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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Reducing the Effect of Transducer Mount Induced Noise on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

Characterization of launch vehicle unsteady aerodynamics is a field best studied through experimentation, which is often carried out in the form of large scale wind tunnel testing. Measurement of the fluctuating pressures induced by the boundary layer noise is customarily made with miniature pressure transducers installed into a model of the vehicle of interest. Literature shows that noise level increases between two to five decibels (dB referenced to 20 micropascal) can be induced when the transducer surface is not mounted perfectly flush with the model outer surface. To reduce this artificially induced noise, special transducer holders have been used for aeroacoustic wind tunnel testing by NASA. This holder is a sleeve into which the transducer fits, with a cap that allows it to be mounted in a recessed hole in the model. A single hole in the holder allows the transport of the tunnel medium so the transducer can discriminate the fluctuating pressure due to the turbulent boundary layer noise. The holder is first dry fitted into the model and any difference in height between the holder and the model surface can be sanded flush. The holder is then removed from the model, the transducer glued inside the holder, and the holder replaced in the model, secured also with glue, thus eliminating the problem of noise level increases due to lack of flushness. In order to work with this holder design, special transducers have been ordered with their standard screen removed and the diaphragm moved as close to the top of the casing as possible to minimize any cavity volume. Although this greatly reduces induced noise due to the transducers being out of flush, the holders can also induce a cavity resonance that is usually at a very high frequency. This noise is termed transducer mount induced noise (XMIN). The peak of the mode can vary with the cavity depth, boundary layer noise that can excite the mode, tunnel flow medium, and the build of the transducers. Because the boundary layer flow interaction with the microphone holder seems to have some effect on the cavity resonance response, the NASA/Marshall Space Flight Center (MSFC) aeroacoustic team postulated that changing the design of the holder might reduce the cavity resonance response. An experiment was performed at the MSFC Trisonic Wind Tunnel comparing a new holder design with the traditional MSFC holder. The new holder design is intended to mimic the Kulite Semiconductors, Inc. B-screen, where the single hole of the traditional MSFC holder is replaced by a series of ten 0.007 inch diameter holes. This particular design serves to protect the transducer diaphragm as in the manufacturer's B-screen, but also increases the frequency of the XMIN such that it is not apparent in the data collected from aeroacoustic wind tunnel tests. MSFC compared the two holder designs by installing two transducers in standard holders on one side of a ten inch long ten degree cone, and two transducers in B-screen holders on the opposite side for a series of runs at different conditions. Over 11-13 February, 2013, the model was run at Mach numbers between 0.80 and 1.96 at angles of attack at 0 deg, plus or minus 1 deg, plus or minus 2 deg, plus or minus 4 deg, and plus or minus 8 deg. Model roll attitudes at 0 deg, 90 deg, and 180 deg were tested to isolate potential tunnel bias. This paper presents results from the test showing that the proposed holder design significantly reduces the influence of XMIN on measured fluctuating pressure levels.

Herron, A. J.↗
NASA NTRS2015

Reducing the Effect of Transducer Mount Induced Noise (XMIN) on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

Characterization of flight vehicle unsteady aerodynamics is often studied via large scale wind tunnel testing. Boundary layer noise is measured by miniature pressure transducers installed in a model. Noise levels (2-5 dB ref. 20 μPa) can be induced when transducer is mounted out of flush with model outer surface. This effect must be minimized to accurately determine aerodynamically induced acoustic environments.

