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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 1,531 records · Page 85
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Ignition and combustion characteristics of metallized propellants

Shakedown and calibration of the experimental apparatus designed to study secondary atomization and ignition characteristics of aluminum slurry propellants was begun and nears completion. The burner was tested over an array of equivalence ratios and oxygen mass fractions to determine the permissible range of operating conditions. A burner flow control system was designed and fabricated which permits independent control of flame stoichiometry and oxidant O2 concentrations. The Sun slurry formulation was selected as the base for initial testing. Laser beam waists at the focal volume of the single particle sizing system were determined to be 396 microns for the Ar-ion and 81 microns for the He-Ne laser. Size calibrations using pinholes were completed and compared to Mie theory for light scattering from particles. A vibrating office droplet generator was used to test the single particle sizing system dynamically by producing a water droplet laden gas flow. The data acquisition and analysis programming was completed and testing is in progress. From initial results, the particle sizing, atomization and ignition detection systems appear to be performing as expected.

Turns, S. R.↗
NASA NTRS1990

Exploratory wind tunnel investigation of the stability and control characteristics of a three-surface, forward-swept wing advanced turboprop model

The purpose of the present investigation was to parametrically study the stability and control characteristics of a forward-swept wing three-surface turboprop model through an extended angle of attack range, including the deep-stall region. As part of a joint research program between North Carolina State University and NASA Langley Research Center, a low-speed wind tunnel investigation was conducted with a three-surface, forward-swept wing, aft-mounted, twin-pusher propeller, model, representative of an advanced turboprop configuration. The tests were conducted in the NASA Langley 12-Foot Low-Speed Wind Tunnel. The model parameters varied in the test were horizontal tail location, canard size, sweep and location, and wing position. The model was equipped with air turbines, housed within the nacelles and driven by compressed air, to model turboprop power effects. A three-surface, forward-swept wing configuration that provided satisfactory static longitudinal and lateral/directional stability was identified. The three-surface configuration was found to have greater longitudinal control and increased center of gravity range relative to a conventional (two-surface) design. The test showed that power had a large favorable effect on stability and control about all three axis in the post-stall regime.

Coe, Paul L., Jr.↗
NASA NTRS1993

Laser velocimeter measurements of the flow field generated by a forward-swept propfan during flutter

Results are presented from an investigation to measure the flow field generated by a forward-swept propfan operating in flutter at a low forward velocity. For comparison to the flutter condition, flow field data are also presented for a slightly reduced rotational speed just below flutter. The forward-swept propfan was tested as the front rotor of a counterrotating pusher propeller. A laser Doppler velocimeter (LDV) was used to measure the velocity field in planes normal to the model centerline downstream of the rotor and in planes of constant radius within the blade passages at each operating condition. The mean time-averaged flow about the blades did not change drastically as the propfan rotational speed was increased from the stable operating point to the flutter condition. No regions of flow separation could be identified in the data plots of the mean intrablade flow field. The relative flow about the blades remained subsonic during flutter operation. The blades were found to have a higher than expected tip loading at both operating conditions. This is thought to have been caused by the outer blade sections twisting under load to higher than expected effective blade angles. This high tip loading resulted in strong vortices and a very nonuniform flow downstream of the tips of the forward-swept blades. This high tip loading may also have caused the blade flutter.

Podboy, Gary G.↗
NASA NTRS1993

Laser velocimeter measurements of the flow field generated by a forward-swept propfan during flutter

Results are presented from an investigation to measure the flow field generated by a forward-swept propfan operating in flutter at a low forward velocity. For comparison to the flutter condition, flow field data are also presented for a slightly reduced rotational speed just below flutter. The forward-swept propfan was tested as the front rotor of a counterrotating pusher propeller. A laser Doppler velocimeter (LDV) was used to measure the velocity field in planes normal to the model centerline downstream of the rotor and in planes of constant radius within the blade passages at each operating condition. A comparison of the data taken at the two different operating conditions indicated that the mean, time-averaged flow about the blades did not change drastically as the propfan rotational speed was increased from the stable operating point to the flutter condition. No regions of flow separation could be identified in the data plots of the mean intrablade flow field. The data also indicate that the relative flow about the blades remained subsonic during flutter operation. The blades were found to have a higher than expected tip loading at both operating conditions. This is thought to have been caused by the outer blade sections twisting under load to higher than expected effective blade angles. This high tip loading resulted in strong vortices and a very nonuniform flow downstream of the tips of the forward-swept blades. This high tip loading may also have caused the blade flutter.

