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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 73 records · Page 4
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NASA NTRS2016

Wind Tunnel Test of Subscale Ringsail and Disk-Gap-Band Parachutes

A subsonic wind tunnel test was conducted to determine the drag and static aerodynamic coefficients, as well as to capture the dynamic motions of a new Supersonic Ringsail parachute developed by the Low Density Supersonic Decelerator Project. To provide a comparison against current Mars parachute technology, the Mars Science Laboratory's Disk-Gap-Band parachute was also included in the test. To account for the effect of fabric permeability, two fabrics ("low" and "standard" permeability) were used to fabricate each parachute canopy type, creating four combinations of canopy type and fabric material. A wide range of test conditions were covered during the test, spanning Mach numbers from 0.09 to 0.5, and static pressures from 103 to 2116 pounds per square inch (psf) (nominal values). The fabric permeability is shown to have a first-order effect on the aerodynamic coefficients and dynamic motions of the parachutes. For example, for a given parachute type and test condition, models fabricated from "low" permeability fabric always have a larger drag coefficient than models fabricated from "standard" permeability material. This paper describes the test setup and conditions, how the results were analyzed, and presents and discusses a sample of the results. The data collected during this test is being used to create and improve parachute aerodynamic databases for use in flight dynamics simulations for missions to Mars.

Zumwalt, Carlie H.↗
NASA NTRSDate not supplied

Thermal Testing and Analysis Techniques for Wires and Wire Bundles

Determining wire and wire bundle amperage capacity (i.e., “ampacity”) currently relies on the use of standards to derate wire ampacity when in a bundle configuration. The feasibility of developing physics-based and regression thermal models of single wires and wire bundles to determine ampacity using a customized test apparatus was investigated during a pathfinder study. A test facility was developed and various wire and wire bundle articles were tested under a variety of temperature and pressure conditions using an efficient test matrix formulated using Design of Experiments (DOE) techniques. Physics-based models were developed and correlated to the test results. Regression models were formulated and compared to test results and standards.

Thermal↗
NASA NTRS1988

Adaptive wall research with two- and three-dimensional models in low speed and transonic tunnels

This paper summarises recent research at the University of Southampton into adaptive wall technology and outlines the direction of current efforts. The work is aimed at developing techniques for use in test sections where the top and bottom walls may be adjusted in single curvature. Wall streamlining eliminates, as far as experimentally possible, the top and bottom wall interference in low speed and transonic aerofoil testing. A streamlining technique has been developed for low speeds which allows the testing of swept wing panels in low interference environments. At higher speeds, a comparison of several two-dimensional transonic streamlining algorithms has been made and a technique for streamlining with a choked test section has also been developed. Three-dimensional work has mainly concentrated on tests of sidewall mounted half-wings and the development of the software packages required to assess interference and to adjust the flexible walls. It has been demonstrated that two-dimensional wall adaptation can significantly modify the level of wall interference around relatively large three-dimensional models. The residual interferences are small and are probably amenable to standard post-test correction methods. Tests on a calibrated wing-body model are planned in the near future to further validate the proposed streamlining technique.

Lewis, M. C.↗
NASA NTRS1978

Performance of preproduction model cesium beam frequency standards for spacecraft applications

A cesium beam frequency standards for spaceflight application on Navigation Development Satellites was designed and fabricated and preliminary testing was completed. The cesium standard evolved from an earlier prototype model launched aboard NTS-2 and the engineering development model to be launched aboard NTS satellites during 1979. A number of design innovations, including a hybrid analog/digital integrator and the replacement of analog filters and phase detectors by clocked digital sampling techniques are discussed. Thermal and thermal-vacuum testing was concluded and test data are presented. Stability data for 10 to 10,000 seconds averaging interval, measured under laboratory conditions, are shown.

Levine, M. W.↗
NASA NTRS1983

Radiation hardness of Efratom M-100 rubidium frequency standard

The effects of nuclear radiation on rubidium gas cell frequency standards and components are presented, including the results of recent tests where a continuously operating rubidium frequency standard (Effratom, Model M-100) was subjected to simultaneous neutron/gamma radiation. At the highest neutron fluence 7.5 10 to the 12th power n/sq cm and total dose 11 krad(Si) tested, the unit operated satisfactorily; the total frequency change over the 2 1/2 hour test period due to all causes, including repeated retraction from and insertion into the reactor, was less than 1 x 10 to the -10th power. The effects of combined neutron/gamma radiation on rubidium frequency standard physics package components were also studied, and the results are presented.