Herron, Andrew J.↗
NASA NTRSDate not supplied

Comparison of Prediction Modeling Methodologies for Aeroacoustic Characterization of Hovering sUAS Rotors

This work compared artificial neural network and multivariate orthogonal function modeling methodologies for the prediction and characterization of isolated hovering sUAS rotor aerodynamics and aeroacoustics. Design of Experiments was used to create input feature spaces over 9 input features: the number of rotor blades, rotor size, rotor speed, the amount of blade twist, blade taper ratio, tip chord length, collective pitch, airfoil camber, and airfoil thickness. CAMRAD~II and AARON were executed at the points defined by the input feature space to predict aerodynamic and aeroacoustic quantities. These predicted aerodynamic and aeroacoustic data were then used to generate artificial neural networks and polynomial response surface models. The two prediction model methodologies were evaluated over test data previously unseen by the models, which showed good prediction capabilities for both model types, with slightly lower prediction error for the artificial neural networks. A characterization study was performed, which showed that input features correspondent to the spanwise sectional blade lift and drag were the most significant factors to the aerodynamic thrust and power, respectively. It was also shown that the aeroacoustic quantities were highly dependent on variations in rotor speed and size, which affect the Doppler factor for tonal noise and the spanwise Reynolds number for broadband noise.

Christopher S Thurman↗
NASA NTRS1985

The structure and characterization of air-assisted swirl atomizer sprays

The detailed aerodynamic structure of air-assisted swirl atomizer sprays is investigated in this paper. In contrast to previously published papers on spray characterization, special emphasis is placed on the region near the atomizer exit where significant variation of the mean drop size and drop velocity occurs. Simple correlations are found for the spray mean drop size and some initial input parameters. Measurements of drop acceleration and deceleration can be well related to the line-of-sight variations in mean drop size and number density. It is found that the initial drop acceleration and spray dispersion cause the initial decrease of axial mean drop diameters and subsequent drop deceleration causes the increase of mean drop diameters.

Mao, C.-P.↗
NASA NTRS1988

Influence of wind shear on the aerodynamic characteristics of airplanes

The aerodynamic effect of shear flow through a series of sensitivity studies of the wind velocity gradients and wing planform geometry parameters is investigated and characterized. The wind shear effect is computed using a modified vortex-lattice computer program and characterized through the formulation of wind shear aerodynamic coefficients. The magnitudes if the aerodynamic effects are demonstrated by computation of the resultant change in the aerodynamics of a conventional wing and horizontal stability configuration on a fixed flight path through a simulated microburst. The results indicate that as much as 20 percent of the control authority of the airplane may be required to counteract the wind-shear-induced forces and moments in the microburst environment.

Vicroy, Dan D.↗
NASA NTRS2009

Findings from the Supersonic Qualification Program of the Mars Science Laboratory Parachute System

In 2012, the Mars Science Laboratory Mission (MSL) will deploy NASA's largest extra-terrestrial parachute, a technology integral to the safe landing of its advanced robotic explorer on the surface. The supersonic parachute system is a mortar deployed 21.5 m disk-gap-band (DGB) parachute, identical in geometric scaling to the Viking era DGB parachutes of the 1970's. The MSL parachute deployment conditions are Mach 2.3 at a dynamic pressure of 750 Pa. The Viking Balloon Launched Decelerator Test (BLDT) successfully demonstrated a maximum of 700 Pa at Mach 2.2 for a 16.1 m DGB parachute in its AV4 flight. All previous Mars deployments have derived their supersonic qualification from the Viking BLDT test series, preventing the need for full scale high altitude supersonic testing. The qualification programs for Mars Pathfinder, Mars Exploration Rover, and Phoenix Scout Missions were all limited to subsonic structural qualification, with supersonic performance and survivability bounded by the BLDT qualification. The MSL parachute, at the edge of the supersonic heritage deployment space and 33% larger than the Viking parachute, accepts a certain degree of risk without addressing the supersonic environment in which it will deploy. In addition, MSL will spend up to 10 seconds above Mach 1.5, an aerodynamic regime that is associated with a known parachute instability characterized by significant canopy projected area fluctuation and dynamic drag variation. This aerodynamic instability, referred to as "area oscillations" by the parachute community has drag performance, inflation stability, and structural implications, introducing risk to mission success if not quantified for the MSL parachute system. To minimize this risk and as an alternative to a prohibitively expensive high altitude test program, a multi-phase qualification program using computation simulation validated by subscale test was developed and implemented for MSL. The first phase consisted of 2% of fullscale supersonic wind tunnel testing of a rigid DGB parachute with entry-vehicle to validate two high fidelity computational fluid dynamics (CFD) tools. The computer codes utilized Large Eddy Simulation and Detached Eddy Simulation numerical approaches to accurately capture the turbulent wake of the entry vehicle and its coupling to the parachute bow-shock. The second phase was the development of fluid structure interaction (FSI) computational tools to predict parachute response to the supersonic flow field. The FSI development included the integration of the CFD from the first phase with a finite element structural model of the parachute membrane and cable elements. In this phase, a 4% of full-scale supersonic flexible parachute test program was conducted to provide validation data to the FSI code and an empirical dataset of the MSL parachute in a flight-like environment. The final phase is FSI simulations of the full-scale MSL parachute in a Mars type deployment. Findings from this program will be presented in terms of code development and validation, empirical findings from the supersonic testing, and drag performance during supersonic operation.