Podboy, Gary G.↗
NASA NTRS1999

Propulsive Small Expendable Deployer System (ProSEDS) Experiment

ProSEDS is a secondary (i.e. piggyback) payload on a Delta-11 GPS 8 mission scheduled for launch in August 2000. It will test the feasibility of generating generate electrodynamic thrust without propellant using a 5 kilometer conducting wire (tether). The ProSEDS obtains thrust as the tether cuts across the magnetic field, a voltage is induced across the wire. Electrons are attracted to the positively based far end of the wire. Electrons flow downward through the conductive tether. Earth's magnetic field exerts a drag force on the current in the tether segments, that is mechanically transferred via the wire to the stage. The primary objective for the ProSEDs mission is to demonstrate that a significant, measurable electrodynamic thrust through a tether in space. The primary mission will last one day, as the primary battery assures at least three orbits of data will be collected, the remaining power will be provided by the secondary battery, which uses tether generated power to recharge. The extended mission begins using the power provided through the tether, and wil terminate when a system ceases to function; (i.e., either degradation of the tether,through Atomic Oxygen contact, a micrometeoroid or other debris impact, or another malfunction.) The technology has many potential applications. Amongst the applications, which are reviewed in detail, are: (1) satellite deorbit, (2) reboost of the International Space Station, (3) propellantless reusable Orbit Transfer Vehicles, (4) Propulsion and power generation for future Jovian missions.

Johnson, Les↗
NASA NTRS2000

Development and Application of Computational/In Vitro Toxicological Methods for Chemical Hazard Risk Reduction of New Materials for Advanced Weapon Systems

The development of quantitative structure-activity relationship (QSAR) is essential for reducing the chemical hazards of new weapon systems. The current collaboration between HEST (toxicology research and testing), MLPJ (computational chemistry) and PRS (computational chemistry, new propellant synthesis) is focusing R&D efforts on basic research goals that will rapidly transition to useful products for propellant development. Computational methods are being investigated that will assist in forecasting cellular toxicological end-points. Models developed from these chemical structure-toxicity relationships are useful for the prediction of the toxicological endpoints of new related compounds. Research is focusing on the evaluation tools to be used for the discovery of such relationships and the development of models of the mechanisms of action. Combinations of computational chemistry techniques, in vitro toxicity methods, and statistical correlations, will be employed to develop and explore potential predictive relationships; results for series of molecular systems that demonstrate the viability of this approach are reported. A number of hydrazine salts have been synthesized for evaluation. Computational chemistry methods are being used to elucidate the mechanism of action of these salts. Toxicity endpoints such as viability (LDH) and changes in enzyme activity (glutahoione peroxidase and catalase) are being experimentally measured as indicators of cellular damage. Extrapolation from computational/in vitro studies to human toxicity, is the ultimate goal. The product of this program will be a predictive tool to assist in the development of new, less toxic propellants.

Frazier, John M.↗
NASA NTRS2001

Development and Lab-Scale Testing of a Gas Generator Hybrid Fuel in Support of the Hydrogen Peroxide Hybrid Upper Stage Program

As part of a NASA funded contract to develop and demonstrate a gas generator cycle hybrid rocket motor for upper stage space motor applications, the development and demonstration of a low sensitivity, high performance fuel composition was undertaken. The ultimate goal of the development program was to demonstrate successful hybrid operation (start, stop, throttling) of the fuel with high concentration (90+%) hydrogen peroxide. The formulation development and lab-scale testing of a simple DOT Class 1.4c gas generator propellant is described. Both forward injected center perforated and aft injected end burner hybrid combustion behavior were evaluated with gaseous oxygen and catalytically decomposed 90% hydrogen peroxide. Cross flow and static environments were found to yield profoundly different combustion behaviors, which were further governed by binder type, oxidizer level and, significantly, oxidizer particle size. Primary extinguishment was accomplished via manipulation of PDL behavior and oxidizer turndown, which is enhanced with the hydrogen peroxide system. Laboratory scale combustor results compared very well with 11-inch and 24-inch sub-scale test results with 90% hydrogen peroxide.

