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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1976

Vibration characteristics of composite fan blades and comparison with measured data

The vibration characteristics of a composite fan blade for high-tip-speed applications were determined theoretically and the results compared with measured data. The theoretical results were obtained using a computerized capability consisting of NASTRAN coupled with composite mechanics by way of pre- and postprocessors. The predicted vibration frequencies and mode shapes were in reasonable agreement with the measured data. Theoretical results showed that different laminate configurations from the same composite system had only small effects on the blade frequency. However, the use of adhesively bonded titanium/beryllium laminar composites may improve considerably the blade vibration characteristics.

Chamis, C. C.↗
NASA NTRS1983

Oil-flow study of a Space Shuttle orbiter tip-fin controller

Possible use of tip-fin controllers instead of a vertical tail on advanced winged entry vehicles was examined. Elimination of the vertical tail and using tip-fins offers the advantages of positive yaw control at high angles of attack and a potential weight savings. Oil-flow technique was used to obtain surface flow patterns on a tip-fin installed on a 0.01-scale Space Shuttle orbiter model for the purpose of assessing the extent of flow interference effects on the wing and tip-fin which might lead to serious heating problems. Tests were conducted in air at Mach 10 for a free-stream Reynolds numbers of .000113 at 20, 30, and 40 degree angle of attack and sideslip angles of 0 and 2 degree. Elevon deflections of -10, 0, and 10 degree and tip-fin control-surface deflections of 0, 20, and 40 degree were employed. Test results were also used to aid in the interpretation of heating data obtained on a Shuttle orbiter tip-fin on another model in a different facility. A limited comparison of oil-flow patterns and heat-transfer data is included. It was determined that elevon deflection angles from -10 to 10 degree and sideslip angles up to 2 degree have very little effect on tip-fin surface flow patterns. Also, there is a minimum of interference between the tip-fin and the wing. The most significant flow interactions occur on the tip-fin onboard surface as a result of its control-surface deflections.

Helms, V. T., III↗
NASA NTRS1990

Helicopter far-field acoustic levels as a function of reduced rotor speeds

This paper will present far-field measured noise levels relative to tests conducted with a model S-76A helicopter. The project was designed to provide supplemental experimental flight data which may be used to further study reduced helicopter rotor speeds (and thus, advancing blade-tip Mach number) effects on far-field acoustic levels. The aircraft was flown in straight and level flight while operating with both the rotor speed and flight speed as test variables. The rotor speed was varied over the range of 107 percent of the main-rotor speed (NR) to 90 percent NR and with the forward flight speed varied over the range of 155 to 35 knots indicated air speed. These conditions produced a wide range of advancing blade-tip Mach numbers to which the noise data are related.

Mueller, Arnold W.↗
DOE OSTI2020

Measurement of crack length in width tapered beam experiments

The width tapered beam method for measuring fracture toughness has been contemplated for use in measuring the adhesion of photovoltaic materials, is promoted by it being a viable method for use directly on modules rather than to engineered test specimens. However, precise determination of crack length is often difficult as failure often involves tendril formation, cavitation voids, and other plastic deformation at the crack front in addition to obfuscation by opaque materials. Because fracture energy varies as the square of crack length, imprecise crack tip length results in large measurement uncertainties. In this work, we develop a method to estimate the crack tip length using the relationship between the applied force and the displacement of the load frame. The crack tip is determined for every data point and can be defined from a calibration curve related to the amount of energy put into the beam, or from a theoretically idealized estimation. Once a calibration curve is set up and the equations programmed, the fracture energy as a function of position on the newly formed crack interface can be accurately estimated. The test additionally provides much easier analysis and insight into the joint failure through more precise knowledge of fracture toughness at each de-bonded area.

42 ENGINEERING↗
NASA NTRS1972

Wind tunnel simulation of full scale vortices.

An experimental investigation has been conducted to determine the important scaling parameters for the flow in the core region of a vortex generated by a rectangular wing tip. The effect of an unconventional planform, the ogee tip, on the tip vortex is also determined. Data were measured for wing pitch angles of 6 and 9 deg at Mach numbers of 0.2, 0.5, and 0.6 for Reynolds numbers ranging from 440,000 to 7,000,000 using a triaxial hot wire probe. The probe was located at 2 and 5 chordlengths downstream from the trailing edge of the 4.25 inch and 26.0 inch chord wings. For rectangular planform wings, the measured vortex core diameter to chord ratios, peak tangential velocity ratios, and axial velocity ratios are shown to be functions only of wing lift coefficient and elapsed time from vortex formation, and appear to be independent of both Mach number and Reynolds number.

Rorke, J. B.↗
NASA NTRS1976

Wind tunnel investigation of a 14 foot vertical axis windmill

A full scale wind tunnel investigation was made to determine the performance characteristics of a 14 ft diameter vertical axis windmill. The parameters measured were wind velocity, shaft torque, shaft rotation rate, along with the drag and yawing moment. A velocity survey of the flow field downstream of the windmill was also made. The results of these tests along with some analytically predicted data are presented in the form of generalized data as a function of tip speed ratio.

