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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 901 records · Page 50
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Three Dimensional Numerical Simulation of Rocket-based Combined-cycle Engine Response During Mode Transition Events

The GTX program at NASA Glenn Research Center is designed to develop a launch vehicle concept based on rocket-based combined-cycle (RBCC) propulsion. Experimental testing, cycle analysis, and computational fluid dynamics modeling have all demonstrated the viability of the GTX concept, yet significant technical issues and challenges still remain. Our research effort develops a unique capability for dynamic CFD simulation of complete high-speed propulsion devices and focuses this technology toward analysis of the GTX response during critical mode transition events. Our principal attention is focused on Mode 1/Mode 2 operation, in which initial rocket propulsion is transitioned into thermal-throat ramjet propulsion. A critical element of the GTX concept is the use of an Independent Ramjet Stream (IRS) cycle to provide propulsion at Mach numbers less than 3. In the IRS cycle, rocket thrust is initially used for primary power, and the hot rocket plume is used as a flame-holding mechanism for hydrogen fuel injected into the secondary air stream. A critical aspect is the establishment of a thermal throat in the secondary stream through the combination of area reduction effects and combustion-induced heat release. This is a necessity to enable the power-down of the rocket and the eventual shift to ramjet mode. Our focus in this first year of the grant has been in three areas, each progressing directly toward the key initial goal of simulating thermal throat formation during the IRS cycle: CFD algorithm development; simulation of Mode 1 experiments conducted at Glenn's Rig 1 facility; and IRS cycle simulations. The remainder of this report discusses each of these efforts in detail and presents a plan of work for the next year.

Edwards, Jack R.↗
NASA NTRSDate not supplied

Hot Hydrogen Testing of Uranium Nitride Cermet for Nuclear Thermal Propulsion

This study investigated the thermochemical stability and performance characteristics of uranium mononitride tungsten molybdenum cermet (UN-Mo30W) as a potential fuel material for Nuclear Thermal Propulsion (NTP) engines. Twelve monolithic cylinders were subjected to conditions simulating an active NTP engine. The tests demonstrated thermal variability within the test article and highlighted the dissociation of UN into U and N2 at elevated temperatures, affecting the microstructure and causing bonding between cermet wafers and cooling tubes. Findings indicate that while UN-Mo30W cermet exhibits some desirable properties for NTP applications, its suitability is limited by high-temperature thermochemical instability.

nuclear fuels↗
NASA NTRS2006

Thermal Simulator Development: Non-Nuclear Testing of Space Fission Systems

Non-nuclear testing can be a valuable tool in the development of a space nuclear power system. At the NASA MSFC Early Flight Fission Test Facility (EFF-TF), highly designed electric heaters are used to simulate the heat from nuclear fuel to test space fission power and propulsion systems. To allow early utilization, nuclear system designs must be relatively simple, easy to fabricate, and easy to test using non-nuclear heaters to closely mimic heat from fission. In this test strategy, highly designed electric heaters are used to simulate the heat from nuclear fuel, allowing one to develop a significant understanding of individual components and integrated system operation without the cost, time and safety concerns associated with nuclear testing.

Bragg-Sitton, Shannon M.↗
DOE OSTI2019

Analysis of the Detonation Wave Structure in a Linearized Rotating Detonation Engine

The rotating detonation engine is increasingly favored as a viable pressure gain combustion technology for both propulsion and power generation applications. Practical designs involve the discrete injection of fuel and air, which then partially mix to produce the reactive mixture that is processed by a continuously moving detonation wave within the detonation chamber. One of the fundamental challenges in the successful operation of rotating detonation engines is the design of a robust fuel-oxidizer injection system. To minimize pressure losses while ensuring efficient mixing that promotes stable detonation, the injection has to satisfy many different constraints. In this work, the impact of discrete injection on wave propagation is studied in an effort to understand the impact of fuel stratification on detonation stability. To this end, the structure of the detonation wave within a linearized rotating detonation engine with fully premixed hydrogen-air and hydrogen-oxygen injection at various equivalence ratios is analyzed. The direct numerical simulation of the linearized model detonation engine is compared with experimental results, and wave behavior is correlated to the wave structure and flow turbulence. As a result, the flow properties and chemical composition of the gases across the detonation wave are studied to examine the characteristics of the reaction zone.

