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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 541 records · Page 30
All sourcesNASA NTRSDOE OSTI
DOE OSTI2014

Performance of Cladding on MOX Fuel with Low 240Pu/239Pu Ratio

The U.S. Department of Energy has decided to dispose of a portion of its surplus plutonium by reconstituting it into mixed oxide (MOX) fuel and irradiating it in commercial power reactors. As part of fuel qualification, four lead assemblies were manufactured and irradiated to a maximum fuel rod average burnup of 47.3 MWd/kg heavy metal. This was the world s first commercial irradiation of MOX fuel with a 240Pu/239Pu ratio less than 0.10. Five fuel rods with varying burnups and plutonium contents were selected from one of the assemblies and shipped to Oak Ridge National Laboratory for hot cell examination. This paper discusses the results of those examinations with emphasis on cladding performance. Exams relevant to the cladding included visual and eddy current exams, profilometry, microscopy, hydrogen analysis, gallium analysis, and mechanical testing. There was no discernible effect of the type of MOX fuel on the performance of the cladding.

FMDP↗
NASA NTRS2004

Experimental Investigation of a High Pressure Ratio Aspirated Fan Stage

The experimental investigation of an aspirated fan stage designed to achieve a pressure ratio of 3.4:1 at 1500 ft/sec is presented in this paper. The low-energy viscous flow is aspirated from diffusion-limiting locations on the blades and flowpath surfaces of the stage, enabling a very high pressure ratio to be achieved in a single stage. The fan stage performance was mapped at various operating speeds from choke to stall in a compressor facility at fully simulated engine conditions. The experimentally determined stage performance, in terms of pressure ratio and corresponding inlet mass flow rate, was found to be in good agreement with the three-dimensional viscous computational prediction, and in turn close to the design intent. Stage pressure ratios exceeding 3:1 were achieved at design speed, with an aspiration flow fraction of 3.5 percent of the stage inlet mass flow. The experimental performance of the stage at various operating conditions, including detailed flowfield measurements, are presented and discussed in the context of the computational analyses. The sensitivity of the stage performance and operability to reduced aspiration flow rates at design and off design conditions are also discussed.

Merchant, Ali↗
DOE OSTI2018

Fatigue crack growth rates of X100 steel welds in high pressure hydrogen gas considering residual stress effects

Fatigue crack growth rate (FCGR) data were measured in high pressure hydrogen gas versus stress intensity factor range (ΔK) in specimens removed from a X100 welded steel pipe. Three distinct regions of the pipe weld were examined: base metal, weld fusion zone, and heat affected zone. Tests were performed at a load ratio (R) of 0.5, frequency of 1 Hz, and at a hydrogen gas pressure of 21 MPa. Tests were also performed in air at 10 Hz as a reference. Fatigue crack growth rates were observed to be over an order of magnitude higher for tests performed in hydrogen compared to the rates from tests in air. Residual stress measurements were collected on identical specimens cut from the base metal, weld, and heat affected zone to account for their influence on measured FCGR data. The slitting method provided residual stress and residual stress intensity factor (K res ), the effect of which was removed from the FCGR data using K norm in order to provide a more direct comparison of crack growth resistance of the base metal, weld and heat affected zone. Prior to accounting for residual stress, FCGR in hydrogen gas appeared to be highest in the weld fusion zone. Furthermore, after accounting for residual stress effects, the weld fusion zone FCGR data converged to the base metal FCGR data, which underscores the importance of accounting for residual stress effects when assessing fatigue performance.

42 ENGINEERING↗
NASA NTRS1993

Design and test of a small two stage counter-rotating turbine for rocket engine application

The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The technology represented by this turbine is being developed for application in an advanced upper stage rocket engine turbopump. This engine will employ an oxygen/hydrogen expander cycle and achieve high performance through efficient combustion, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low flow rates result in very small airfoil diameter, height and chord. The high efficiency and small size requirements present a challenging turbine design problem. The unconventional approach employed to meet this challenge is described, along with the detailed design process and resulting airfoil configurations. The method and results of full scale aerodynamic performance evaluation testing of both one and two stage configurations, as well as operation without the secondary stage stator are presented. The overall results of this effort illustrate that advanced aerodynamic design tools and hardware fabrication techniques have provided improved capability to produce small high performance turbines for advanced rocket engines.