Muraca, R. J.↗
NASA NTRS1979

Pressure and thermal distributions on wings and adjacent surfaces induced by elevon deflections at Mach 6

Surface pressure distributions and heat transfer distributions were obtained on wing half-models in regions where three dimensional separated flow effects are prominent. Unswept and 50 deg and 70 deg swept semispan wings were tested, for trailing-edge-elevon ramp angles of 0 deg, 10 deg, 20 deg, and 30 deg, with and without cylindrical and flat plate center bodies and with and without various wing-tip plates and fins. The data, obtained for a free stream Mach number of 6 and a wing-root-chord Reynolds number of 18.5 million, reveal considerably larger regions of increased pressure and thermal loads than would be anticipated using non-separated flow analyses.

Kaufman, L. G., II↗
NASA NTRS1986

Dual throat thruster test results

Test results are presented for a dual throat thruster burning gaseous oxygen and hydrogen at primary (inner) chamber pressures from 380 to 680 psia. Heat flux profiles were obtained from calorimetric cooling channels in the inner nozzle, outer or secondary chamber and the tip of the inner nozzle. Data were obtained for two nozzle spacings over a chamber pressure ratio (secondary/primary) range of 0.45 to 0.83 with both chambers firing (Mode I). Fluxes near the end of the inner nozzle were significantly higher than in Mode II when only the inner chamber was fired, due to flow separation and recirculation caused by the back pressure imposed by the secondary chamber. As the pressure ratio increased, these heat fluxes increased and the region of high heat flux relative to Mode II extended farther upstream. The use of a gaseous hydrogen bleed flow in the secondary chamber to control heat fluxes in the primary plume attachment region was investigated in Mode II testing. A bleed flow of less than two percent of the primary chamber flow was required.

Ewen, R. L.↗
NASA NTRS1986

Dual-throat thruster thermal model

The dual-throat engine is one of the dual nozzle engine concepts studied for advanced space transportation applications. It provides a thrust change and an in-flight area ratio change through the use of two concentric combustors with their throats arranged in series. Test results are presented for a dual throat thruster burning gaseous oxygen and hydrogen at primary (inner) chamber pressures from 380 to 680 psia. Heat flux profiles were obtained from calorimetric cooling channels in the inner nozzle, outer or secondary chamber and the tip of the inner nozzle. Data were obtained for two nozzle spacings over a chamber pressure ratio (secondary/primary) range of 0.45 to 0.83 with both chambers firing (Mode I). Fluxes near the end of the inner nozzle were significantly higher than in Mode II when only the inner chamber was fired, due to the flow separation and recirculation caused by the back pressure imposed by the secondary chamber. As the pressure ratio increased, these heat fluxes increased and the region of high heat flux relative to Mode II extended farther upstream. The use of the gaseous hydrogen bleed flow in the secondary chamber to control heat fluxes in the primary plume attachment region was investigated in Mode II testing. A thermal model of a dual throat thruster was developed and upgraded using the experimental data.

Ewen, R. L.↗
NASA NTRS1987

Geometrical acoustics and transonic helicopter sound

A new method is presented for predicting the impulsive noise generated by a transonic rotor blade. The method is a combined approach involving computational fluid dynamics and geometrical acoustics. A full-potential finite-difference method is used to obtain the pressure field close to the blade. A Kirchhoff integral formulation is then used to extend these finite-difference results into the far field. This Kirchhoff formula is based on geometrical acoustics approximations. It requires initial data across a plane at the sonic radius in a blade-fixed coordinate system. This data is provided by the finite-difference solution. Acoustic pressure predictions show good agreement with hover experimental data for cases with hover tip Mach numbers of 0.88 through 0.96. The cases above 0.92 tip Mach number are dominated by non-linear transonic effects seen as strong shocks on and off the blade tip. This paper gives the first successful predictions of far-field acoustic pressures for high-speed impulsive noise over a range of Mach numbers after delocalization.

Isom, Morris↗
NASA NTRS1988

Low-speed vortical flow over a 5-degree cone with tip geometry variations

Experimental results on the surface pressures and sectional side forces on a 5-deg cone were obtained for the cases of three different nose tips. The sectional side force data for both the sharp cone and the blunt nose cone configurations showed a dependence on roll orientation. The blunt nose configurations were found to be effective in reducing the sectional side force for angles of attack up to 25 deg.

Chu, J.↗
NASA NTRS1990

An analysis of the viscous flow through a compact radial turbine by the average passage approach

A steady, three-dimensional viscous average passage computer code is used to analyze the flow through a compact radial turbine rotor. The code models the flow as spatially periodic from blade passage to blade passage. Results from the code using varying computational models are compared with each other and with experimental data. These results include blade surface velocities and pressures, exit vorticity and entropy contour plots, shroud pressures, and spanwise exit total temperature, total pressure, and swirl distributions. The three computational models used are inviscid, viscous with no blade clearance, and viscous with blade clearance. It is found that modeling viscous effects improves correlation with experimental data, while modeling hub and tip clearances further improves some comparisons. Experimental results such as a local maximum of exit swirl, reduced exit total pressures at the walls, and exit total temperature magnitudes are explained by interpretation of the flow physics and computed secondary flows. Trends in the computed blade loading diagrams are similarly explained.

