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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 973 records · Page 54
All sourcesNASA NTRSDOE OSTI
DOE OSTI2022

Decarbonizing Medium- & Heavy-Duty On-Road Vehicles: Zero-Emission Vehicles Cost Analysis

We simulate adoption and energy consumption of zero-emission vehicles (ZEVs) in the medium and heavy duty (MD/HD) sector. With continued improvement in ZEV technologies, total-cost-of-driving parity with conventional vehicles is achievable by 2035 for all MD/HD vehicle classes. Two technological solutions – battery electric vehicles (BEVs) and fuel cell electric vehicles (FCEVs) are viable in multiple market segments, offering alternative pathways for decarbonization. In this scenario, ZEV sales reach 99+% by 2045 and 42% by 2030. 80% of the MD/HD stock transitions to ZEVs by 2050, reducing CO2 emissions by 69% from 2019. Disparities in adoption rate and vehicle usage between vehicle classes and between short and long-haul applications highlight opportunities for targeted policies (to reduce emissions it is critical to understand what vehicles transition to ZEV, not just how many vehicles). Results are highly sensitive to assumed fuel prices, technology improvement trajectories, adoption decision-making, and assumptions about future freight demand and vehicle use.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1979

Inert gas thrusters

Inert gas thrusters considered for space propulsion systems were investigated. Electron diffusion across a magnetic field was examined utilizing a basic model. The production of doubly charged ions was correlated using only overall performance parameters. The use of this correlation is therefore possible in the design stage of large gas thrusters, where detailed plasma properties are not available. Argon hollow cathode performance was investigated over a range of emission currents, with the positions of the inert, keeper, and anode varied. A general trend observed was that the maximum ratio of emission to flow rate increased at higher propellant flow rates. It was also found that an enclosed keeper enhances maximum cathode emission at high flow rates. The maximum cathode emission at a given flow rate was associated with a noisy high voltage mode. Although this mode has some similarities to the plume mode found at low flows and emissions, it is encountered by being initially in the spot mode and increasing emission. A detailed analysis of large, inert-gas thruster performance was carried out. For maximum thruster efficiency, the optimum beam diameter increases from less than a meter at under 2000 sec specific impulse to several meters at 10,000 sec. The corresponding range in input power ranges from several kilowatts to megawatts.

Kaufman, H. R.↗
NASA NTRS1981

NASA progress in aircraft noise prediction

Langley Research Center efforts to develop a methodology for predicting the effective perceived noise level (EPNL) produced by jet-powered CTOL aircraft to an accuracy of + or - 1.5 dB are summarized with emphasis on the aircraft noise prediction program (ANOPP) which contains a complete set of prediction methods for CTOL aircraft including propulsion system noise sources, aerodynamic or airframe noise sources, forward speed effects, a layered atmospheric model with molecular absorption, ground impedance effects including excess ground attenuation, and a received noise contouring capability. The present state of ANOPP is described and its accuracy and applicability to the preliminary aircraft design process is assessed. Areas are indicated where further theoretical and experimental research on noise prediction are needed. Topics covered include the elements of the noise prediction problem which are incorporated in ANOPP, results of comparisons of ANOPP calculations with measured noise levels, and progress toward treating noise as a design constraint in aircraft system studies.

Raney, J. P.↗
NASA NTRS1986

A guidance-motivated sensitivity analysis of an aero-assisted boost vehicle

A simple model of an aero-assisted booster is used to examine the contributions of propulsion system type, aerodynamic lift and flight trajectory to the efficiency with which payloads can be placed into low earth orbit. The higher propulsive efficiency of ramjet and scramjet propulsion has the potential of increasing the payload mass ratio significantly. The contributions of turbojet propulsion and aerodynamic lift are less significant. The additional complexity involved in using aerodynamic propulsion and lift requires dealing with a more comprehensive set of design variables than for rocket boosters. The approach taken is to derive a set of sensitivity functions which relate booster performance to the design variables. The problems of optimum mixing of aerodynamic lift with thrust and determining the optimal boost trajectory is treated. The potential payload capacity of a horizontal take-off air-breathing boost vehicle is examined. The optimization problem which considers propulsive efficiency, aerodynamic configuration, and control and guidance issues is discussed.