English, T. C.↗
NASA NTRS1986

Validation of large space structures by ground tests

The paper presents concepts designed to validate, through the use of ground tests, mathematical models of continuous type structures and structures comprised of interconnecting subsystems. For continuous-type structures, a multiple boundary condition test approach is considered in which the basic idea is to perform a large number of tests using artificial boundary conditions from which good ground test data can be obtained. The test ensures an arbitrarily large number of test data for use in the validation and updating of the mathematical model. Another approach, applicable to structures comprised of subsystems, involves the identification of significant structural elements for the system dynamic model which are not validated by standard modal tests of the subsystems.

Wada, B. K.↗
NASA NTRS2011

Testing Einstein's Gravity on Large Scales

A little over a decade has passed since two teams studying high redshift Type Ia supernovae announced the discovery that the expansion of the universe was accelerating. After all this time, we?re still not sure how cosmic acceleration fits into the theory that tells us about the large-scale universe: General Relativity (GR). As part of our search for answers, we have been forced to question GR itself. But how will we test our ideas? We are fortunate enough to be entering the era of precision cosmology, where the standard model of gravity can be subjected to more rigorous testing. Various techniques will be employed over the next decade or two in the effort to better understand cosmic acceleration and the theory behind it. In this talk, I will describe cosmic acceleration, current proposals to explain it, and weak gravitational lensing, an observational effect that allows us to do the necessary precision cosmology.

Prescod-Weinstein, Chandra↗
NASA NTRS2010

NASA Ares I Launch Vehicle Upper Stage Reaction Control System (ReCS) Cold Flow Development Test Overview

NASA s Ares I launch vehicle, consisting of a five segment solid rocket booster first stage and a liquid bi-propellant J2-X engine Upper Stage, is the vehicle that s been chosen to launch the Orion Crew Module, which will return humans to the Moon, Mars, and beyond. After First Stage booster separation, the Reaction Control System (ReCS), a monopropellant hydrazine system, will provide the Upper Stage element with three degrees of freedom control as needed. This paper provides an overview of the system level development testing that has taken place on the Ares I launch vehicle Upper Stage ReCS. The ReCS System Development Test Article (SDTA) was built as a flight representative water flow test article whose primary test objective was to obtain fluid system performance data to evaluate the integrate system performance characteristics and verify analytical models. Water is the industry standard for cold flow testing of hydrazine systems, because the densities are very close and the speeds of sound are well characterized. The completion of this development level test program was considered necessary to support the ReCS Critical Design Review. This paper will address the design approach taken in building the test article, the objectives of the test program, types of testing completed, general results, the ability of the program to meet the test objectives, and lessons learned

Dervan, Melanie↗
NASA NTRS1999

Uncertainty Analysis of Inertial Model Attitude Sensor Calibration and Application with a Recommended New Calibration Method

Statistical tools, previously developed for nonlinear least-squares estimation of multivariate sensor calibration parameters and the associated calibration uncertainty analysis, have been applied to single- and multiple-axis inertial model attitude sensors used in wind tunnel testing to measure angle of attack and roll angle. The analysis provides confidence and prediction intervals of calibrated sensor measurement uncertainty as functions of applied input pitch and roll angles. A comparative performance study of various experimental designs for inertial sensor calibration is presented along with corroborating experimental data. The importance of replicated calibrations over extended time periods has been emphasized; replication provides independent estimates of calibration precision and bias uncertainties, statistical tests for calibration or modeling bias uncertainty, and statistical tests for sensor parameter drift over time. A set of recommendations for a new standardized model attitude sensor calibration method and usage procedures is included. The statistical information provided by these procedures is necessary for the uncertainty analysis of aerospace test results now required by users of industrial wind tunnel test facilities.

Tripp, John S.↗
NASA NTRS1984

A review and analysis of boundary layer transition data for turbine application

A symposium on transition in turbines was held at the NASA Lewis Research Center. One recommendation of the working groups was the collection of existing transition data to provide standard cases against which models could be tested. A number of data sets from the open literature that include heat transfer data in apparently transitional boundary layers, with particular application to the turbine environment, were reviewed and analyzed to extract transition information from the heat transfer data. The data sets reviewed cover a wide range of flow conditions, from low speed, flat plate tests to full scale turbine airfoils operating at simulated turbine engine conditions. The results indicate that free stream turbulence and pressure gradient have strong, and opposite, effects on the location of the start of transition and on the length of the transition zone.