Huber, F. W.↗
DOE OSTI2022

Effects of Port Angle on Scavenging of an Opposed Piston Two-Stroke Engine

Opposed-piston 2-stroke (OP-2S) engines have the potential to achieve higher thermal efficiency than a typical diesel engine. However, the uniflow scavenging process is difficult to control over a wide range of speeds and loads. Scavenging performance is highly sensitive to pressure dynamics, port timings, and port design. This study proposes an analysis of the effects of port vane angle on the scavenging performance of an opposed-piston 2-stroke engine via simulation. A CFD model of a three-cylinder opposed-piston 2-stroke was developed and validated against experimental data collected by Achates Power Inc. One of the three cylinders was then isolated in a new model and simulated using cycle-averaged and cylinder-averaged initial/ boundary conditions. This isolated cylinder model was used to efficiently sweep port angles from 12 degrees to 29 degrees at different pressure ratios. Results indicate that scavenging performance is correlated with the bulk swirl ratio generated by these port angles. Scavenging performance is also sensitive to the pressure ratio across the engine. It was concluded that, for a given pressure ratio, the smallest port angle produces the best scavenging; however, previous work indicates lower bulk swirl ratio during combustion can have an impact on emissions formation.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1977

Acoustic and aerodynamic performance of a variable-pitch 1.83-meter-(6-ft) diameter 1.20-pressure-ratio fan stage (QF-9)

Far field noise data and related aerodynamic performance are presented for a variable pitch fan stage having characteristics suitable for low noise, STOL engine application. However, no acoustic suppression material was used in the flow passages. The fan was externally driven by an electric motor. Tests were made at several forward thrust rotor blade pitch angles and one for reverse thrust. Fan speed was varied from 60 to 120 percent of takeoff (design) speed, and exhaust nozzles having areas 92 to 105 percent of design were tested. The fan noise level was at a minimum at the design rotor blade pitch angles of 64 deg for takeoff thrust and at 57 deg for approach (50 percent takeoff thrust). Perceived noise along a 152.4-m sideline reached 100.1 PNdb for the takeoff (design) configuration for a stage pressure ratio of 1.17 and thrust of 57,600 N. For reverse thrust the PNL values were 4 to 5 PNdb above the takeoff values at comparable fan speeds.

Glaser, F. W.↗
NASA NTRS1987

A detailed description of the uncertainty analysis for high area ratio rocket nozzle tests at the NASA Lewis Research Center

A preliminary uncertainty analysis was performed for the High Area Ratio Rocket Nozzle test program which took place at the altitude test capsule of the Rocket Engine Test Facility at the NASA Lewis Research Center. Results from the study establish the uncertainty of measured and calculated parameters required for the calculation of rocket engine specific impulse. A generalized description of the uncertainty methodology used is provided. Specific equations and a detailed description of the analysis is presented. Verification of the uncertainty analysis model was performed by comparison with results from the experimental program's data reduction code. Final results include an uncertainty for specific impulse of 1.30 percent. The largest contributors to this uncertainty were calibration errors from the test capsule pressure and thrust measurement devices.

Davidian, Kenneth J.↗
NASA NTRS1977

Quiet Clean Short-haul Experimental Engine (QCSEE). Aerodynamic and aeromechanical performance of a 50.8 cm (20 inch) diameter 1.34 PR variable pitch fan with core flow

The fan aerodynamic and aeromechanical performance tests of the quiet clean short haul experimental engine under the wing fan and inlet with a simulated core flow are described. Overall forward mode fan performance is presented at each rotor pitch angle setting with conventional flow pressure ratio efficiency fan maps, distinguishing the performance characteristics of the fan bypass and fan core regions. Effects of off design bypass ratio, hybrid inlet geometry, and tip radial inlet distortion on fan performance are determined. The nonaxisymmetric bypass OGV and pylon configuration is assessed relative to both total pressure loss and induced circumferential flow distortion. Reverse mode performance, obtained by resetting the rotor blades through both the stall pitch and flat pitch directions, is discussed in terms of the conventional flow pressure ratio relationship and its implications upon achievable reverse thrust. Core performance in reverse mode operation is presented in terms of overall recovery levels and radial profiles existing at the simulated core inlet plane. Observations of the starting phenomena associated with the initiation of stable rotor flow during acceleration in the reverse mode are briefly discussed. Aeromechanical response characteristics of the fan blades are presented as a separate appendix, along with a description of the vehicle instrumentation and method of data reduction.