Heidmann, James D.↗
NASA NTRS1990

An analysis of the viscous flow through a compact radial turbine by the average passage approach

A steady, three-dimensional viscous average passage computer code is used to analyze the flow through a compact radial turbine rotor. The code models the flow as spatially periodic from blade passage to blade passage. Results from the code using varying computational models are compared with each other and with experimental data. These results include blade surface velocities and pressures, exit vorticity and entropy contour plots, shroud pressures, and spanwise exit total temperature, total pressure, and swirl distributions. The three computational models used are inviscid, viscous with no blade clearance, and viscous with blade clearance. It is found that modeling viscous effects improves correlation with experimental data, while modeling hub and tip clearances further improves some comparisons. Experimental results such as a local maximum of exit swirl, reduced exit total pressures at the walls, and exit total temperature magnitudes are explained by interpretation of the flow physics and computed secondary flows. Trends in the computed blade loading diagrams are similarly explained.

Heidmann, James D.↗
NASA NTRS2001

A Practical Tutorial on Modified Condition/Decision Coverage

This tutorial provides a practical approach to assessing modified condition/decision coverage (MC/DC) for aviation software products that must comply with regulatory guidance for DO-178B level A software. The tutorial's approach to MC/DC is a 5-step process that allows a certification authority or verification analyst to evaluate MC/DC claims without the aid of a coverage tool. In addition to the MC/DC approach, the tutorial addresses factors to consider in selecting and qualifying a structural coverage analysis tool, tips for reviewing life cycle data related to MC/DC, and pitfalls common to structural coverage analysis.

Hayhurst, Kelly J.↗
NASA NTRS1960

Landing-Impact Characteristics of Load-Alleviating Struts on a Model of a Winged Space Vehicle

A brief experimental investigation was made of the landing-impact characteristics of a 1/9-scale dynamic model of a winged space vehicle. The landing tests were made by catapulting a free model onto a hard; surface runway and onto water. The model had a conical fuselage and a flat - plate wing with a basic delta planform and 75 deg sweepback of the leading edge. The use of yielding-metal shock absorbers and various landing-gear arrangements was investigated during landing impact. The basic landing gear consisted of a dual rubber-tired nose wheel and twin main skids aft of the center of gravity near the wing tips. landing motion and acceleration data were obtained over a range of landing attitudes, gross weights, and initial sinking speeds. Brief tests were made with an alternate nose-wheel location. An all-skid configuration also was briefly evaluated for hard-surface and water landings. The landing gear employing yielding struts for impact-energy absorption during hard-surface landings resulted in accelerations of approximately 5 1/2 g near the nose gear over a range of landing parameters. Replacing the nose wheel and tire with a skid did not significantly change the accelerations. Landings in smooth water with rigid struts and adequate planing area at the nose skid resulted in a maximum landing acceleration of approximately 4g.

Blanchard, Ulysse J.↗
NASA NTRS2010

FSSC Science Tools: Pulsar Analysis

This slide presentation reviews the typical pulsar analysis, giving tips for screening of the data, the use of time series analysis, and utility tools. Specific information about analyzing Vela data is reviewed.

Thompson, Dave↗
NASA NTRS1986

Development of a ROT22 - DATAMAP interface

This report (Contract NAS2-10331- Mod 10), outlines the development and validation of an interface between the three-dimensional transonic analysis program ROT22 and the Data from Aeromechanics Test and Analytics-Management and Analysis Package (DATAMAP). After development of the interface, the validation is carried out as follows. First, the DATAMAP program is used to analyze a portion of the Tip Aerodynamics and Acoustics Test (TAAT) data. Specifically, records 2872 and 2873 are analyzed at an azimuth of 90 deg, and record 2806 is analyzed at 60 deg. Trim conditions for these flight conditions are then calculated using the Bell performance prediction program ARAM45. Equivalent shaft, pitch, and twist angles are calculated from ARAM45 results and used as input to the ROT22 program. The interface uses the ROT22 results and creates DATAMAP information files from which the surface pressure contours and sectional pressure coefficients are plotted. Twist angles input to ROT22 program are then iteratively modified in the tip region until the computed pressure coefficients closely match the measurements. In all cases studied, the location of the shock is well predicted. However, the negative pressure coefficients were underpredicted. This could be accounted for by blade vortex interaction effects.

Shenoy, K. R.↗
NASA NTRS1985

The geometric properties of Landsat-5 Thematic Mapper data and their conformity to Landsat-4 data and to earth's surface geometry

The geometry of two TIPS processed Landsat-5 Thematic Mapper scenes was analyzed and compared with that of SCROUNGE processed Landsat-4 data. Swath-to-swath and band-to-band registration of Washington, DC, and northeastern Iowa scenes was found to be similar to or better than that of Landsat-4 data. Results indicate a high degree of geometric conformity between the images produced by the different systems. The geometric conformity of the TIPS processed images to the Space Oblique Mercator projection, however, proved to be less accurate than the targeted processing error of 15 meters.

Walker, R. E.↗