42 ENGINEERING↗
NASA NTRS1992

Aerospace Test Facilities at NASA LeRC Plumbrook

An overview of the facilities and research being conducted at LeRC's Plumbrook field station is given. The video highlights four main structures and explains their uses. The Space Power Facility is the world's largest space environment simulation chamber, where spacebound hardware is tested in simulations of the vacuum and extreme heat and cold of the space plasma environment. This facility was used to prepare Atlas 1 rockets to ferry CRRES into orbit; it will also be used to test space nuclear electric power generation systems. The Spacecraft Propulsion Research Facility allows rocket vehicles to be hot fired in a simulated space environment. In the Cryogenic Propellant Tank Facility, researchers are developing technology for storing and transferring liquid hydrogen in space. There is also a Hypersonic Wind Tunnel which can perform flow tests with winds up to Mach 7.

Source record↗
NASA NTRS2011

Time Domain Tool Validation Using ARES I-X Flight Data

The ARES I-X vehicle was launched from NASA's Kennedy Space Center (KSC) on October 28, 2009 at approximately 11:30 EDT. ARES I-X was the first test flight for NASA s ARES I launch vehicle, and it was the first non-Shuttle launch vehicle designed and flown by NASA since Saturn. The ARES I-X had a 4-segment solid rocket booster (SRB) first stage and a dummy upper stage (US) to emulate the properties of the ARES I US. During ARES I-X pre-flight modeling and analysis, six (6) independent time domain simulation tools were developed and cross validated. Each tool represents an independent implementation of a common set of models and parameters in a different simulation framework and architecture. Post flight data and reconstructed models provide the means to validate a subset of the simulations against actual flight data and to assess the accuracy of pre-flight dispersion analysis. Post flight data consists of telemetered Operational Flight Instrumentation (OFI) data primarily focused on flight computer outputs and sensor measurements as well as Best Estimated Trajectory (BET) data that estimates vehicle state information from all available measurement sources. While pre-flight models were found to provide a reasonable prediction of the vehicle flight, reconstructed models were generated to better represent and simulate the ARES I-X flight. Post flight reconstructed models include: SRB propulsion model, thrust vector bias models, mass properties, base aerodynamics, and Meteorological Estimated Trajectory (wind and atmospheric data). The result of the effort is a set of independently developed, high fidelity, time-domain simulation tools that have been cross validated and validated against flight data. This paper presents the process and results of high fidelity aerospace modeling, simulation, analysis and tool validation in the time domain.

Hough, Steven↗
DOE OSTI2013

Electric Drive Dynamic Thermal System Model for Advanced Vehicle Propulsion Technologies: Cooperative Research and Development Final Report, CRADA Number CRD-09-360

Electric drive systems, which include electric machines and power electronics, are a key enabling technology for advanced vehicle propulsion systems that reduce the dependence of the U.S. transportation sector on petroleum. However, to penetrate the market, these electric drive technologies must enable vehicle solutions that are economically viable. The push to make critical electric drivesystems smaller, lighter, and more cost-effective brings respective challenges associated with heat removal and system efficiency. In addition, the wide application of electric drive systems to alternative propulsion technologies ranging from integrated starter generators, to hybrid electric vehicles, to full electric vehicles presents challenges in terms of sizing critical components andthermal management systems over a range of in-use operating conditions. This effort focused on developing a modular modeling methodology to enable multi-scale and multi-physics simulation capabilities leading to generic electric drive system models applicable to alternative vehicle propulsion configurations. The primary benefit for the National Renewable Energy Laboratory (NREL) is the abilityto define operating losses with the respective impact on component sizing, temperature, and thermal management at the component, subsystem, and system level. However, the flexible nature of the model also allows other uses related to evaluating the impacts of alternative component designs or control schemes depending on the interests of other parties.