Gaugler, R. E.↗
NASA NTRS1985

A review and analysis of boundary layer transition data for turbine application

A symposium on transition in turbines was held at the NASA Lewis Research Center. One recommendation of the working groups was the collection of existing transition data to provide standard cases against which models could be tested. A number of data sets from the open literature that include heat transfer data in apparently transitional boundary layers, with particular application to the turbine environment, were reviewed and analyzed to extract transition information from the heat transfer data. The data sets reviewed cover a wide range of flow conditions, from low speed, flat plate tests to full scale turbine airfoils operating at simulated turbine engine conditions. The results indicate that free stream turbulence and pressure gradient have strong, and opposite, effects on the location of the start of transition and on the length of the transition zone.

Gaugler, R. E.↗
NASA NTRS1993

Laboratory studies of low temperature rate coefficients: The atmospheric chemistry of the outer planets

The objectives are to measure laboratory rate coefficients for key reactions of hydrocarbon molecules and radicals at low temperatures, which are relevant to the atmospheric photochemistry of Saturn, Jupiter, and Titan. Upcoming NASA planetary missions, such as Cassini, will probe the atmosphere of Titan in more detail, offering an excellent opportunity to test kinetic models and to establish fiducial standards for using kinetic models to interpret various parameters of the outer planets. Accurate low temperature kinetic data, which are presently lacking, may require crucial revisions to the rates of formation and destruction and are of utmost importance to the success of these efforts. In this program, several key reactions of ethynyl radicals (C2H) with acetylene (C2H2), methane (CH4), and oxygen (O2), down to temperatures of 170 K were successfully investigated. The experimental apparatus developed in our laboratory for measuring reaction kinetics at low temperatures consists of a laser photolysis/infrared probe laser setup. The rate measurements are carried out as a function of (low) temperature with a transverse flow cell designed specifically for these studies. A 193 nm argon fluoride pulsed excimer laser is used to photolyze a suitable precursor molecule, such as acetylene to produce C2H, and a high resolution, tunable infrared F-center laser (2.3-3.35 mu m) probes the transient concentrations of the radical species directly in absorption to extract the kinetic rate coefficients.

Leone, Stephen R.↗
NASA NTRS2000

Test Cases for a Rectangular Supercritical Wing Undergoing Pitching Oscillations

Steady and unsteady measured pressures for a Rectangular Supercritical Wing (RSW) undergoing pitching oscillations have been presented. From the several hundred compiled data points, 27 static and 36 pitching oscillation cases have been proposed for computational Test Cases to illustrate the trends with Mach number, reduced frequency, and angle of attack. The wing was designed to be a simple configuration for Computational Fluid Dynamics (CFD) comparisons. The wing had an unswept rectangular planform plus a tip of revolution, a panel aspect ratio of 2.0, a twelve per cent thick supercritical airfoil section, and no twist. The model was tested over a wide range of Mach numbers, from 0.27 to 0.90, corresponding to low subsonic flows up to strong transonic flows. The higher Mach numbers are well beyond the design Mach number such as might be required for flutter verification beyond cruise conditions. The pitching oscillations covered a broad range of reduced frequencies. Some early calculations for this wing are given for lifting pressure as calculated from a linear lifting surface program and from a transonic small perturbation program. The unsteady results were given primarily for a mild transonic condition at M = 0.70. For these cases the agreement with the data was only fair, possibly resulting from the omission of viscous effects. Supercritical airfoil sections are known to be sensitive to viscous effects (for example, one case cited). Calculations using a higher level code with the full potential equations have been presented for one of the same cases, and with the Euler equations. The agreement around the leading edge was improved, but overall the agreement was not completely satisfactory. Typically for low-aspect-ratio rectangular wings, transonic shock waves on the wing tend to sweep forward from root to tip such that there are strong three-dimensional effects. It might also be noted that for most of the test, the model was tested with free transition, but a few points were taken with an added transition strip for comparison. Some unpublished results of a rigid wing of the same airfoil and planform that was tested on the pitch and plunge apparatus mount system (PAPA) showed effects of the lower surface transition Strip on flutter at the lower subsonic Mach numbers. Significant effects of a transition strip were also obtained on a wing with a thicker supercritical section on the PAPA mount system. Both of these flutter tests on the PAPA resulted in very low reduced frequencies that may be a factor in this influence of the transition strip. However, these results indicate that correlation studies for RSW may require some attention to the estimation of transition location to accurately treat viscous effects. In this report several Test Cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flow effects. An overview of the model and tests is given and the standard formulary for these data is listed. Sample data points are presented in both tabular and graphical form. A complete tabulation and plotting of all the Test Cases is given. Only the static pressures and the real and imaginary parts of the first harmonic of the unsteady pressures are available. All the data for the test are available in electronic file form. The Test Cases are also available as separate electronic files.