Taylor, L. W.↗
NASA NTRS1992

Exhaust plumes and their interaction with missile airframes - A new viewpoint

The present, novel treatment of missile airframe-exhaust plume interactions emphasizes their simulation via a formal solution of the Reynolds-averaged Navier-Stokes (RNS) equation and is accordingly able to address the simulation requirements of novel missiles with nonconventional/integrated propulsion systems. The method is made possible by implicit RNS codes with improved artificial dissipation models, generalized geometric capabilities, and improved two-equation turbulence models, as well as by such codes' recent incorporation of plume thermochemistry and multiphase flow effects.

Dash, S. M.↗
NASA NTRS1991

Mixing and mixing enhancement in supersonic reacting flowfields

Propulsion systems planned for use late in this century and beyond will require appropriate physical models for describing supersonic combustion and numerical techniques for solving the model governing equations. A computer program to study these flows is reported which considers the multicomponent diffusion and convection of important chemical species, the finite-rate reaction of these species, and the resulting interaction of the field mechanics and the chemistry. The application of the program to a spatially developing and reacting mixing layer, which serves an an excellent physical model for the mixing and reaction processes that take place in a scramjet combustor, is reported. Several techniques to enhance the fuel-air mixing and growth of that layer and improve its overall combustion efficiency are considered.

Drummond, J. P.↗
NASA NTRS1996

Waterjet Impeller

North American Marine Jet (NAMJ), Inc. received assistance from Marshall Space Flight Center engineers in the Computational Fluid Dynamics (CFD) branch of the Structure and Dynamics Laboratory in improving the proposed design of a new impeller for their jet-propulsion systems. Marshall used advanced CFD techniques, which included creating a three-dimensional computer model of the impeller for analysis. With Marshall input, the company modified the design, then Marshall used a computer model to make a solid polycarbonate model. The rapid prototyping allowed the company to avoid many time- consuming and costly steps in creating the impeller model. NAMJ is now able to compete with Pacific-area and European manufacturers who have traditionally dominated the market.

Source record↗
NASA NTRS2002

20th JANNAF Propulsion Systems Hazards Subcommittee Meeting

This volume, the first of two volumes, is a collection of 24 unclassified/unlimited-distribution papers which were presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 20th Propulsion Systems Hazards Subcommittee (PSHS), 38th Combustion Subcommittee (CS), 26th Airbreathing Propulsion Subcommittee (APS), and 21 Modeling and Simulation Subcommittee meeting. The meeting was held 8-12 April 2002 at the Bayside Inn at The Sandestin Golf & Beach Resort and Eglin Air Force Base, Destin, Florida. Topics covered include: insensitive munitions and hazard classification testing of solid rocket motors and other munitions; vulnerability of gun propellants to impact stimuli; thermal decomposition and cookoff properties of energetic materials; burn-to-violent reaction phenomena in energetic materials; and shock-to-detonation properties of solid propellants and energetic materials.