33 ADVANCED PROPULSION SYSTEMS↗
DOE OSTI2023

Evaluation of Spray and Combustion Models for Simulating Dilute Combustion in a Direct-Injection Spark-Ignition Engine

Dilute combustion in spark-ignition engines has the potential to improve thermal efficiency by mitigating knock and by reducing throttling and wall heat losses. However, ignition and combustion processes can become unstable for dilute operation due to a lowered laminar flame speed, resulting in excessive cycle-to-cycle variability (CCV) of the combustion process. To compensate for the slower combustion in less reactive mixtures, a modified intake port geometry can be employed to generate a strong tumble flow in the cylinder and elevate turbulence levels around the spark plug, thereby promoting a faster transition to turbulent deflagration. Consequently, optimizing combustion chamber geometry and operating strategy is crucial to maximizing the benefits of using dilute combustion with enhanced in-cylinder turbulence across a wide range of operating conditions. Computational fluid dynamics (CFD) simulations can be utilized for virtual engine optimization tasks, but this would require the models to be truly predictive regarding the impact of changes to the engine design and operational parameters.In this study, multicycle large-eddy simulations (LES) are performed for a direct-injection spark-ignition engine to investigate the model performance in predicting engine combustion characteristics with respect to changes in the intake configuration. A tumble plate that blocks the lower part of the intake port inlet is used to vary the tumble. A set of CFD models that have been recently developed are employed, which takes into account the drag of nonspherical droplets, flash-boiling behavior of liquid sprays, spray-wall interaction, surrogate formulation of a research-grade E10 gasoline, and fast chemical kinetic solvers. Simulation results are compared to experimental engine data in terms of cylinder pressure, apparent heat release rate, mass fraction burned timing, and flame images. It is found that LES employing the state-of-the-art CFD models are capable of properly predicting the spray processes and reproducing the measured mean cylinder pressure for the case with the tumble plate. On the other hand, the LES over-predicts the combustion rate during the early combustion stage and under-estimates the CCV, and these discrepancies become larger when the tumble plate is removed.

computational fluid dynamics simulation↗
NASA NTRS2014

Plasmonic Force Propulsion Revolutionizes Nano/Picosatellite Capability

This project investigated a new type of small spacecraft propulsion for attitude control, specifically proximity and precision pointing control. Plasmonic force propulsion uses solar light focused on deep-subwavelength nanostructures to excite strong optical forces that accelerate and expel nanoparticle propellant. The goal of the project was to assess the feasibility of plasmonic force propulsion for nano/pico-satellite applications by evaluating key mission parameters for a nano/pico-satellite using plasmonic force propulsion in a NASA-relevant mission context. We achieved this goal and objective by evaluating plasmonic force propulsion within a NASA mission that required attitude control and precision pointing of a small satellite. We numerically simulated plasmonic force fields with asymmetric/gradient geometry and relevant solar light constraints, predicted nanoparticle velocity, mass flow rate, and resulting propulsion performance (thrust, specific impulse), and evaluated spacecraft position control resolution and pointing precision. Additionally we compared the precision pointing capabilities of plasmonic propulsion, as well as the mass, volume, and power requirements, with other state-of-the-art control techniques, such as reaction wheels and colloid/electrospray electric propulsion. The results are very exciting. Plasmonic force propulsion can significantly enhance the state-of-the-art in small spacecraft position and attitude control by 1-2 orders of magnitude. This is most succinctly shown in the figure below, which compares proximity and attitude control of plasmonic force propulsion (PFP) with other state-of-the-art thruster systems (μCAT, VAT, electrospray). Additionally this figure also shows the proximity and attitude control required for different existing (James Webb Space Telescope, Hubble) and future (LISA and Stellar Imager) NASA missions. While some of these NASA missions are not small spacecraft missions, the requirements serve to illustrate the fact that more precise proximity and attitude control will be required for future NASA science missions. Stellar imager is a proposed NASA missions that requires an extremely high pointing precision of 0.1 milliarcseconds (2.7x10(exp -7) deg.) for an ultraviolet telescope that has over 200× the resolution of the Hubble Space Telescope, is able to take images showing details on the surfaces of other stars, consists of 20-30 small "mirror sats" flying in formation to produce a giant mirror, and requires each mirror-sat to be placed with nanometer precision and control its attitude with milliarcsecond precision.

