3230 to 60,000, heat flux up to 583,200 Btu per hr per ft2 of heat transfer area, and exit Mach numbe r up to 1.0. The results indicate that, in the turbulent range of Reynolds number, good correlation of the local heat transfer coefficients is obtained when the physical properties and density of helium are evaluated at the surface temperature. The average heat transfer coefficients are best correlated on the basis that the coefficient varies with [1 + (L/D))(sup -0,7)] and that the physical properties and density are evaluated at the surface temperature. The average friction coefficients for the tests with no heat addition are in complete agreement with the Karman-Nikuradse line. The average friction coefficients for heat addition are in poor agreement with the accepted line.

Taylor, Maynard F.↗
NASA NTRS1971

Liquid transfer demonstration on board Apollo 14 during transearth coast

The transfer of liquid from one container to another in a weightless environment was demonstrated by the crew of Apollo 14. A scale-model liquid-transfer system was used on board the spacecraft during the transearth coast period. The liquid transfer unit consisted of a surface tension baffled tank system containing two baffle designs. Liquid was transferred between tanks with a hand pump operated by the astronaut. The results showed that liquid was efficiently transferred to and from either baffled tank to within two percent of the design value residual liquid without reaching gas ingestion. The liquid-vapor interface in the receiver tank was positioned successfully with the gas located at the vent.

Abdalla, K. L.↗