Giffin, R. G.↗
NASA NTRS1972

Performance and control study of a low-pressure-ratio turbojet engine for a drone aircraft

The results of analog and digital computer studies of a low-pressure-ratio turbojet engine system for use in a drone vehicle are presented. The turbojet engine consists of a four-stage axial compressor, single-stage turbine, and a fixed area exhaust nozzle. Three simplified fuel schedules and a generalized parameter fuel control for the engine system are presented and evaluated. The evaluation is based on the performance of each schedule or control during engine acceleration from a windmill start at Mach 0.8 and 6100 meters to 100 percent corrected speed. It was found that, because of the higher acceleration margin permitted by the control, the generalized parameter control exhibited the best dynamic performance.

Seldner, K.↗
NASA NTRS1987

Ion-molecule calculation of the abundance ratio of CCD to CCH in dense interstellar clouds

Laboratory measurements and calculations have been performed to determine the abundance ratio of the deuterated ethynyl radical (CCD) to the normal radical (CCH) which can be achieved in dense interstellar clouds via isotopic fractionation in the C2H2(+) (HD)=C2HD(+)(H2) system of reactions. According to this limited treatment, the CCD/CCH abundance ratio which can be attained is in the range 0.02-0.03 for the Orion molecular cloud and 0.0l-0.02 for TMC-1. These ranges of numbers are in reasonable agreement with the observed values in Orion and TMC-1. However, the analysis of the CCD/CCH abundance ratio is complicated via the presence of competing fractionation mechanisms, especially in the low-temperature source TMC-1.

Herbst, Eric↗
DOE OSTI2017

Polypeptide Composite Particle-Assisted Organization of π-Conjugated Polymers into Highly Crystalline “Coffee Stains”

We demonstrate that homopolypeptides covalently tethered to anisotropically shaped silica particles induce crystalline ordering of representative semiconducting polymers. Films drop-cast from chloroform dispersions of poly(γ-stearyl-l-glutamate) (PSLG) composite particles and poly(3-hexythiophene) (P3HT) led to highly ordered crystalline structures of P3HT. Hydrophobic–hydrophobic interactions between the alkyl side chains of P3HT and PSLG were the main driving force for P3HT chain ordering into the crystalline assemblies. It was found that the orientation of rigid P3HT fibrils on the substrate adopted the directionality of the evaporating front. Regardless of the PSLG-coated particle dimensions used, the drop-cast films displayed patterns that were shaped by the coffee ring and Marangoni effects. PSLG-coated particles of high axial ratio (4.2) were more efficient in enhancing the electronic performance of P3HT than low axial ratio (2.6) homologues. Devices fabricated from the ordered assemblies displayed improved charge-carrier transport performance when compared to devices fabricated from P3HT alone. Furthermore, these results suggest that PSLG can favorably mediate the organization of semiconducting polymers.

36 MATERIALS SCIENCE↗
DOE OSTI2014

A Comparative Study of the Adsorption of Water and Methanol in Zeolite BEA: A Molecular Simulation Study

