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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Thermal-mechanical fatigue test apparatus for metal matrix composites and joint attachments

Two thermal-mechanical fatigue (TMF) test facilities were designed and developed, one to test tungsten fiber reinforced metal matrix composite specimens at temperature up to 1430C (2600F) and another to test composite/metal attachment bond joints at temperatures up to 760C (1400 F). The TMF facility designed for testing tungsten fiber reinforced metal matrix composites permits test specimen temperature excursions from room temperature to 1430C (2600F) with controlled heating and loading rates. A strain-measuring device measures the strain in the test section of the specimen during each heating and cooling cycle with superimposed loads. Data is collected and recorded by a computer. The second facility is designed to test composite/metal attachment bond joints and to permit heating to a maximum temperature of 760C (1400F) within 10 min and cooling to 150C (300F) within 3 min. A computer controls specimen temperature and load cycling.

Westfall, L. J.↗
NASA NTRS1988

Thermal-mechanical fatigue test apparatus for metal matrix composites and joint attachments

Two thermal-mechanical fatigue (TMF) test facilities were designed and developed, one to test tungsten fiber reinforced metal matrix composite specimens at temperature up to 1430C (2600F) and another to test composite/metal attachment bond joints at temperatures up to 760F (1400F). The TMF facility designed for testing tungsten fiber reinforced metal matrix composites permits test specimen temperature excursions from room temperature to 1430C (2600F) with controlled heating and loading rates. A strain-measuring device measures the strain in the test section of the specimen during each heating and cooling cycle with superimposed loads. Data is collected and recorded by a computer. The second facility is designed to test composite/metal attachment bond joints and to permit heating to a maximum temperature of 760C (1400F) within 10 min and cooling to 150C (300F) within 3 min. A computer controls specimen temperature and load cycling.

Westfall, Leonard J.↗
NASA NTRS2015

Aerodynamic Load Analysis of a Variable Camber Continuous Trailing Edge Flap System on a Flexible Wing Aircraft

This paper presents an investigative aerodynamic analysis conducted on the novel control surface known as a Variable Camber Continuous Trailing Edge Flap (VCCTEF). The VCCTEF is modeled as a control effector on the NASA Generic Transport Model (GTM) where wing flexibility is considered. Aerodynamic modeling of the aircraft is conducted using vortex-lattice method (VLM), and an aeroelastic model of the aircraft that utilizes a coupled finite-element analysis (FEA) vortex-lattice solution is employed. VLM solutions are used to determine quasi-steady aerodynamic loading over the aeroelastic wing structures with VCCTEF. The load data is used to calculate aerodynamic sensitivities to control surface deployment and is also integrated to determine overall hinge moments. This analysis is conducted for different flight conditions, where control sensitivities compare the VCCTEF effectiveness against conventional control surfaces. Hinge moment results provide insight into aeroelastic wing loads, and worst case hinge moments for the VCCTEF can be estimated. Results show that the VCCTEF offers greater control authority than conventional ailerons at cruise, and nominal flap settings are determined for a low-speed take-off condition where the VCCTEF maintains comparable control effectiveness. Worst case hinge moment values are presented, where for a possible VCCTEF configuration, the stiff wing model demonstrates up to a 5.90 percent increase in flap hinge moment relative to a rigid model, and a reduced stiffness model demonstrates up to an 11.42 percent increase.

Aeroelasticity↗
NASA NTRS1981

Systems design and analysis of the microwave radiometer spacecraft

Systems design and analysis data were generated for microwave radiometer spacecraft concept using the Large Advanced Space Systems (LASS) computer aided design and analysis program. Parametric analyses were conducted for perturbations off the nominal-orbital-altitude/antenna-reflector-size and for control/propulsion system options. Optimized spacecraft mass, structural element design, and on-orbit loading data are presented. Propulsion and rigid-body control systems sensitivities to current and advanced technology are established. Spacecraft-induced and environmental effects on antenna performance (surface accuracy, defocus, and boresight off-set) are quantified and structured material frequencies and modal shapes are defined.