EDL↗
NASA NTRS2005

Time Series Vegetation Aerodynamic Roughness Fields Estimated from MODIS Observations

Most land surface models used today require estimates of aerodynamic roughness length in order to characterize momentum transfer between the surface and atmosphere. The most common method of prescribing roughness is through the use of empirical look-up tables based solely on land cover class. Theoretical approaches that employ satellite-based estimates of canopy density present an attractive alternative to current look-up table approaches based on vegetation cover type that do not account for within-class variability and are oftentimes simplistic with respect to temporal variability. The current research applies Raupach s formulation of momentum aerodynamic roughness to MODIS data on a regional scale in order to estimate seasonally variable roughness and zero-plane displacement height fields using bulk land cover parameters estimated by [Jasinski, M.F., Borak, J., Crago, R., 2005. Bulk surface momentum parameters for satellite-derived vegetation fields. Agric. For. Meteorol. 133, 55-68]. Results indicate promising advances over look-up approaches with respect to characterization of vegetation roughness variability in land surface and atmospheric circulation models.

Borak, Jordan S.↗
NASA NTRSDate not supplied

Evaluating the Utility of Pressure Scanners for Unsteady Pressure Measurements in Wind Tunnel Characterization of the Space Launch System

In many aerodynamic measurement applications, the ability to make time-accurate measurements with pressure scanners provides practical advantages over making the same measurements directly using flush-mounted transducers. The pneumatic distortion due to the tubing between the sensing module and surface port limits the usefulness of such measurements to averaged, steady data only. This issue has restricted the application of pressure scanners for unsteady pressure measurements in wind tunnel and flight-testing applications. The Wiener-filtered inverse system response model has been proven to be effective in reconstructing complex unsteady pressure signals acquired remotely using pressure scanners. In this work, the feasibility of using pressure scanning systems for time-resolved pressure measurements in the NASA Langley 14- by 22-Foot Subsonic Wind Tunnel was evaluated. Results demonstrated good agreement between the reference transducers and reconstructed pressure scanner measurements up to frequencies of approximately 500 Hz.

Unsteady pressure↗
NASA NTRS1972

Unsteady rotor aerodynamics at low inflow and its effect on flutter.

A comprehensive flutter analysis has been developed utilizing compressible unsteady aerodynamics for a helicopter rotor at a low inflow condition. The rotor is treated as a nonuniform rotating beam with finite bending and torsional stiffness. The unsteady aerodynamic representation incorporates the effects of spanwise variations in Mach number and reduced frequency by applying available two-dimensional theories in a strip theory fashion. The results are characterized by an oscillation of the apparent aerodynamic damping with decreasing frequency ratio (oscillation frequency/rotational frequency) which suggests a new flutter criteria for rotors. These results are correlated with available experimental data.