Lund, Gary K.↗
NASA NTRS2017

Numerical Modeling of an Integrated Vehicle Fluids System Loop for Pressurizing a Cryogenic Tank

This paper presents a numerical model of the pressurization loop of the Integrated Vehicle Fluids (IVF) system using the Generalized Fluid System Simulation Program (GFSSP). The IVF propulsion system, being developed by United Launch Alliance to reduce system weight and enhance reliability, uses boiloff propellants to drive thrusters for the reaction control system as well as to run internal combustion engines to develop power and drive compressors to pressurize propellant tanks. NASA Marshall Space Flight Center (MSFC) conducted tests to verify the functioning of the IVF system using a flight-like tank. GFSSP, a finite volume based flow network analysis software developed at MSFC, has been used to support the test program. This paper presents the simulation of three different test series, comparison of numerical prediction and test data and a novel method of presenting data in a dimensionless form. The paper also presents a methodology of implementing a compressor map in a system level code.

LeClair, A. C.↗
NASA NTRSDate not supplied

Hybrid Modeling for Complex Systems Health Management

The research work presents application of hybrid physics-informed machine learning to a representative electric powertrain for unmanned aerial vehicles. The model is composed of physics-derived and empirical equations, integrated with connected networks that are strategically placed within the model to substitute equations that are subject to large uncertainty. Polynomial fit driven by heuristics or empirical observations can be substituted by more flexible networks that can minimize the error between model predictions and observations without being restricted to a predefined functional form. This modeling strategy allows training of networks deep inside the model and unknown parameters in a single learning stage. The powertrain model consists of Li-ion batteries, electronic speed controller with pulse-width modulation, and brush-less DC motor with connected propeller. Results obtained from combination of laboratory and simulation tests are discussed in this work.

PINNS↗
NASA NTRS1975

Interior noise levels of two propeller-driven light aircraft

The relationships between aircraft operating conditions and interior noise and the degree to which ground testing can be used in lieu of flight testing for performing interior noise research were studied. The results show that the noise inside light aircraft is strongly influenced by the rotational speed of the engine and propeller. Both the overall noise and low frequency spectra levels were observed to decrease with increasing high speed rpm operations during flight. This phenomenon and its significance is not presently understood. Comparison of spectra obtained in flight with spectra obtained on the ground suggests that identification of frequency components and relative amplitude of propeller and engine noise sources may be evaluated on stationary aircraft.

Catherines, J. J.↗
DOE OSTI2020

Analysis of the Shattered Pellet Injection Fragment Plumes Generated by Machine Specific Shatter Tube Designs

Shattered pellet injector systems have been installed on DIII-D, JET, and KSTAR and used to experimentally determine the effectiveness of the shattered pellet injection (SPI) process in mitigating the deleterious effects of a tokamak plasma disruption. Pellets are fired, and before entering the plasma, strike a bent tube known as a shatter tube causing the pellet to shatter. The process of pellet fragmentation is a chaotic process that can be described in terms of fragment size distribution through a statistical model that incorporates the effects of the pellet material and impact characteristics. In addition to the fragment size distribution, the shatter plume has other characteristics of interest, such as a fragment velocity distribution and temporal mass evolution. The fragment velocity distribution is important because it is needed to accurately model the spread and location of the ablation and the deposition of impurities in the plasma over time. The temporal mass evolution is necessary to determine the time-resolved delivery of mass to the plasma.Due to installation constraints, the shatter tube currently installed on JET has a unique geometry with a modest S-bend followed by a 20-deg bend at the end of the tube. The DIII-D and KSTAR shatter tube design is a simple tube bent through an angle of 20 deg followed by a straight section. The resulting shatter sprays from the JET shatter tube and a 20-deg miter bend shatter tube were experimentally characterized for various pellet materials and speeds. Laboratory testing of these shatter tubes allows for the use of fast cameras to capture the fragment spray traveling through a large vacuum chamber. These high-speed videos of the shatter plumes allow the fragment size distribution, temporal mass evolution, and velocity distribution of the fragments within the plume to be determined. Overall, this paper presents a comparison of the unique geometry of the JET shatter tube to the miter bend geometries used for shattering and some insight into the variables that may be adjusted to produce the optimal shatter spray. The impact of entrained propellant gas on the resulting shatter spray was examined during testing.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1985