Rovey, Joshua L.↗
NASA NTRS2003

Propulsion Options for the Global Precipitation Measurement Core Satellite

This study was conducted to evaluate several propulsion system options for the Global Precipitation Measurement (GPM) core satellite. Orbital simulations showed clear benefits for the scientific data to be obtained at a constant orbital altitude rather than with a decay/reboost approach. An orbital analysis estimated the drag force on the satellite will be 1 to 12 mN during the five-year mission. Four electric propulsion systems were identified that are able to compensate for these drag forces and maintain a circular orbit. The four systems were the UK-10/TS and the NASA 8 cm ion engines, and the ESA RMT and RITl0 EVO radio-frequency ion engines. The mass, cost, and power requirements were examined for these four systems. The systems were also evaluated for the transfer time from the initial orbit of 400 x 650 km altitude orbit to a circular 400 km orbit. The transfer times were excessive, and as a consequence a dual system concept (with a hydrazine monopropellant system for the orbit transfer and electric propulsion for drag compensation) was examined. Clear mass benefits were obtained with the dual system, but cost remains an issue because of the larger power system required for the electric propulsion system. An electrodynamic tether was also evaluated in this trade study.

Cardiff, Eric H.↗
NASA NTRSDate not supplied

Computational Investigation of Retropropulsion Operating Environments with a GPU-Enabled Detached Eddy Simulation Approach

Human exploration of the surface of Mars will require an extended powered descent phase of flight, during which aerodynamic-propulsive interference effects can be significant. Characterization of these environments to enable implementation of this technology into a flight vehicle will rely heavily on computational simulation. This work advances the understanding of retropropulsion aerodynamics through application of a massively parallel detached eddy simulation approach on a GPU-accelerated computational framework, yielding data that are largely unachievable with conventional high-performance computing resources. This work includes time-dependent and time-averaged forces and moments on a conceptual, full-scale vehicle in environments and operating conditions relevant to human Mars exploration. Conditions are examined where the engine exhaust flow transitions between over-expanded and under-expanded flow structures, and flight operation will require the ability to maintain control of the vehicle during such a transition. These transitions occur as the vehicle decelerates, and as such, this investigation includes supersonic, transonic, and subsonic flight conditions. Options for vehicle control during powered flight include differential throttling of the engines. This paper provides an overview of the computational campaign, approach, and discussion of results in characterizing the resulting aerodynamics for differential throttling with retropropulsion in atmospheric environments.

EDL↗
NASA NTRS2006

Test Capability Enhancements to the NASA Langley 8-Foot High Temperature Tunnel

The NASA Langley 8-Foot High Temperature Tunnel produces true enthalpy environments simulating flight from Mach 4 to Mach 7, primarily for airbreathing propulsion and aerothermal/thermo-structural testing. Flow conditions are achieved through a methane-air heater and nozzles producing aerodynamic Mach numbers of 4, 5 or 7 and have exit diameters of 8 feet or 4.5 feet. The 12-ft long free-jet test section, housed inside a 26-ft vacuum sphere, accommodates large test articles. Recently, the facility underwent significant upgrades to support hydrocarbon fueled scramjet engine testing and to expand flight simulation capability. The upgrades were required to meet engine system development and flight clearance verification requirements originally defined by the joint NASA-Air Force X-43C Hypersonic Flight Demonstrator Project and now the Air Force X-51A Program. Enhancements to the 8-Ft. HTT were made in four areas: 1) hydrocarbon fuel delivery; 2) flight simulation capability; 3) controls and communication; and 4) data acquisition/processing. The upgrades include the addition of systems to supply ethylene and liquid JP-7 to test articles; a Mach 5 nozzle with dynamic pressure simulation capability up to 3200 psf, the addition of a real-time model angle-of-attack system; a new programmable logic controller sub-system to improve process controls and communication with model controls; the addition of MIL-STD-1553B and high speed data acquisition systems and a classified data processing environment. These additions represent a significant increase to the already unique test capability and flexibility of the facility, and complement the existing array of test support hardware such as a model injection system, radiant heaters, six-component force measurement system, and optical flow field visualization hardware. The new systems support complex test programs that require sophisticated test sequences and precise management of process fluids. Furthermore, the new systems, such as the real-time angle of attack system and the new programmable logic controller enhance the test efficiency of the facility. The motivation for the upgrades and the expanded capabilities is described here.