Garrett, L. B.↗
NASA NTRS2015

Microstructure-Sensitive Investigation of Fracture Using Acoustic Emission Coupled With Electron Microscopy

A novel technique using Scanning Electron Microscopy (SEM) in conjunction with Acoustic Emission (AE) monitoring is proposed to investigate microstructure-sensitive fatigue and fracture of metals. The coupling between quasi in situ microscopy with actual in situ nondestructive evaluation falls into the ICME framework and the idea of quantitative data-driven characterization of material behavior. To validate the use of AE monitoring inside the SEM chamber, Aluminum 2024-B sharp notch specimen were tested both inside and outside the microscope using a small scale mechanical testing device. Subsequently, the same type of specimen was tested inside the SEM chamber. Load data were correlated with both AE information and observations of microcracks around grain boundaries as well as secondary cracks, voids, and slip bands. The preliminary results are in excellent agreement with similar findings at the mesoscale. Extensions of the application of this novel technique are discussed.

Wisner, Brian↗
NASA NTRS1995

Comparison of measured and calculated dynamic loads for the Mod-2 2.5 mW wind turbine system

The Boeing Company, under contract to the Electric Power Research Institute (EPRI), has completed a test program on the Mod-2 wind turbines at Goodnoe Hills, Washington. The objectives were to update fatigue load spectra, discern site and machine differences, measure vortex generator effects, and to evaluate rotational sampling techniques. This paper shows the test setup and loads instrumentation, loads data comparisons and test/analysis correlations. Test data are correlated with DYLOSAT predictions using both the NASA interim turbulence model and rotationally sampled winds as inputs. The latter is demonstrated to have the potential to improve the test/analysis correlations. The paper concludes with an assessment of the importance of vortex generators, site dependence, and machine differences on fatigue loads. The adequacy of prediction techniques used are evaluated and recommendations are made for improvements to the methodology.

Zimmerman, D. K.↗
NASA NTRS2017

Microscopic Evaluation of Friction Plug Welds- Correlation to a Processing Analysis

Recently an analysis of dynamic forge load data from the friction plug weld (FPW) process and the corresponding tensile test results showed that good plug welds fit well within an analytically determined processing parameter box. There were, however, some outliers that compromised the predictions. Here the microstructure of the plug weld material is presented in view of the load analysis with the intent of further understanding the FPW process and how it is affected by the grain structure and subsequent mechanical properties.

Rabenberg, Ellen M.↗
NASA NTRS1985

Free-Tip Rotor Wind Tunnel Test Results

A new rotor blade tip design called the free-tip has been proposed as a means to improve forward flight performance characteristics and reduce oscillatory loads. The free-tip design incorporates a tip that is free to pitch independently of the rest of the blade. Pitching about an axis forward of the quarter-chord, the tip weathervanes into its local wind, thus reducing angle of attack perturbations and the resulting oscillatory lift loadings. A nearly constant nose-up pitching moment is applied mechanically to the tip so that the tip, to maintain pitching equilibrium, produces nearly steady positive lift around the azimuth. A wind-tunnel test of a small-scale, 5.1 m diameter model rotor was conducted to obtain comparative forward flight performance and oscillatory loads data with the tips free and fixed. The free-up was shown to reduce power in trimmed flight over a wide range of advance ratio and thrust; at an advance ratio of 0.3 and C(sub L)/sigma of 0.08 the reduction is 12%. Oscillatory flapwise bending-moments and oscillatory pitch link loads are also reduced, but the oscillatory in-plane bending moments increase.

Stroub, Robert H.↗
NASA NTRS2019

Aeroelastic Stability Analysis of a Full-Scale Isolated Proprotor on the Tiltrotor Test Rig

A full-scale isolated proprotor test was recently conducted in the USAF National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel at NASA Ames. The test article was a 3-bladed research rotor derived from the right-hand rotor of the AW609. For this test, the NASA Tiltrotor Test Rig (TTR) and rotor were installed in the 40- by 80-Foot Wind Tunnel. This paper covers the analyses and testing done to prepare for a safe entry. Included are brief descriptions of the following: NASTRAN models of the TTR, ground vibration tests of the TTR (and resulting modal data), loads analyses, and stability predictions using the comprehensive analysis CAMRAD II. The evolution of these analyses from early in the TTR program until the initiation of actual testing is also discussed. The intent is to show how all of these efforts were integrated to ensure a successful test. This paper includes stability predictions based on NASTRAN modal data and worst-case damping test data. The stability predictions covered all test conditions: hover, cruise (airplane mode), conversion, and helicopter mode. The predictions showed that the TTR and rotor are stable within the test envelope.