Pierce, G. A.↗
NASA NTRS1972

Aerodynamic parameters of the Navion airplane extracted from flight

An iterative method, which is characterized as a maximum-likelihood minimum-variance technique, was used to extract the aerodynamic parameters of a Navion airplane from flight data. The purposes were to compare the results with parameters obtained from wind-tunnel tests and with results obtained by analog matching the same data, and to develop techniques for application of the parameter extraction program. Results from the study showed that the parameter-extraction program can produce aerodynamic parameters which will permit close estimation of the aircraft time histories used in the extraction process. The program determined an estimate of the standard deviations of the states and parameters. These estimates were used to indicate how well the calculated states fit the flight data and the confidence in the values of the estimated parameters. The study also showed that the values of the parameters were affected by the data and mathematical model used during the extraction process. Because of the lack of confidence in the parameters extracted by use of some of the sets of data, several parameters were estimated by other methods. By using a combination of methods, a set of parameters which gave a fit to the data was obtained.

Suit, W. T.↗
NASA NTRSDate not supplied

Uncertainty Analysis of the Pitch Damping Coefficient of Blunt Bodies, Measured from Magnetic Suspension Wind Tunnel Tests

Blunt body aerodynamics in re-entry conditions are characterized by oscillatory behavior that is often encapsulated by the pitch damping coefficient (Cmq). Magnetic Suspension Wind Tunnel (MSWT) testing with a Magnetic Suspension Balance System (MSBS) aims to improve blunt body wind tunnel testing by removing the sometimes unknown aerodynamic effects of traditional sting mounts, as well as generate more constrained outputs than vertical spin tunnel testing. Efforts to calculate the uncertainty of the pitch damping coefficient with a proposed Earth Entry Vehicle (EEV) capsule model have been made by fitting an analytic prediction of the attitude time history to video data of a free-to-oscillate MSBS test.

Magnetic suspension, dynamic stability, uncertaint↗
NASA NTRS1991

Automatic fringe analysis

To satisfy the requirement for fast, accurate interferometric analytical tools, the Fringe Analysis Workstation (FAW) has been developed to analyze complex fringe image data easily and rapidly. FAW is employed for flow studies in hydrodynamics and aerodynamics experiments, and for target shell characterization in inertial confinement fusion research. Three major components of the FAW system: fringe analysis/image processing, input/output, and visualization/graphical user interface are described.

Chiu, Arnold↗
NASA NTRS2003

An Aerodynamic Simulation Process for Iced Lifting Surfaces and Associated Issues

This paper discusses technologies and software tools that are being implemented in a software toolkit currently under development at NASA Glenn Research Center. Its purpose is to help study the effects of icing on airfoil performance and assist with the aerodynamic simulation process which consists of characterization and modeling of ice geometry, application of block topology and grid generation, and flow simulation. Tools and technologies for each task have been carefully chosen based on their contribution to the overall process. For the geometry characterization and modeling, we have chosen an interactive rather than automatic process in order to handle numerous ice shapes. An Appendix presents features of a software toolkit developed to support the interactive process. Approaches taken for the generation of block topology and grids, and flow simulation, though not yet implemented in the software, are discussed with reasons for why particular methods are chosen. Some of the issues that need to be addressed and discussed by the icing community are also included.

Choo, Yung K.↗
NASA NTRS2004

Multiobjective Aerodynamic Shape Optimization Using Pareto Differential Evolution and Generalized Response Surface Metamodels

Differential Evolution (DE) is a simple, fast, and robust evolutionary algorithm that has proven effective in determining the global optimum for several difficult single-objective optimization problems. The DE algorithm has been recently extended to multiobjective optimization problem by using a Pareto-based approach. In this paper, a Pareto DE algorithm is applied to multiobjective aerodynamic shape optimization problems that are characterized by computationally expensive objective function evaluations. To improve computational expensive the algorithm is coupled with generalized response surface meta-models based on artificial neural networks. Results are presented for some test optimization problems from the literature to demonstrate the capabilities of the method.