Harvin, S. F.↗
DOE OSTI2023

Temperature Rise on Liner Surfaces of Fuel Cell Electric Vehicle Tanks during Fueling Process

Herein, the temperature of the inner tank walls, or plastic liners, of composite pressure tanks in fuel cell electric vehicles during fueling using three‐dimensional computational fluid dynamics (CFD) models is evaluated. The liner materials are limited to 85.0 °C to prevent thermal stress causing material failure that would result in a hydrogen leak. Therefore, the temperature of hydrogen gas must be limited to below 85.0 °C during the fueling process as dictated by the current fueling protocol. However, there are limited experimental or simulation data confirming that the temperature changes do not exceed the threshold. Herein, the liner temperatures with CFD tank models for two sizes of type IV tanks representative of the upper and lower system bounds that are close to the SAE J2601 fueling protocol (36.0 and 244.0 L) are evaluated. First, each model's reliability is validated with experimental data and then analyzed, and the data are used to evaluate the maximum hydrogen and liner temperatures under real‐world fueling conditions. The evaluation shows that the maximum liner surface temperature of each tank model is at least 7 °C lower than that of the hydrogen. Additionally, there is at least 12 °C difference found between the upper limit and actual liner temperatures.

33 ADVANCED PROPULSION SYSTEMS↗
DOE OSTI2025

Chemical timescale effects on detonation convergence

Numerical simulations of detonation-containing flows have emerged as crucial tools for designing next- generation power and propulsion devices. As these tools mature, it is important for the combustion community to properly understand and isolate grid resolution effects when simulating detonations. To this end, the objective of this work is to provide a comprehensive analysis of the numerical convergence of unsteady detonation simulations, with focus on isolating the impacts of chemical timescale modifications on convergence characteristics in the context of operator splitting. With the aid of an AMReX-based adaptive mesh refinement flow solver-which enables resolutions up to ($\mathcal{O}$ (1000) cells-per-induction length-the convergence analysis is conducted using two kinetics configurations: (1) the simplified three-step Arrhenius-based model mechanism of Short and Quirk (1997), where chemical timescales in the detonation are modified by adjusting activation energies in the initiation and branching reactions, and (2) a detailed hydrogen- air mechanism, where the chemical timescales are adjusted by varying the ambient pressure. The convergence of unsteady self-sustained detonations in one-dimensional channels is then analyzed with reference to steady-state theoretical baseline solutions using these mechanisms. The goal of the analysis is to provide a detailed comparison of the effects of grid resolution on both macroscopic (peak pressures and wave speeds) and microscopic (wave structure) quantities of interest, drawing connections between the deviations from steady-state baselines and minimum chemical timescales. In particular, chemical timescale reductions were found to have minimal impact on the convergence of macroscopic properties. Furthermore, analyses of microscopic convergence trends, particularly in the reaction front location, revealed a key insight: maintaining the induction time while eliminating prohibitive chemical timescales through mechanism simplifications and combustion modeling can significantly enhance detonation convergence properties. Ultimately, this work uncovers resolution-dependent unsteady detonation convergence regimes and highlights the important role played by not only the chemical timescales, but also the ratio between the chemical timescale and induction time on the numerical convergence of the detonation wave structure.