Full-Scale Isolated Proprotor↗
NASA NTRS1974

Results of tests 0A12 and IA9 in the Ames Research Center unitary plan wind tunnels on an 0.030 scale model of the space shuttle vehicle 2A to determine aerodynamic loads, volume 16

Tests were conducted in the NASA/ARC Unitary Plan Wind Tunnels on an 0.030 scale replica of the space shuttle vehicle Configuration 2A. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 3.5. The integrated vehicle was tested at angles of attack and sideslip from -8 degrees to +8 degrees. The isolated orbiter was tested at angles of attack from -15 degrees to +40 degrees and angles of sideslip from -10 degrees to +10 degrees as dictated by trajectory considerations. The effects of orbiter/external tank incidence angle and deflected control surfaces on aerodynamic loads were also investigated.

Spangler, R. H.↗
NASA NTRS1973

Results of tests OA12 and IA9 in the Ames Research Center unitary plan wind tunnels on an 0.030 scale model of the space shuttle vehicle 2A to determine aerodynamic loads, volume 7

Tests were conducted in wind tunnels during April and May 1973, on an 0.030-scale replica of the Space Shuttle Vehicle Configuration 2A. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 3.5. The investigation included tests on the integrated (launch) configuration and on the isolated orbiter (entry configuration). The integrated vehicle was tested at angles of attack and sideslip from -8 deg. The isolated orbiter was tested at angles of attack from -15 deg to +40 deg and angles of sideslip from -10 deg to +10 deg as dictated by trajectory considerations. The effects of orbiter/external tank incidence angle and deflected control surfaces on aerodynamic loads were also investigated.

Spangler, R. H.↗
NASA NTRS1992

National Launch System cycle 1 loads and models data book

This document contains preliminary cycle 1 loads for the National Launch System (NLS) 1 and 2 vehicles. The loads provided and recommended as design loads represent the maximum load expected during prelaunch and flight regimes, i.e., limit loads, except that propellant tank ullage pressure has not been included. Ullage pressure should be added to the loads book values for cases where the addition results in higher loads. The loads must be multiplied by the appropriate factors of safety to determine the ultimate loads for which the structure must be capable.

Bugg, F.↗
NASA NTRS1974

Results of tests OA12 and IA9 in the Ames Research Center unitary plan wind tunnels on an 0.030-scale model of the space shuttle vehicle 2A to determine aerodynamic loads, volume 18

Tests were conducted, during April and May 1973, on an 0.030-scale replica of the Space Shuttle Vehicle Configuration 2A. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 3.5. The investigation included tests on the integrated (launch) configuration, and on the isolated orbiter (entry configuration). The integrated vehicle was tested at angles of attack and sideslip from -8 degrees to +8 degrees. The isolated orbiter was tested at angles of attack from -15 degrees to +40 degrees and angles of sideslip from -10 degrees to +10 degrees as dictated by trajectory considerations. The effect of orbiter/external tank incidence angle and deflected control surfaces on aerodynamic loads were also investigated.

Spangler, R. H.↗
NASA NTRS1947

Effect of Slot-Entry Skirt Extensions on Aerodynamic Characteristics of a Wing Section of the XB-36 Airplane Equipped with a Double Slotted Flap

An investigation was made in the Langley two-dimensional low-turbulence tunnel on a wing section for the XB-36 airplane equipped with a double slotted flap to determine the effect on lift and drag of various slot-entry skirt extension. A skirt extension of 0.787 deg. was found to provide the best combination of high maximum lift with flap deflected and law drag with flap retracted. The data showed that the maximum lift at intermediate (20 deg. to 45 deg.) flap deflections was lowered considerably by the slot-entry extension; but at high flap deflections the effect was small. An increase in Reynolds number from 2.4 million to 6.0 million increased the maximum.lift coefficient at a flap deflection of 55 deg. from 3.12 to 3.30 and from 1.18 to 1.40 for the flap retracted condition, but did not greatly affect the maximum lift coefficient for intermediate flap deflections. The flap and fore flap load data indicated that the maximum lift coefficients at high flap deflections are limited by a breakdown in the flow over the .flaps.