Ignition characterization of LOX/hydrocarbon propellants

The results of an evaluation of the ignition characteristics of the gaseous oxygen (Gox)/Ethanol propellant combination are summarized. Ignition characterization was accomplished through the analysis, fabrication and testing of a spark initiated torch igniter and prototype 620 1bF thruster/igniter assembly. The igniter was tested over a chamber pressure range of 3.3 to 262 psia and a mixture ratio range of 0.4 to 40. The prototype thruster/igniter assembly was tested over the chamber pressure range of 74 to 197 psia and mixture ratio range of 0.778 to 3.29. Cold (-92 to -165 F) and ambient (44 to 80 F) propellant temperatures were used. Spark igniter ignition limits and thruster steady state and pulse mode, performance, cooling and stability data are presented. Spark igniter ignition limits are presented in terms of cold flow pressure, ignition chamber diameter and mixture ratio. Thruster performance is presented in terms of vacuum specific impulse versus engine mixture ratio. Gox/Ethanol propellants were shown to be ignitable over a wide range of mixture ratios. Cold propellants were shown to have a minor effect on igniter ignition limits. Thruster pulse mode capability was demonstrated with multiple pulses of 0.08 sec duration and less.

Lawver, B. R.↗
NASA NTRS2010

Status of the NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test After 30,352 Hours of Operation

The NASA s Evolutionary Xenon Thruster (NEXT) program is tasked with significantly improving and extending the capabilities of current state-of-the-art NSTAR thruster. The service life capability of the NEXT ion thruster is being assessed by thruster wear test and life-modeling of critical thruster components, such as the ion optics and cathodes. The NEXT Long-Duration Test (LDT) was initiated to validate and qualify the NEXT thruster propellant throughput capability. The NEXT thruster completed the primary goal of the LDT; namely to demonstrate the project qualification throughput of 450 kg by the end of calendar year 2009. The NEXT LDT has demonstrated 30,352 hr of operation and processed 490 kg of xenon throughput--surpassing the NSTAR Extended Life Test hours demonstrated and more than double the throughput demonstrated by the NSTAR flight-spare. Thruster performance changes have been consistent with a priori predictions. Thruster erosion has been minimal and consistent with the thruster service life assessment, which predicts the first failure mode at greater than 750 kg throughput. The life-limiting failure mode for NEXT is predicted to be loss of structural integrity of the accelerator grid due to erosion by charge-exchange ions.

Herman, Daniel A.↗
NASA NTRSDate not supplied

Survivability of NASA’s 10k RDRE

Rotating detonation rocket engines (RDRE) have widespread global industry and academic interest. Lessons learned in survivability is critical to furthering this technology. To support RDRE development, NASA hot-fire tested several 10k lbf thrust class thrust chamber assemblies with various propellent combinations including liquid oxygen/ liquid methane, liquid oxygen/ liquid hydrogen, and liquid oxygen/RP-1. The thrust chamber assemblies include injectors, inner bodies, and outer bodies that were additively manufactured (AM) laser powder bed fusion (L-PBF) GRCop-42. This paper will cover hardware damage identified in the 2023 and 2024 RDRE testing campaigns and lessons learned to improve survivability for future RDREs. Particularly notable were the findings concerning the number of detonation waves formed and their effects on the bolted interface and its sealing.

Tessa M. Fedotowsky↗
NASA NTRS1992

Results of a 1462 hour ammonia arcjet endurance test

A total of 1462 hours of operaton were demonstrated in an endurance test of a 30 kWe-class ammonia arcjet operated at 10 kWe. The propellant flow rate was 0.170 g/s, and the measured performance increased from approximately 650 s specific impulse at 36 percent efficiency at the beginning of the test to 675 s at 39 percent near the end. The voltage increased and the current dropped slightly over the first 400 hours, and then remained approximately constant for the remainder of the test. The test, scheduled for 1500 hours, was terminated after an insulator in the rear of the engine failed. Post-test examination of the electrodes revealed only minimal damage. Although there was evidence of a number of mass transport processes occurring in the discharge chamber, the primary life-limiting wear mechanisms appear to be cathode tip erosion and constrictor melting.