Adaptive mesh refinement↗
NASA NTRS1992

High spatial resolution measurements in a single stage ram accelerator

High spatial resolution experimental tube wall pressure measurements of ram accelerator gas dynamic phenomena are presented in this paper. The ram accelerator is a ramjet-in-tube device which operates in a manner similar to that of a conventional ramjet. The projectile resembles the centerbody of a ramjet and travels supersonically through a tube filled with a combustible gaseous mixture, with the tube acting as the outer cowling. Pressure data are recorded as the projectile passes by sensors mounted in the tube wall at various locations along the tube. Utilization of special highly instrumented sections of tube has allowed the recording of gas dynamic phenomena with high resolution. High spatial resolution tube wall pressure data from the three regimes of propulsion studied to date (subdetonative, transdetonative, and superdetonative) in a single stage gas mixture are presented and reveal the three-dimensional character of the flow field induced by projectile fins and the canting of the fins and the canting of the projectile body relative to the tube wall. Also presented for comparison to the experimental data are calculations made with an inviscid, three-dimensional CFD code. The knowledge gained from these experiments and simulations is useful in understanding the underlying nature of ram accelerator propulsive regimes, as well as assisting in the validation of three-dimensional CFD coded which model unsteady, chemically reactive flows.

Hinkey, J. B.↗
NASA NTRS2004

Mars Exploration Rovers navigation results

The twin Mars Exploration Rovers, Spirit and Opportunity, were launched on June 10, 2003, and July 8, 2003, from Cape Canaveral, Florida. Spirit and Opportunity were targeted for landings at Gusev Crater (arrival on January 4, 2004) and Meridiani Planum (arrival on January 25, 2004). The primary navigation challenge was to deliver each spacecraft to the desired atmospheric entry interface point with sufficient accuracy such that each lander would touch down within a specified landing ellipse (about 70 km x 5 km) determined to be safe for landing and also judged to be scientifically interesting. In order to achieve landing within the target ellipse, precise control of the inertial entry flight path angle (FPA) at atmospheric entry was required. The maximum allowable errors in FPA following TCM-5 (trajectory correction maneuver #5) at Entry (E) -2 days were +/-0.12 deg(3 sigma) for Spirit and +/-0.14 deg(3 sigma) for Opportunity. Achieving these entry delivery accuracies necessitated significant improvements to the interplanetary navigation system used for MER. These improvements included new processes and software for orbit determination, propulsive maneuver design, and entry, descent, and landing (EDL) trajectory simulation. The actual achieved atmospheric entry accuracies for Spirit and Opportunity significantly exceeded the requirements. At the navigation data cutoff for the TCM-5 final design, the orbit determination FPA knowledge error was 0.028 deg(3 sigma) for Spirit and 0.035 deg(3 sigma) for Opportunity. Because of exceptionally accurate navigation performance, TCM-5 (E - 2 days) and TCM-6 (E - 4 hours) were canceled for both Spirit and Opportunity. The actual landing locations (determined from in-situ Doppler tracking between the MER rovers and the Mars Odyssey orbiter) differed from the target landing points by 10.1 km (downtrack) for Spirit and 24.6 km (downtrack) for Opportunity. The majority of the landing position offsets for both landers was primarily caused by variations in atmosphere and spacecraft aerodynamic modeling from what was predicted. The amount of the landing position offset caused by navigation-only errors was only 3.3 km (uptrack) for Spirit and 9.7 km (downtrack) for Opportunity.

Mars Exploration Rover (MER)↗
DOE OSTI2021

Dynamic Ride-Matching for Large-Scale Transportation Systems

Efficient dynamic ride-matching (DRM) in large-scale transportation systems is a key driver in transport simulations to yield answers to challenging problems. Although the DRM problem is simple to solve, it quickly becomes a computationally challenging problem in large-scale transportation system simulations. Therefore, this study thoroughly examines the DRM problem dynamics and proposes an optimization-based solution framework to solve the problem efficiently. To benefit from parallel computing and reduce computational times, the problem’s network is divided into clusters utilizing a commonly used unsupervised machine learning algorithm along with a linear programming model. Then, these sub-problems are solved using another linear program to finalize the ride-matching. At the clustering level, the framework allows users adjusting cluster sizes to balance the trade-off between the computational time savings and the solution quality deviation. A case study in the Chicago Metropolitan Area, U.S., illustrates that the framework can reduce the average computational time by 58% at the cost of increasing the average pick up time by 26% compared with a system optimum, that is, non-clustered, approach. Another case study in a relatively small city, Bloomington, Illinois, U.S., shows that the framework provides quite similar results to the system-optimum approach in approximately 62% less computational time.

33 ADVANCED PROPULSION SYSTEMS↗