Bennett, Robert M.↗
NASA NTRS2000

Test Cases for the Benchmark Active Controls: Spoiler and Control Surface Oscillations and Flutter

As a portion of the Benchmark Models Program at NASA Langley, a simple generic model was developed for active controls research and was called BACT for Benchmark Active Controls Technology model. This model was based on the previously-tested Benchmark Models rectangular wing with the NACA 0012 airfoil section that was mounted on the Pitch and Plunge Apparatus (PAPA) for flutter testing. The BACT model had an upper surface spoiler, a lower surface spoiler, and a trailing edge control surface for use in flutter suppression and dynamic response excitation. Previous experience with flutter suppression indicated a need for measured control surface aerodynamics for accurate control law design. Three different types of flutter instability boundaries had also been determined for the NACA 0012/PAPA model, a classical flutter boundary, a transonic stall flutter boundary at angle of attack, and a plunge instability near M = 0.9. Therefore an extensive set of steady and control surface oscillation data was generated spanning the range of the three types of instabilities. This information was subsequently used to design control laws to suppress each flutter instability. There have been three tests of the BACT model. The objective of the first test, TDT Test 485, was to generate a data set of steady and unsteady control surface effectiveness data, and to determine the open loop dynamic characteristics of the control systems including the actuators. Unsteady pressures, loads, and transfer functions were measured. The other two tests, TDT Test 502 and TDT Test 5 18, were primarily oriented towards active controls research, but some data supplementary to the first test were obtained. Dynamic response of the flexible system to control surface excitation and open loop flutter characteristics were determined during Test 502. Loads were not measured during the last two tests. During these tests, a database of over 3000 data sets was obtained. A reasonably extensive subset of the data sets from the first two tests have been chosen for Test Cases for computational comparisons concentrating on static conditions and cases with harmonically oscillating control surfaces. Several flutter Test Cases from both tests have also been included. Some aerodynamic comparisons with the BACT data have been made using computational fluid dynamics codes at the Navier-Stokes level (and in the accompanying chapter SC). Some mechanical and active control studies have been presented. In this report several Test Cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flow effects. Cases for static angles of attack, static trailing-edge and upper-surface spoiler deflections are included for a range of conditions near those for the oscillation cases. Cases for trailing-edge control and upper-surface spoiler oscillations for a range of Mach numbers, angle of attack, and static control deflections are included. Cases for all three types of flutter instability are selected. In addition some cases are included for dynamic response measurements during forced oscillations of the controls on the flexible mount. An overview of the model and tests is given, and the standard formulary for these data is listed. Some sample data and sample results of calculations are presented. Only the static pressures and the first harmonic real and imaginary parts of the pressures are included in the data for the Test Cases, but digitized time histories have been archived. The data for the Test Cases are also available as separate electronic files.

Bennett, Robert M.↗
NASA NTRS2000

Test Cases for a Clipped Delta Wing with Pitching and Trailing-Edge Control Surface Oscillations

Steady and unsteady measured pressures for a Clipped Delta Wing (CDW) undergoing pitching oscillations and trailing-edge control surface oscillations have been presented . From the several hundred compiled data points, 22 static cases, 12 pitching-oscillation cases, and 12 control-surface-oscillation cases have been proposed for Computational Test Cases to illustrate the trends with Mach number, reduced frequency, and angle of attack. The planform for this wing was derived by simplifying the planform of a proposed design for a supersonic transport which is described as the Boeing 2707-300. The strake was deleted, the resulting planform was approximated by a trapezoid with an unswept trailing edge, and the twist and camber were removed. In order to facilitate pressure instrumentation, the thickness was increased to 6 percent from the typical 2.5 to 3 percent for the supersonic transport. The airfoil is thus a symmetrical circular arc section with t/c = 0.06. A wing of similar planform but with a thinner airfoil of t/c = 0.03 was used in the flutter investigations, and the buffet and stall flutter investigation . Flutter results are also reported both for the 3 per cent thick simplified wing and for a more complex SST model. One of the consequences of the increased thickness of the clipped delta wing is that transonic effects are enhanced for Mach numbers near one. They are significantly stronger than would be the case for the thinner wing. Also, with the combination of high leading edge sweep of 50.5, and the sharp leading edge, a leading edge vortex forms on the wing at relatively low angles of attack, on the order of three degrees. The Appendix discusses some of the vortex flow effects. In addition, a shock develops over the aft portion of the wing at transonic speeds such that at some angles of attack, there is both a leading edge vortex and a shock wave on the wing. Such cases are a computational challenge. Some previous applications of this data set have been for the evaluation of an aerodynamic panel method and for evaluation of a Navier-Stokes capability. Linear theory and panel method results are also presented, which demonstrated the need for inclusion of transonic effects. Flutter calculations for the related wing with t/c=O.O3 are given. In this report several Test Cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flow effects. An overview of the model and tests are given, and the standard formulary for these data is listed. For each type of data, a sample table and a sample plot of the measured pressures are presented. A complete tabulation and plotting of the Test Cases is given. Only the static pressures and the 1st harmonic real and imaginary parts of the pressures are available. All of the data for the test are included in a microfiche document in the original report and are available in electronic file form. The Test Cases are also available as separate electronic files.