Polk, J. E.↗
NASA NTRS2008

Igniters for Liquid Oxygen/Liquid Methane Technology Development

As part of NASA's technology development of liquid methane / liquid oxygen engines for future exploration missions, two different igniters were recently studied at NASA Marshall Space Flight Center. The first igniter tested was an impinging injection, spark-initiated torch igniter, and the second was a microwave-generated plasma igniter. The purpose of the ignition tests was to define the ignition limits under vacuum conditions and characterize the transient start-up performance as a function of propellant mixture ratio (MR), mass flow rates, inlet temperatures, and pre-ignition chamber pressure. In addition, for the impinging igniter two different spark plugs were tested, and for the microwave igniter the magnetron filament warm-up time and the magnetron input power were both varied. The results gathered from these tests indicated that the impinging igniter is capable of operating over an MR range of 2 - 27, with methane and oxygen inlet temperatures as low as -161 F and -233 F, respectively. The microwave igniter was tested over an MR range of 2 - 9, with methane and oxygen inlet temperatures as low as -90 F and -200 F, respectively. The microwave igniter achieved ignition over this range, although an upper ignition limit was determined for the oxidizer mass flow rate. In general, the torch exhaust temperatures for the microwave igniter were not as high as those attained with the impinging igniter. The microwave igniter, however, was hot-fired 17 times and was still operational, whereas the impinging igniter spark plugs experienced thermal shock and erosion over nine hot-fire tests. It was concluded that for the microwave igniter better mixing of the propellants might be required in order to both raise the torch exhaust temperature and decrease the required magnetron input power, and for the impinging igniter the spark plug position within the igniter chamber should be varied in future tests to identify a more optimal location. All of the igniter tests were supported by the Propulsion & Cryogenics Advanced Development project, which is part of NASA's Exploration Technology Development Program.

Osborne, Robin J.↗
NASA NTRS1992

Analysis of an advanced ducted propeller subsonic inlet

A time marching Navier-Stokes code called PARC (PARC2D for 2-D/axisymmetric and PARC3D for 3-D flow simulations) was validated for an advanced ducted propeller (ADP) subsonic inlet. The code validation for an advanced ducted propeller (ADP) subsonic inlet. The code validation was implemented for a non-separated flow condition associated with the inlet operating at angles-of-attack of 0 and 25 degrees. The inlet test data were obtained in the 9 x 15 ft Low Speed Wind Tunnel at NASA Lewis Research Center as part of a cooperative study with Pratt and Whitney. The experimental study focused on the ADP inlet performance for take-off and approach conditions. The inlet was tested at a free stream Mach number of 0.2, at angles-of-attack between O and 35 degrees, and at a maximum propeller speed of 12,000 RPM which induced a corrected air flow rate of about 46 lb/sec based on standard day conditions. The computational grid and flow boundary conditions (BC) were based on the actual inlet geometry and the funnel flow conditions. At the propeller face, two types of BC's were applied: a mass flow BC and a fixed flow properties BC. The fixed flow properties BC was based on a combination of data obtained from the experiment and calculations using a potential flow code. Comparison of the computational results with the test data indicates that the PARC code with the propeller face fixed flow properties BC provided a better prediction of the inlet surface static pressures than the predictions when the mass flow BC was used. For an angle-of-attack of 0 degrees, the PARC2D code with the propeller face mass flow BC provided a good prediction of inlet static pressures except in the region of high pressure gradient. With the propeller face fixed flow properties BC, the PARC2D code provided a good prediction of the inlet static pressures. For an angle-of-attack of 25 degrees with the mass flow BC, the PARC3D code predicted statis pressures which deviated significantly from the test data; however, with the fixed flow properties BC, a good comparison with the test data was obtained.

Iek, Chanthy↗