Cocchiaro, James E.↗
NASA NTRS2004

Plasma Sail Concept Fundamentals

The mini-magnetospheric plasma propulsion (M2P2) device, originally proposed by Winglee et al., predicts that a 15-km standoff distance (or 20-km cross-sectional dimension) of the magnetic bubble will provide for sufficient momentum transfer from the solar wind to accelerate a spacecraft to unprecedented speeds of 50 C80 km/s after an acceleration period of 3 mo. Such velocities will enable travel out of the solar system in period of 7 yr almost an order of magnitude improvement over present chemical-based propulsion systems. However, for the parameters of the simulation of Winglee et al., a fluid model for the interaction of M2P2 with the solar wind is not valid. It is assumed in the magnetohydrodynamic (MHD) fluid model, normally applied to planetary magnetospheres, that the characteristic scale size is much greater than the Larmor radius and ion skin depth of the solar wind. In the case of M2P2, the size of the magnetic bubble is actually less than or comparable to the scale of these characteristic parameters. Therefore, a kinetic approach, which addresses the small-scale physical mechanisms, must be used. A two-component approach to determining a preliminary estimate of the momentum transfer to the plasma sail has been adopted. The first component is a self-consistent MHD simulation of the small-scale expansion phase of the magnetic bubble. The fluid treatment is valid to roughly 5 km from the source and the steady-state MHD solution at the 5 km boundary was then used as initial conditions for the hybrid simulation. The hybrid simulations showed that the forces delivered to the innermost regions of the plasma sail are considerably ( 10 times) smaller than the MHD counterpart, are dominated by the magnetic field pressure gradient, and are directed primarily in the transverse direction.

Khazanov, G. V.↗
NASA NTRS2009

Performance Characteristics of the NEXT Long-Duration Test After 16,550 h and 337 kg of Xenon Processed

The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated to verify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the anticipated throughput requirement of 300 kg from mission analyses conducted utilizing the NEXT propulsion system. The LDT is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of June 25, 2008, the thruster has accumulated 16,550 h of operation: the first 13,042 h at the thruster full-input-power of 6.9 kW with 3.52 A beam current and 1800 V beam power supply voltage. Operation since 13,042 h, i.e., the most recent 3,508 h, has been at an input power of 4.7 kW with 3.52 A beam current and 1180 V beam power supply voltage. The thruster has processed 337 kg of xenon (Xe) surpassing the NSTAR propellant throughput demonstrated during the extended life testing of the Deep Space 1 flight spare ion thruster. The NEXT LDT has demonstrated a total impulse of 13.3 106 N s; the highest total impulse ever demonstrated by an ion thruster. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. This paper presents the performance of the NEXT LDT to date with emphasis on performance variations following throttling of the thruster to the new operating condition and comparison of performance to the NSTAR extended life test.

Soulas, George C.↗
NASA NTRS1996

Development of an Intelligent Monitoring and Control System for a Heterogeneous Numerical Propulsion System Simulation

The NASA Numerical Propulsion System Simulation (NPSS) project is exploring the use of computer simulation to facilitate the design of new jet engines. Several key issues raised in this research are being examined in an NPSS-related research project: zooming, monitoring and control, and support for heterogeneity. The design of a simulation executive that addresses each of these issues is described. In this work, the strategy of zooming, which allows codes that model at different levels of fidelity to be integrated within a single simulation, is applied to the fan component of a turbofan propulsion system. A prototype monitoring and control system has been designed for this simulation to support experimentation with expert system techniques for active control of the simulation. An interconnection system provides a transparent means of connecting the heterogeneous systems that comprise the prototype.

Reed, John A.↗
NASA NTRS2016

Benchmarking Model Variants in Development of a Hardware-in-the-Loop Simulation System

Distributed engine control architecture presents a significant increase in complexity over traditional implementations when viewed from the perspective of system simulation and hardware design and test. Even if the overall function of the control scheme remains the same, the hardware implementation can have a significant effect on the overall system performance due to differences in the creation and flow of data between control elements. A Hardware-in-the-Loop (HIL) simulation system is under development at NASA Glenn Research Center that enables the exploration of these hardware dependent issues. The system is based on, but not limited to, the Commercial Modular Aero-Propulsion System Simulation 40k (C-MAPSS40k). This paper describes the step-by-step conversion from the self-contained baseline model to the hardware in the loop model, and the validation of each step. As the control model hardware fidelity was improved during HIL system development, benchmarking simulations were performed to verify that engine system performance characteristics remained the same. The results demonstrate the goal of the effort; the new HIL configurations have similar functionality and performance compared to the baseline C-MAPSS40k system.