Bennett, Robert M.↗
NASA NTRS2014

Evaluation of Human and Anthropomorphic Test Device Finite Element Models under Spaceflight Loading Conditions

In an effort to develop occupant protection standards for future multipurpose crew vehicles, the National Aeronautics and Space Administration (NASA) has looked to evaluate the test device for human occupant restraint with the modification kit (THOR‐K) anthropomorphic test device (ATD) in relevant impact test scenarios. With the allowance and support of the National Highway Traffic Safety Administration, NASA has performed a series of sled impact tests on the latest developed THOR‐K ATD. These tests were performed to match test conditions from human volunteer data previously collected by the U.S. Air Force. The objective of this study was to evaluate the THOR‐K finite element (FE) model and the Total HUman Model for Safety (THUMS) FE model with respect to the tests performed. These models were evaluated in spinal and frontal impacts against kinematic and kinetic data recorded in ATD and human testing. Methods: The FE simulations were developed based on recorded pretest ATD/human position and sled acceleration pulses measured during testing. Predicted responses by both human and ATD models were compared to test data recorded under the same impact conditions. The kinematic responses of the models were quantitatively evaluated using the ISO‐metric curve rating system. In addition, ATD injury criteria and human stress/strain data were calculated to evaluate the risk of injury predicted by the ATD and human model, respectively. Results: Preliminary results show well‐correlated response between both FE models and their physical counterparts. In addition, predicted ATD injury criteria and human model stress/strain values are shown to positively relate. Kinematic comparison between human and ATD models indicates promising biofidelic response, although a slightly stiffer response is observed within the ATD. Conclusion: As a compliment to ATD testing, numerical simulation provides efficient means to assess vehicle safety throughout the design process and further improve the design of physical ATDs. The assessment of the THOR‐K and THUMS FE models in a spaceflight testing condition is an essential first step to implementing these models in the computational evaluation of spacecraft occupant safety. Promising results suggest future use of these models in the aerospace field.

Putnam, Jacob P.↗
NASA NTRS2020

Structural Model Tuning Tool: User's Reference Manual

This report presents an efficient approach for tuning finite element models to match the measured ground vibration test and static test data. Frequencies, mode shapes, total weight, location of the center of gravity, and static deformation computed from the finite element model are matched to the measured data. The model tuning procedure used in this work is based on solving an optimization problem in which the errors for the considered metrics, between the finite element prediction and the measured data, are minimized. Analytical sensitivity values of performance indices are computed using the NASTRAN-generated sensitivity values together with the in-house computer codes, which allow for faster computational time and the use of gradient-based optimizers. The method is applied to the Aerostructures Test Wing IV model. The study shows that the military standard and the National Aeronautics and Space Administration standard for comparing analytical and experimental modal data are all satisfied. The final finite element model correlates well with the test data. The flutter speed decreases by 8.91 percent after model tuning compared with the original Aerostructures Test Wing IV design.

Chan-gi Pak↗
NASA NTRS2014

Repeatability Modeling for Wind-Tunnel Measurements: Results for Three Langley Facilities

Data from extensive check standard tests of seven measurement processes in three NASA Langley Research Center wind tunnels are statistically analyzed to test a simple model previously presented in 2000 for characterizing short-term, within-test and across-test repeatability. The analysis is intended to support process improvement and development of uncertainty models for the measurements. The analysis suggests that the repeatability can be estimated adequately as a function of only the test section dynamic pressure over a two-orders- of-magnitude dynamic pressure range. As expected for low instrument loading, short-term coefficient repeatability is determined by the resolution of the instrument alone (air off). However, as previously pointed out, for the highest dynamic pressure range the coefficient repeatability appears to be independent of dynamic pressure, thus presenting a lower floor for the standard deviation for all three time frames. The simple repeatability model is shown to be adequate for all of the cases presented and for all three time frames.

Hemsch, Michael J.↗