Madavan, Nateri K.↗
NASA NTRS2011

Results of an Advanced Fan Stage Operating Over a Wide Range of Speed and Bypass Ratio: Fan Stage Design and Experimental Results - Part 1

NASA s Fundamental Aeronautics Program is investigating turbine-based combined cycle (TBCC) propulsion systems for access to space because it provides the potential for aircraft-like, space-launch operations that may significantly reduce launch costs and improve safety. To this end, National Aeronautics and Space Administration (NASA) and General Electric (GE) teamed to design a Mach 4 variable cycle turbofan/ramjet engine for access to space. To enable the wide operating range of a Mach 4+ variable cycle turbofan ramjet required the development of a unique fan stage design capable of multi-point operation to accommodate variations in bypass ratio (10 ), fan speed (7 ), inlet mass flow (3.5 ), inlet pressure (8 ), and inlet temperature (3 ). In this paper, NASA has set out to characterize a TBCC engine fan stage aerodynamic performance and stability limits over a wide operating range including power-on and hypersonic-unique "windmill" operation. Herein, we will present the fan stage design, and the experimental test results of the fan stage operating from 15 to 100 percent corrected design speed. Whereas, in the companion paper, we will provide an assessment of NASA s APNASA code s ability to predict the fan stage performance and operability over a wide range of speed and bypass ratio.

Suder, Kenneth L.↗
NASA NTRS2016

Overview of the Space Launch System Ascent Aeroacoustic Environment Test Program

Characterization of accurate flight vehicle unsteady aerodynamics is critical for component and secondary structure vibroacoustic design. The Aerosciences Branch at the National Aeronautics and Space Administration (NASA) Marshall Space Flight Center has conducted a test at the NASA Ames Research Center (ARC) Unitary Plan Wind Tunnels (UPWT) to determine such ascent aeroacoustic environments for the Space Launch System (SLS). Surface static pressure measurements were also collected to aid in determination of local environments for venting, CFD substantiation, and calibration of the flush air data system located on the launch abort system. Additionally, this test supported a NASA Engineering and Safety Center study of alternate booster nose caps. Testing occurred during two test campaigns: August - September 2013 and December 2013 - January 2014. Four primary model configurations were tested for ascent aeroacoustic environment definition. The SLS Block 1 vehicle was represented by a 2.5% full stack model and a 4% truncated model. Preliminary Block 1B payload and manned configurations were also tested, using 2.5% full stack and 4% truncated models respectively. This test utilized the 11 x 11 foot transonic and 9 x 7 foot supersonic tunnel sections at the ARC UPWT to collect data from Mach 0.7 through 2.5 at various total angles of attack. SLS Block 1 design environments were developed primarily using these data. SLS Block 1B preliminary environments have also been prepared using these data. This paper discusses the test and analysis methodology utilized, with a focus on the unsteady data collection and processing.

Herron, Andrew J.↗
NASA NTRS2019

Space Launch System Aeroacoustic Wind Tunnel Test Results

Characterization of accurate launch vehicle unsteady aerodynamics is critical for component and secondary structure vibroacoustic design. For the National Aeronautics and Space Administration (NASA) Space Launch System (SLS), aeroacoustic environments have been derived primarily through sub-scale wind tunnel testing. Both optical techniques and high frequency pressure measurements have been utilized across multiple testing facilities and numerous vehicle configurations to develop a range of preliminary and detailed environments. As the vehicle has matured and evolved, the data collected from each subsequent configuration has allowed for comparison studies which isolate the effects of certain outer mold line (OML) features on measured fluctuating pressure levels. This paper presents observations on some of those effects for features which include abort system protuberances, various fairings geometries, interstage flanges, and multibody interactions between a central core and fall away boosters. These features, and the flow conditions produced by them, are broadly applicable to many launch vehicle configurations.

Steva, Thomas B.↗
NASA NTRS1989

Effect of control surface mass unbalance on the stability of a closed-loop active control system

An inertial energy approach similar to the aerodynamic energy method for flutter suppression is used to investigate the effects of mass-unbalanced control surfaces on the stability of a closed-loop system. It is demonstrated that a spanwise section for sensor location can be obtained which ensures minimum sensitivity to the mode shapes of the aircraft. Leading-edge control is characterized by a compatibility between inertial stabilization and aerodynamic stabilization that trailing-edge control lacks.

Nissim, E.↗