Grand Canonical Monte Carlo (GCMC) simulations were carried out to study the equilibrium adsorption concentration of methanol and water in all-silica zeolite BEA over the wide temperature and pressure ranges. For both water and methanol, their adsorptive capacity increases with increasing pressure and decreasing temperature. The onset of methanol adsorption occurs at much lower pressures than water adsorption at all temperatures. Our GCMC simulation results also indicate that the adsorption isotherms of methanol exhibit a gradual change with pressure while water adsorption shows a sharp first-order phase transition at low temperatures. To explore the effects of Si/Al ratio on adsorption, a series of GCMC simulations of water and methanol adsorption in zeolites HBEA with Si/Al=7, 15, 31, 63 were performed. As the Si/Al ratio decreases, the onsets of both water and methanol adsorption dramatically shift to lower pressures. The type V isotherm obtained for water adsorption in hydrophobic BEA progressively changes to type I isotherm with decreasing Si/Al ratio in hydrophilic HBEA. This work was supported by the US Department of Energy, Office of Basic Energy Sciences, Division of Chemical Sciences, Geosciences & Biosciences. Pacific Northwest National Laboratory (PNNL) is a multiprogram national laboratory operated for DOE by Battelle.

Environmental Molecular Sciences Laboratory↗
NASA NTRS1975

Hypersonic research engine project. Phase 2: Preliminary report on the performance of the HRE/AIM at Mach 6

Test results of the Aerothermodynamic Integration Model are presented. A program was initiated to develop a hydrogen-fueled research-oriented scramjet for operation between Mach 3 and 8. The primary objectives were to investigate the internal aerothermodynamic characteristics of the engine, to provide realistic design parameters for future hypersonic engine development as well as to evaluate the ground test facility and testing techniques. The engine was tested at the NASA hypersonic tunnel facility with synthetic air at Mach 5, 6, and 7. The hydrogen fuel was heated up to 1500 R prior to injection to simulate a regeneratively cooled system. The engine and component performance at Mach 6 is reported. Inlet performance compared very well both with theory and with subscale model tests. Combustor efficiencies up to 95 percent were attained at an equivalence ratio of unity. Nozzle performance was lower than expected. The overall engine performance was computed using two different methods. The performance was also compared with test data from other sources.

Sun, Y. H.↗
NASA NTRS1991

Performance of a high-work, low-aspect-ratio turbine stator tested with a realistic inlet radial temperature gradient

A 0.767-scale model of a turbine stator designed for the core of a high-bypass-ratio aircraft engine was tested with uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The principal measurements were radial and circumferential surveys of stator-exit total temperature, total pressure, and flow angle. The stator-exit flow field was also computed by using a three-dimensional Navier-Stokes solver. Other than temperature, there were no apparent differences in performance due to the inlet conditions. The computed results compared quite well with the experimental results.

Stabe, Roy G.↗
NASA NTRS1987

A detailed description of the uncertainty analysis for High Area Ratio Rocket Nozzle tests at the NASA Lewis Research Center

A preliminary uncertainty analysis has been performed for the High Area Ratio Rocket Nozzle test program which took place at the altitude test capsule of the Rocket Engine Test Facility at the NASA Lewis Research Center. Results from the study establish the uncertainty of measured and calculated parameters required for the calculation of rocket engine specific impulse. A generalized description of the uncertainty methodology used is provided. Specific equations and a detailed description of the analysis are presented. Verification of the uncertainty analysis model was performed by comparison with results from the experimental program's data reduction code. Final results include an uncertainty for specific impulse of 1.30 percent. The largest contributors to this uncertainty were calibration errors from the test capsule pressure and thrust measurement devices.

Kenneth J. Davidian↗
NASA NTRS1974

Acoustic and aerodynamic performance of a 1.83 meter (6 foot) diameter 1.2 pressure ratio fan (QF-6)

A 1.2-pressure-ratio, 1.83-meter-(6-ft-) diameter experimental fan stage with characteristics suitable for use in STOL aircraft engines was tested for acoustic and aerodynamic performance. The design incorporated features for low noise, including absence of inlet guide vanes, low rotor-blade-tip speed, low aerodynamic blade loading, and long axial spacing between the rotor and stator rows. The stage was run with four nozzles of different area. The perceived noise along a 152.4 meter (500-ft) sideline was rear-quadrant dominated with a maximum design-point level of 103.9 PNdb. The acoustic 1/3-octave results were analytically separated into broadband and pure-tone components. It was found that the stage noise levels generally increase with a decrease in nozzle area, with this increase observed primarily in the broadband noise component. A stall condition was documented acoustically with a 90-percent-of-design-area nozzle.

Woodward, R. P.↗