Networked Systems↗
DOE OSTI2024

Evaluating the Impacts of Autonomous Electric Vehicles Adoption on Vehicle Miles Traveled and CO2 Emissions

Autonomous electric vehicles (AEVs) can potentially revolutionize the transportation landscape, offering a safer, contact-free, easily accessible, and more eco-friendly mode of travel. Prior to the market uptake of AEVs, it is critical to understand the consumer segments that are most likely to adopt these vehicles. Beyond market adoption, it is also important to quantify the impact of AEVs on broader transportation systems and the environment, such as impacts on the annual vehicle miles traveled (VMT) and greenhouse gas (GHG) emissions. In this pilot study, using survey data, a statistical model correlating AEV adoption intention and socioeconomic and built environment attributes was estimated, and a sensitivity analysis was conducted to understand the importance of factors impacting AEV adoption. We found that the market segments range from early adopters who are wealthy, technologically savvy, and relatively young to non-adopters who are more cautious to new technologies. This is followed by a synthetic population microsimulation of market penetration for the San Francisco Bay Area. With five household vehicle replacement scenarios, we assessed the annual VMT and tailpipe carbon dioxide (CO2) emissions change associated with vehicle replacement. It is found that adopting AEVs can potentially reduce more than 5 megatons of CO2 yearly, which is approximately 30% of the total CO2 emitted by internal combustion engine (ICE) cars in the region.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS2005

The Importance of Detailed Component Simulations in the Feedsystem Development for a Two-Stage-to Orbit Reusable Launch Vehicle

To meet the requirements for the 2nd Generation Reusable Launch Vehicle (RLV), a unique propulsion feed system concept was identified using crossfeed between the booster and orbiter stages that could reduce the Two-Stage-to-Orbit (TSTO) vehicle weight and development cost by approximately 25%. A Main Propulsion System (MPS) crossfeed water demonstration test program was configured to address all the activities required to reduce the risks for the MPS crossfeed system. A transient, one-dimensional system simulation was developed for the subscale crossfeed water flow tests. To ensure accurate representation of the crossfeed valve's dynamics in the system model, a high-fidelity, three-dimensional, computational fluid-dynamics (CFD) model was employed. The results from the CFD model were used to specify the valve's flow characteristics in the system simulation. This yielded a crossfeed system model that was anchored to the specific valve hardware and achieved good agreement with the measured test data. These results allowed the transient models to be correlated and validated and used for full scale mission predictions. The full scale model simulations indicate crossfeed is ' viable with the system pressure disturbances at the crossfeed transition being less than experienced by the propulsion system during engine start and shutdown transients.

Mazurkivich, Pete↗
NASA NTRS2005

Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

Pulse detonation rocket engines (PDREs) have generated research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred considerable interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models. In recent work by the author, a quasi-one-dimensional, finite rate chemistry CFD model was utilized to study the gasdynamics and performance characteristics of PDREs over a range of blowdown pressure ratios from 1-1000. Models of this type are computationally inexpensive, and enable first-order parametric studies of the effect of several nozzle and extension geometries on PDRE performance over a wide range of conditions. However, the quasi-one-dimensional approach is limited in that it cannot properly capture the multidimensional blast wave and flow expansion downstream of the PDRE, nor can it resolve nozzle flow separation if present. Moreover, the previous work was limited to single-pulse calculations. In this paper, an axisymmetric finite rate chemistry model is described and utilized to study these issues in greater detail. Example Mach number contour plots showing the multidimensional blast wave and nozzle exhaust plume are shown. The performance results are compared with the quasi-one-dimensional results from the previous paper. Both Euler and Navier-Stokes solutions are calculated in order to determine the effect of viscous effects in the nozzle flowfield. Additionally, comparisons of the model results to performance data from CalTech, as well as experimental flowfield measurements from Stanford University, are also reported.

Morris, Christopher I.↗
NASA NTRS1976

Simulation of an ion thruster control system

The results of an initial effort to model the control loops of a 30-cm diameter electron bombardment thruster and a transistorized power processor predicting its operation were described. Data from which the model is made is presented as well as comparisons between the computer outputs and test data from the JPL Solar Electric Propulsion systems laboratory.

Kudo, I.↗