Cahill, Jones F.↗
NASA NTRS1973

Results of Tests OA12 and IA9 in the Ames Research Center unitary plan wind tunnels on an 0.030-scale model of the Space Shuttle Vehicle 2A to determine aerodynamic loads, volume 4

Tests were conducted in wind tunnels during April and May 1973, on an 0.030-scale replica of the Space Shuttle Vehicle Configuration 2A. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 3.5. The investigation included tests on the integrated (launch) configuration and on the isolated orbiter (entry configuration). The integrated vehicle was tested at angles of attack and sideslip from minus 8 degrees to +8 degrees. The isolated orbiter was tested at angles of attack from minus 15 degrees to +40 degrees and angles of sideslip from minus 10 degrees to +10 degrees as dictated by trajectory considerations. The effects of orbiter/external tank incidence angle and deflected control surfaces on aerodynamic loads were also investigated.

Spangler, R. H.↗
NASA NTRS1973

Results of tests OA12 and IA9 in the Ames Research Center unitary plan wind tunnels on an 0.030-scale model of the space shuttle vehicle 2A to determine aerodynamic loads, volume 1

Tests were conducted in unitary plan wind tunnels on an 0.030-scale replica of the space shuttle vehicle configuration 2A. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 3.5. The investigation included tests on the integrated (launch) configuration and tests on the isolated orbiter (entry configuration). The integrated vehicle was tested at angles of attack and sideslip from minus 8 deg to plus 8 deg. The isolated orbiter was tested at angles of attack from minus 15 deg to plus 40 deg and angles of sideslip from minus 10 deg to plus 10 deg are dictated by trajectory considerations. The effects of orbiter/external tank incidence and deflected control surfaces on aerodynamic loads were also investigated.

Spangler, R. H.↗
NASA NTRS1940

Pressure-distribution investigation of an NACA 0009 airfoil with an 80-percent-chord plain flap and three tabs

Pressure-distribution tests of an NACA 0009 airfoil with an 80-percent-chord plain flap and three plain tabs, having chord of 10, 20, and 30 percent of the flap chord, were made. Section data suitable for application to the design of horizontal and vertical tail surfaces were obtained. Resultant-pressure diagrams for the airfoil with the flap and the 20-percent-chord tab are presented. Plots are also given of increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs. A comparison of some characteristic slopes for the 30-, the 50-, and the 80-percent-chord flaps, tested in the general investigation of plain flaps for control surfaces, is included. Section aerodynamic and load data have been made available for a wide range of flap and a tab chords to be used on an NACA 0009 airfoil or on other conventional sections.

Ames, Milton B , Jr↗
NASA NTRS1973

Results of tests OA12 and IA9 in the Ames Research Center Unitary Plan Wind Tunnels on an 0.030-scale model of the Space Shuttle Vehicle 2A to determine aerodynamic loads, volume 3

Tests were conducted in the NASA/ARC Unitary Plan Wind Tunnels during April and May 1973, on an 0.030-scale replica of the Space Shuttle Vehicle Configuration 2A. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 3.5. The investigation included tests IA9A, B and C on the integrated (launch) configuration and tests OA12A and C on the isolated orbiter (entry configuration). The integrated vehicle was tested at angles of attack and sideslip from -8 degrees to +8 degrees. The isolated orbiter was tested at angles of attack from -15 degrees to +40 degrees and angles of sideslip from -10 degrees to +10 degrees to as dictated by trajectory considerations. The effects of orbiter/external tank incidence angle and deflected control surfaces on aerodynamic loads were also investigated.

Spangler, R. H.↗